TM-55-1520-234-23-1

TM-55-1520-234-23-1
TM 55-1520-234-23-1
VOLUME I
TECHNICAL MANUAL
AVUM AND AVIM MAINTENANCE MANUAL
AH-1S (MOD)
This copy is a reprint which includes current
pages from Changes 1 through 75.
HEADQUARTERS, DEPARTMENT OF THE ARMY
30 SEPTEMBER 1976
TM 55 1520-234-23-1
C 77
CHANGE
HEADQUARTERS
DEPARTMENT OF THE ARMY,
WASHINGTON, D.C., 31 August 1994
NO. 77
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited
TM 55-1520-23023-1, 30 September 1976, is changed as follows:
1.
2.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
1-13 and 1-14
5-31 and 5-32
1-13 and 1-14
5-31 and 5-32
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army.
GORDON R. SULLIVAN
General, United States Army
Chief of Staff
Official:
MILTON H. HAMILTON
Administrative Assistant to the
Secretary of the Army
07371
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31-E, block no. 0145, requirements for
TM 55-1520-234-23-1.
URGENT
TM 55 1520-234-23-1
C76
CHANGE
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 29 July 1994
NO. 76
AVUM AND AVIM Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited
TM 55-1520-234-23-1, 30 September 1 976, is changed as follows:
1.
2.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
5-3 and 5-4
5-3 and 5-4
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
GORDON R. SULLIVAN
General, United States Army
Chief of Staff
Official:
MILTON H. HAMILTON
Administrative Assistant to the
Secretary of the Army
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31-E, block no. 0145, requirements for TM 55-1520-234-23 -1.
URGENT
URGENT
NOTICE: THIS CHANGE HAS BEEN PRINTED AND DISTRIBUTED OUT OF SEQUENCE. IT SHOULD BE
INSERTED IN THE MANUAL AND USED. UPON RECEIPT OF THE EARLIER SEQUENCED CHANGE
ENSURE A MORE CURRENT CHANGE PAGE IS NOT REPLACED WITH A LESS CURRENT PAGE.
TM 55-1520-234-23-1
C 75
CHANGE
NO. 75
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 16 September 1992
}
AVIATION UNIT AND AVIATION INTERMEDIATE
Maintenance Manual
HELICOPTER, ATTACK AH-IS (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1.
2.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
1-48A/ 1-48B
5-30G and 5-30H
1-48A/ 1-48B
5-30G and 5-30H
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
GORDON R. SULLIVAN
General, United States Army
Chief of Staff
Official:
MILTON H. HAMILTON
Administrative Assistant to the
Secretary of the Army
02543
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31-E, block no. 0145, requirements for TM 55-1520-234-23-1.
DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited.
URGENT
URGENT
NOTICE:
THIS CHANGE HAS BEEN PRINTED AND DISTRIBUTED OUT OF SEQUENCE. IT SHOULD BE
INSERTED IN THE MANUAL AND USED. UPON RECEIPT OF THE EARLIER SEQUENCED CHANGE
ENSURE A MORE CURRENT CHANGE PAGE IS NOT REPLACED WITH A LESS CURRENT PAGE.
TM 55-1520-234-23-1
C 74
CHANGE
NO. 74
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 15 July 1992
}
AVIATION UNIT AND AVIATION INTERMEDIATE
Maintenance Manual
HELICOPTER, ATTACK AH-IS (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1.
2.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
---5-5 and 5-6
5-4A/5-4B
5-5 and 5-6
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
GORDON R. SULLIVAN
General, United States Army
Chief of Staff
Official:
MILTON H. HAMILTON
Administrative Assistant to the
Secretary of the Army
01493
DISTRIBUTION :
To be distributed in accordance with DA Form 12-31-E, block no.
requirements for TM 55-1520-234-23-1.
0145, AVUM and AVIM maintenance
DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited.
URGENT
URGENT
NOTICE:
THIS CHANGE HAS BEEN PRINTED AND DISTRIBUTED OUT OF SEQUENCE. IT SHOULD BE
INSERTED IN THE MANUAL AND USED. UPON RECEIPT OF THE EARLIER SEQUENCED CHANGE
ENSURE A MORE CURRENT CHANGE PAGE IS NOT REPLACED WITH A LESS CURRENT PAGE.
TM 55-1520-234-23-1
C 73
CHANGE
NO. 73
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 29 May 1992
}
AVIATION UNIT AND AVIATION INTERMEDIATE
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1.
2.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
v through viii
---ix through xii
1-11 and 1-12
1-22A/1-22B
6-71 and 6-72
6-72A/6-72B
6-73 and 6-74
6-79 and 6-80
6-80A and 6-80B
6-80E and 6-80F
6-81 and 6-82
6-90E and 6-90F
6-91 and 6-92
6-92A and 6-92B
6-101 and 6-102
----
v through viii
viii A/viii B
ix through xii
1-11 and 1-12
1-22A/1-22B
6-71 and 6-72
6-72A/6-72B
6-73 and 6-74
6-79 and 6-80
6-80A and 6-80B
6-80E and 6-80F
6-81 and 6-82
6-90E and 6-90F
6-91 and 6-92
6-92A and 6-92B
6-101 and 6-102
6-104A through 6-104C/6-104D
Retain this sheet in front of manual for reference purposes.
URGENT
TM 55-1520-234-23-1
C 73
By Order of the Secretary of the Army:
GORDON R. SULLIVAN
General, United States Army
Chief of Staff
Official:
MILTON H. HAMILTON
Administrative Assistant to the
Secretary of the Army
01483
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31-E, block no., 0145, AVUM and AVIM maintenance
requirements for TM 55-1520-234-23-1.
URGENT
NOTICE:
THIS CHANGE HAS BEEN PRINTED AND DISTRIBUTED OUT OF SEQUENCE. IT SHOULD BE
INSERTED IN THE MANUAL AND USED. UPON RECEIPT OF THE EARLIER SEQUENCED CHANGE
ENSURE A MORE CURRENT CHANGE PAGE IS NOT REPLACED WITH A LESS CURRENT PAGE.
TM 55-1520-234-23-1
C 72
CHANGE
NO. 72
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 20 June 1992
}
AVIATION UNIT AND AVIATION INTERMEDIATE
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1.
2.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
5-54A/5-54B
5-55 and 5-56
5-54A/5-54B
5-55 and 5-56
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
CARL E. VUONO
General, United States Army
Chief of Staff
Official:
PATRICIA P. HICKERSON
Colonel, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31-E, block no., 0145, AVUM and AVIM maintenance
requirements for TM 55-1520-234-23-1.
URGENT
TM 55-1520-234-23-1
C 71
CHANGE
NO. 71
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 31 May 1991
}
AVIATION UNIT AND AVIATION INTERMEDIATE
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
iii and iv
xiii and xiv
1-2A and 1-2B
1-15 and 1-16
1-32A/1-32B
1-33 and 1-34
1-47 and 1-48
1-53 through 1-56
1-66A and 1-66B
1-67 and 1-63
1-70C and 1-70D
1-71 and 1-72
1-72A/ 1-72B
3-1 and 3-2
3-2C and 3-2D
3-3 and 3-4
3-4A/3-4B
4-4C and 4-4D
4-5 and 4-6
4-13 and 4-14
4-43 and 4-44
5-1 and 5-2
---5-3 and 5-4
5-9 and 5-10
5-10A and 5-10B
5-11 and 5-12
5- 1 2A/5- 1 2B
5-13 and 5-14
5-14A/5-14B
5-15 and 5-16
5-16A/5-16B
5-29 and 5-30
5-30AA and 5-30AB
---5-30BB and 5-30BC
5-30CF and 5-30CG
Insert pages
iii and iv
xiii and xiv
1-2A and 1-2B
1-15 and 1-16
1-32A and 1-32B
1-33 and 1-34
1-47 and 1-43
1-53 through 1-56
1-66A and 1-66B
1-67 and 1-68
1-70C and 1-70D
1-71 and 1-72
1-72A/ 1-72B
3-1 and 3-2
3-2C and 3-2D
3-3 and 3-4
3-4A/3-4B
4-4C through 4-4H
4-5 and 4-6
4-13 and 4-14
4-43 and 4-44
5-1 and 5-2
5-2C and 5-2D
5-3 and 5-4
5-9 and 5-10
5-10A through 5-10D
5-11 and 5-12
5- 1 2A/ 5- 1 2B
5-13 and 5-14
5-14A and 5-14B
5-15 and 5-16
5-16A and 5-16B
5-29 and 5-30
5-30AA and 5-30AB
5-30AB.1 and 5-30AB.2
5-30BB and 5-30BC
5-30CF and 5-30CG
DISTRIBUTION STATEMENT A: Approved for public release; distribution is unlimited.
TM 55-1520-234-23-1
C 71
Remove pages
Insert pages
5-31 and 5-32
5-39 and 5-40
5-43 and 5-44
5-44A/5-44B
5-31 and 5-32
5-39 and 5-40
5-43 and 5-44
5-44A through 5-44C/
5-44D
5-47 and 5-48
5-63 and 5-64
5-66A/ 5-66B
5-83 and 5-84
5-93 and 5-94
5-94A through 5-94D
5-95 and 5-96
5-96A/5-96B
5-132A/5-132B
6-3 and 6-4
6-10C and 6-10D
6-17 and 6-18
6-29 and 6-30
6-38A/6-38B
6-39 and 6-40
6-42A and 6-42B
6-51 through 6-54
6-63 and 6-64
6-64A through 6-64C/6-64D
6-77 and 6-78
6-80C and 6-80D
6-90C and 6-90D
6-94A and 6-94B
6-103 and 6-104
6-112A through 6-112C/
6-112D
7-21 and 7-22
7-50A through 7-50K
8-12A/8-12B
9-39 and 9-40
9-44A and 9-44B
9-67 and 9-68
5-47 and 5-48
5-63 and 5-64
5-66A/ 5-66B
5-83 and 5-84
5-93 and 5-94
5-94A and 5-94B
5-95 and 5-96
5-96A/5-96B
5-132A/5-132B
6-3 and 6-4
6-10C and 6-10D
6-17 and 6-18
6-29 and 6-30
---6-39 and 6-40
6-42A and 6-42B
6-51 through 6-54
6-63 and 6-64
6-64A through 6-64C/6-64D
6-77 and 6-78
6-80C and 6-80D
6-90C and 6-90D
6-94A and 6-94B
6-103 and 6-104
6-112A and 6-112B
7-21 and 7-22
7-50A through 7-50C/7-50D
8-12A/8-12B
9-39 and 9-40
9-44A and 9-44B
9-67 and 9-68
2.
Retain these sheets in front of manual for reference purposes.
TM 55-1520-234-23-1
C 71
By Order of the Secretary of the Army:
GORDON R. SULLIVAN
General, United States Army
Chief of Staff
Official:
MILTON H. HAMILTON
Administrative Assistant to the
Secretary of the Army
02932
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31-E, block no. 0145, requirements for TM 55-1520-234-23-1.
URGENT
TM 55-1520-234-23-1
C 70
CHANGE
NO. 70
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 14 May 1991
}
AVIATION UNIT AND AVIATION INTERMEDIATE
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1.
2.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
5-54A/5-54B
5-55 and 5-56
5-54A/5-54B
5-55 and 5-56
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
CARL E. VUONO
General, United States Army
Chief of Staff
Official:
PATRICIA P. HICKERSON
Colonel, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31-E, block no.
requirements for TM 55-1520-234-23-1.
URGENT
0145, AVUM and AVIM maintenance
URGENT
NOTICE:
THIS CHANGE HAS BEEN PRINTED AND DISTRIBUTED OUT OF SEQUENCE. IT SHOULD BE
INSERTED IN THE MANUAL AND USED. UPON RECEIPT OF THE EARLIER SEQUENCED CHANGE
ENSURE A MORE CURRENT CHANGE PAGE IS NOT REPLACED WITH A LESS CURRENT PAGE.
TM 55-1520-234-23-1
C 69
CHANGE
NO. 69
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 26 OCTOBER 1990
}
AVIATION UNIT AND AVIATION INTERMEDIATE
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1.
2.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
1-71 and 1-72
1-72A/1-72B
1-71 and 1-72
1-72A/1-72B
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
CARL E. VUONO
General, United States Army
Chief of Staff
Official:
THOMAS F. SIKORA
Colonel, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-IS
Helicopter, Attack.
*This TM Change 69 is a result of USAAVSCOM SOF Message 121830Z March 1990.
URGENT
URGENT
NOTICE:
THIS CHANGE HAS BEEN PRINTED AND DISTRIBUTED OUT OF SEQUENCE. IT SHOULD BE
INSERTED IN THE MANUAL AND USED. UPON RECEIPT OF THE EARLIER SEQUENCED CHANGE
ENSURE A MORE CURRENT CHANGE PAGE IS NOT REPLACED WITH A LESS CURRENT PAGE.
TM 55-1520-234-23-1
C 68
CHANGE
NO. 68
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 17 May 1990
}
AVIATION UNIT AND AVIATION INTERMEDIATE
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1.
2.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
5-31 and 5-32
5-55 and 5-56
5-31 and 5-32
5-55 and 5-56
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
CARL E. VUONO
General, United States Army
Chief of Staff
Official:
WILLIAM J. MEEHAN II
Colonel, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-IS
Helicopter, Attack.
*This TM Change 68 is a result of USAAVSCOM SOF Message 031830Z April 1990.
URGENT
URGENT
NOTICE:
THIS CHANGE HAS BEEN PRINTED AND DISTRIBUTED OUT OF SEQUENCE. IT SHOULD BE
INSERTED IN THE MANUAL AND USED. UPON RECEIPT OF THE EARLIER SEQUENCED CHANGE
ENSURE A MORE CURRENT CHANGE PAGE IS NOT REPLACED WITH A LESS CURRENT PAGE.
TM 55-1520-234-23-1
C 67
CHANGE
NO. 67
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 1 March 1990
}
AVIATION UNIT AND AVIATION INTERMEDIATE
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1.
2.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
5-30AA through 5-30AD
5-30AG and 5-30AH
5-30AJ
5-30AK and 5-30AL
5-30AR and 5-30AS
5-30AZ through 5-30BC
5-30AA through 5-30AD
5-30AG and 5-30AH
5-30AJ
5-30AK and 5-30AL
5-30AR and 5-30AS
5-30AZ through 5-30BC
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
CARL E. VUONO
General, United States Army
Chief of Staff
Official:
WILLIAM J. MEEHAN II
Colonel, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-IS
Helicopter, Attack.
URGENT
URGENT
NOTICE:
THIS CHANGE HAS BEEN PRINTED AND DISTRIBUTED OUT OF SEQUENCE. IT SHOULD BE
INSERTED IN THE MANUAL AND USED. UPON RECEIPT OF THE EARLIER SEQUENCED CHANGE
ENSURE A MORE CURRENT CHANGE PAGE IS NOT REPLACED WITH A LESS CURRENT PAGE.
TM 55-1520-234-23-1
C 66
CHANGE
NO. 66
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 28 February 1990
}
AVIATION UNIT AND AVIATION INTERMEDIATE
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1.
2.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
6-10A through 6-l0C/6-10D
6-11 and 6-12
6-10A through 6-10D
6-11 and 6-12
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
CARL E. VUONO
General, United States Army
Chief of Staff
Official:
WILLIAM J. MEEHAN II
Colonel, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-IS
Helicopter, Attack.
URGENT
TM 55-1520-234-23-1
C 65
CHANGE
NO. 65
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 22 November 1990
}
AVIATION UNIT AND AVIATION INTERMEDIATE
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
1-17 and 1-18
1-19 and 1-20
---1-23 and 1-24
1-66A and 1-66B
1-67 and 1-68
1-70A through 1-70E/1-70F
2-100A/2-100B
2-104A and 2-104B
2-104E through 2-104H
2-109 and 2-110
2-126E through 2-126H
---2-132A and 2-132B
2-133 and 2-134
2-142A and 2-142B
3-3 and 3-4
---3-5 and 3-6
4-4C and 4-4D
4-5 through 4-10
4-15 and 4-16
4-42 through 4-44
5-12A/5- 12B
5-13 and 5-14
5- 20A/ 5-20B
5-21 and 5-22
5-29 and 5-30
5-30C through 5-30M
---5-30Q through 5-30T
5-30Y and 5-30Z
---5-30AK and 5-30AL
5-30AX through 5-30BC
5-30BM through 5-30BY
5-30CD through 5-30CG
5-41 and 5-42
1-17 and 1-13
1-19 and 1-20
1-22A/1-22B
1-23 and 1-24
1-66A and 1-66B
1-67 and 1-6S
1-70A through 1-70E/1-70F
2-100A/2-100B
2-104A and 2-104B
2-104E through 2-104H
2-109 and 2-110
2-126E through 2-126H
2-126J through 2-126M
2-132A and 2-132B
2-133 and 2-134
2-142A and 2-142B
3-3 and 3-4
3-4A/3-4B
3-5 and 3-6
4-4C and 4-4D
4-5 through 4-10
4-15 and 4-16
4-42 through 4-44
5-1 2A/5-1 2B
5-13 and 5-14
5-20A/ 5- 20B
5-21 and 5-?2
5-29 and 5-30
5-30A/5-30B through 5-30M
5-30P.1 and 5-30P.2
5-30Q through 5-30T
5-30Y and 5-30Z
5-30Z.1 through 5-30Z.3/5-30Z.4
5-30AK and 5-30AL
5-30AX through 5-30BC
5-30BM through 5-30BY
5-30CD through 5-30CG
5-41 and 5-42
TM 55-1520-234-23-1
C 65
Remove pages
Insert pages
---5-50A/5-50B
5-51 and 5-52
5-66A/5-66B
5-69 and 5-70
5-73 and 5-74
---5-78A and 5-78B
5-79 and 5-80
---5-96A/5-96B
5-101 and 5-102
5-111 and 5-112
5-115 and 5-116
5-116A/5-116B
5-123 and 5-124
5-127 and 5-128
5-131 and 5-132
5-132A/5-132B
5-135 and 5-136
---6-20A through 6-20D
6-39 through 6-42
---6-43 and 6-44
6-63 and 6-64
6-65 and 6-66
6-71 and 6-72
6-73 and 6-74!
6-80C through 6-80F
6-90C through 6-90F
6-93 and 6-94
6-107 and 6-108
6-111 and 6-112
---7-34A and 7-34B
7-43 and 7-44
7-44A and 7-44B
8-18A/8-18B
8-31 and 8-32
8-37 and 8-38
9-1 and 9-2
9-15 and 9-16
9-16A/9-16B
9-23 and 9-24
9-24A and 9-24B
9-72A and 9-72B
5-44A/5-44B
5-50A/5-50B
5-51 and 5-52
5-66A/5-66B
5-69 and 5-70
5-73 and 5-74
5-74A/5-74B
5-78A through 5-78C/5-78D
5-79 and 5-80
5-80A/5-80B
5-96A/5-96B
5-101 and 5-102
5-111 and 5-112
5-115 and 5-116
5-116A/5-116B
5-123 and 5-124
5-127 and 5-128
5-131 and 5-132
5-132A/5-132B
5-135 and 5-136
5-136A/5-136B
6-20A through 6-20D
6-39 through 6-42
6-42A through 6-42E/6-42F
6-43 and 6-44
6-63 and 6-64
6-65 and 6-66
6-71 and 6-72
6-73 and 6-74
6-80C through 6-80G/6-80H
6-90C through 6-90F
6-93 and 6-94
6-107 and 6-108
6-111 and 6-112
6-112A and. 6-112B
7-34A and 7-34B
7-43 and 7-44
7-44A and 7-44B
8-18A/8-18B
8-31 and 8-32
8-37 and Y,-38
9-1 and 9-2
9-15 and 9-16
9-16A/9-16B
9-23 and 9-24
9-24A and 9-24B
9-72A and 9-72B
2.
Retain these sheets in front of manual for reference purposes.
TM 55-1520-234-23-1
C 65
By Order of the Secretary of the Army:
Official:
CARL E. VUONO
General, United States Army
Chief of Staff
THOMAS F. SIKORA
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION :
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-IS Helicopter,
Attack.
URGENT
TM 55-1520-234-23-1
C 64
CHANGE
NO. 64
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 1 February 1990
}
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
2.
Remove pages
Insert pages
1-55 and 1-56
2-101
5-79 and 5-80
1-55 and 1-56
2-101
5-79 and 5-80
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
CARL E. VUONO
General, United States Army
Chief of Staff
Official:
WILLIAM J. MEEHAN II
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-IS
Helicopter, Attack.
URGENT
URGENT
NOTE:
THIS CHANGE HAS BEEN PRINTED AND DISTRIBUTED OUT OF SEQUENCE. IT SHOULD BE
INSERTED IN THE MANUAL AND USED. UPON RECEIPT OF THE EARLIER SEQUENCED CHANGE
ENSURE A MORE CURRENT CHANGE PAGE IS NOT REPLACED WITH A LESS CURRENT PAGE.
TM 55-1520-234-23-1
C 63
CHANGE
NO. 63
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 1 December 1989
}
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-IS (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
2.
Remove pages
Insert pages
1-69 and 1-70
1-71 and 1-72
1-72A/1-72B
6-79 and 6-80
6-80A and 6-80B
1-69 and 1-70
1-17 and 1-72
1-72A/1-72B
6-79 and 6-80
6-80A and 6-80B
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
CARL E.VUONO
General, United States Army
Chief of Staff
Official:
WILLIAM J. MEEHAN II
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-IS
Helicopter, Attack.
URGENT
TM 55-1520-234-23-1
C 62
CHANGE
NO. 62
}
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 31 January 1990
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-IS (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
2.
Remove pages
Insert pages
iii and iv
xiii and xiv
1-35 through 1-38
---------------------
iii and iv
xiii and xiv
1-35 through 1-38
1-38A/1-38B
1-40A/1-40B
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
Official:
CARL E. VUONO
General, United States Army
Chief of Staff
WILLIAM J. MEEHAN II
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-1S
Helicopter, Attack.
URGENT
TM 55-1520-234-23-1
C 61
CHANGE
NO. 61
}
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 1 JUNE 1989
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-IS (MOD)
TM 55-1520-234-23-1, 30 September, 1976, is changed as follows:
1.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
2.
Remove pages
Insert pages
9-44A and 9-44B
9-44A and 9-44B
Retain this page in front of manual for reference purposes.
By Order of the Secretary of the Army:
CARL E. VUONO
General, United States Army
Chief of Staff
Official:
WILLIAM J. MEEHAN II
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-1S
Helicopter, Attack.
URGENT
URGENT
TM 55-1520-234-23-1
C 60
CHANGE
NO. 60
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 15 May 1989
}
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-IS (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
2.
Remove pages
Insert pages
1-67 and 1-68
1-71 and 1-72
6-33 and 6-34
1-67 and 1-68
1-71 and 1-72
6-33 and 6-34
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
CARL E. VUONO
General, United States Army
Chief of Staff
Official:
WILLIAM J. MEEHAN II
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-1S
Helicopter, Attack.
URGENT
URGENT
NOTICE:
THIS CHANGE HAS BEEN PRINTED AND DISTRIBUTED OUT OF SEQUENCE. IT SHOULD BE
INSERTED IN THE MANUAL AND USED. UPON RECEIPT OF THE EARLIER SEQUENCED
CHANGE INSURE A MORE CURRENT CHANGE PAGE IS NOT REPLACED WITH A LESS CURRENT
PAGE.
TM 55-1520-234-23-1
C 59
CHANGE
NO. 59
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 12 April 1989
}
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-IS (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
2.
Remove pages
Insert pages
5-41 and 5-42
7-44A/7-44B
7-45 and 7-46
5-41 and 5-42
7-44A and 7-44B
7-45 and 7-46
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
CARL E. VUONO
General, United States Army
Chief of Staff
Official:
WILLIAM J. MEEHAN II
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-1S
Helicopter, Attack.
URGENT
URGENT
NOTICE:
THIS CHANGE HAS BEEN PRINTED AND DISTRIBUTED OUT OF SEQUENCE. IT SHOULD BE
INSERTED IN THE MANUAL AND USED. UPON RECEIPT OF THE EARLIER SEQUENCED
CHANGE INSURE A MORE CURRENT CHANGE PAGE IS NOT REPLACED WITH A LESS CURRENT
PAGE.
TM 55-1520-234-23-1
C 58
CHANGE
NO. 58
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 28 February 1989
}
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-IS (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
2.
Remove pages
Insert pages
a and b
1-66A and 1-66B
1-71 and 1-72
1-72A/1-72B
5-30G and 5-30H
9-75 and 9-76
9-97 and 9-98
a and b
1-66A and 1-66B
1-71 and 1-72
1-72A/1-72B
5-30G and 5-30H
9-75 and 9-76
9-97 and 9-98
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
CARL E. VUONO
General, United States Army
Chief of Staff
Official:
WILLIAM J. MEEHAN II
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM requirements for AH-1S Helicopter, Attack.
URGENT
TM 55-1520-234-23-1
C 57
CHANGE
NO. 57
}
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 3 February 1989
Aviation Unit and Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-IS (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
2.
Remove pages
Insert pages
ix through xiv
1-11 and 1-12
1-15 through 1-18
9-36A through 9-36C/9-36D
9-37 through 9-42
--9-43 and 9-44
---
ix through xiv
1-11 and 1-12
1-15 through 1-18
9-36A through 9-36D
9-37 through 9-42
9-42A through 9-42Z
9-43 and 9-44
9-44A and 9-44B
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
Official:
CARL E. VUONO
General, United States Army
Chief of Staff
WILLIAM J. MEEHAN II
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM requirements for AH-1S Helicopter, Attack.
TM 55-1520-234-23-1
C 56
CHANGE
NO. 56
}
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 30 December 1988
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-IS (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
c/d
1-29 and 1-30
------1-31 and 1-32
1-41 and 1-42
-----1-55 and 1-56
1-67 through 1-70
1-71 and 1-72
1-72A/1-72B
2-96A and 2-96B
2-110E and 2-110F
3-1 and 3-2
4-7 through 4-10
4-15 and 4-16
4-53 and 4-54
5-14A/5-14B
5-30Y and 5-30Z
5-43 and 5-44
5-57 and 5-58
------5-79 and 5-80
5-87 and 5-88
5-94A and 5-94B
5-113 and 5-114
6-27 and 6-28
6-31 and 6-32
-----6-33 and 6-34
6-34A/6-34B
6-77 and 6-78
6-92A and 6-92B
6-93 and 6-94
c and d
1-29 and 1-30
1-30A/1-30B
1-31 and 1-32
1-41 and 1-42
1-42A/1-42B
1-55 and 1-56
1-67 through 1-70
1-71 and 1-72
1-72A/1-72B
2-96A and 2-96B
2-110E and 2-110F
3-1 and 3-2
4-7 and 4-10
4-15 and 4-16
4-53 and 4-54
5-14A/5-14B
5-30Y and 5-30Z
5-43 and 5-44
5-57 and 5-58
5-58A/5-58B
5-79 and 5-80
5-87 and 5-88
5-94A and 5-94B
5-113 and 5-114
6-27 and 6-28
6-31 and 6-32
6-32A/6-32B
6-33 and 6-34
6-34A/6-34B
6-77 and 6-78
6-92A and 6-92B
6-93 and 6-94
TM 55-1520-234-23-1
C 56
2.
Remove pages
Insert pages
7-25 and 7-26
7-37 through 7-38
7-39 and 7-42
7-42A/7-42B
7-51 and 7-52
8-15 and 8-16
8-18A/8-18B
9-43 through 9-48
7-25 and 7-26
7-37 through 7-38
7-39 and 7-42
7-42A/7-42B
7-51 and 7-52
8-15 and 8-16
8-18A/8-18B
9-43 through 9-48
Retain these sheets in front of manual for reference purposes.
By Order of the Secretary of the Army:
Official:
CARL E. VUONO
General, United States Army
Chief of Staff
WILLIAM J. MEEHAN II
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Requirements for AH-1S Helicopter, Attack.
URGENT
NOTICE:
THIS CHANGE HAS BEEN PRINTED AND DISTRIBUTED OUT OF SEQUENCE. IT SHOULD BE
INSERTED IN THE MANUAL AND USED. UPON RECEIPT OF THE EARLIER SEQUENCED
CHANGE INSURE A MORE CURRENT CHANGE PAGE IS NOT REPLACED WITH A LESS CURRENT
PAGE.
TM 55-1520-234-23-1
C 55
CHANGE
NO. 55
}
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 30 September 1988
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-1S(MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
2.
Remove pages
Insert pages
5-78A and 5-78B
5-78A and 5-78B
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
CARL E. VUONO
General, United States Army
Chief of Staff
Official:
R. L. DILWORTH
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Requirements for AH-1S Helicopter, Attack.
URGENT
TM 55-1520-234-23-1
C 54
CHANGE
NO. 54
}
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 5 October 1988
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-1S(MOD)
TM 55-1520-234-23-1, 30 September, 1976, is changed as follows:
1.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
iii thru xii
1-7 and 1-8
1-11 thru 1-18
1-18A/1-18B
1-19 and 1-20
1-23 and 1-24
1-37 and 1-38
1-48A/1-48B
1-50C/1-50D
1-59 and 1-60
1-67 and 1-68
1-70C and 1-70D
1-71 and 1-72
1-72A/1-72B
2-1 and 2-2
2-7 and 2-8
2-73 and 2-74
2-100A/2-100B
2-143 and 2-144
4-5 and 4-6
4-15 and 4-16
5-5 and 5-6
5-37 and 5-38
5-65 and 5-66
5-78A and 5-78B
5-79 and 5-80
5-97 and 5-98
5-101 and 5-102
5-121 thru 5-124
5-135 and 5-136
6-24C and 6-24D
6-33 and 6-34
6-63 and 6-64
6-65 thru 6-70
iii thru xiv
1-7 and 1-8
1-11 thru 1-18
1-18A/1-18B
1-19 and 1-20
1-23 and 1-24
1-37 and 1-38
1-48A/1-48B
1-50C/1-50D
1-59 and 1-60
1-67 and 1-68
1-70C and 1-70D
1-71 and 1-72
1-72A/1-72B
2-1 and 2-2
2-7 and 2-8
2-73 and 2-74
2-100A/2-100B
2-143 and 2-144
4-5 and 4-6
4-15 and 4-16
5-5 and 5-6
5-37 and 5-38
5-65 and 5-66
5-78A and 5-78B
5-79 and 5-80
5-97 and 5-98
5-101 and 5-102
5-121 thru 5-124
5-135 and 5-136
6-24C and 6-24D
6-33 and 6-34
6-63 and 6-64
6-65 thru 6-70
TM 55-1520-234-23-1
C 54
2.
Remove pages
Insert pages
6-71 and 6-72
6-73 and 6-74
6-77 and 6-73
6-90C thru 6-90F
6-93 and 6-94
6-94A and 6-94B
6-103 and 6-104
9-1 and 9-2
9-23 and 9-24
6-71 and 6-72
6-73 and 6-74
6-77 and 6-78
6-90C thru 6-90F
6-93 and 6-94
6-94A and 6-94B
6-103 and 6-104
9-1 and 9-2
9-23 and 9-24
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
Official:
CARL E. VUONO
General, United States Army
Chief of Staff
WILLIAM J. MEEHAN II
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-1S
Helicopter, Attack.
URGENT
TM 55-1520-234-23-1
C 53
CHANGE
NO. 53
}
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 29 July 1988
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-1S(MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1.
Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
2.
Remove pages
Insert pages
1-29 and 1-30
1-29 and 1-30
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
CARL E. VUONO
General, United States Army
Chief of Staff
Official:
R. L. DILWORTH
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Requirements for AH-1S Helicopter, Attack.
URGENT
URGENT
TM 55-1520-234-23-1
C 52
CHANGE
NO. 52
}
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 17 June 1988
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-1S(MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1.
Remove and insert pages as indicated below. New or changed text material is Indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
2.
Remove pages
Insert pages
1-29 and 1-30
1-29 and 1-30
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
CARL E. VUONO
General, United States Army
Chief of Staff
Official:
R. L. DILWORTH
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Requirements for AH-1S Helicopter, Attack.
URGENT
URGENT
TM 55-1520-234-23-1
C 51
CHANGE
NO. 51
}
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 6 April 1988
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-1S(MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1.
Remove and insert pages as indicated below. New or changed text material is Indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
2.
Remove pages
Insert pages
6-13
6-13 and 6-14
Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
CARL E. VUONO
General, United States Army
Chief of Staff
Official:
R. L. DILWORTH
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Requirements for AH-1S Helicopter, Attack.
URGENT
URGENT
TM 55-1520-234-23-1
C 50
CHANGE
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 3 March 1988
NO. 50
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand
Insert pages
1-29 and 1-30
1-55 and 1-56
1-56A/1-56B
5-29 and 5-30
--5-30CF and 5-30CG
5-67 and 5-68
Remove pages
1-29 and 1-30
1-55 and 1-56
--5-29 and 5-30
5-30A and 5-30B
5-30CF/5-30CG
5-67 and 5-68
2. Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
CARL E. VUONO
General, United States Army
Chief of Staff
Official:
R. L. DILWORTH
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-1S
Helicopter, Attack.
URGENT
URGENT
NOTICE: THIS CHANCE HAS BEEN PRINTED AND DISTRIBUTED OUT OF SEQUENCE. IT SHOULD BE
INSERTED IN THE MANUAL AND USED UPON RECEIPT OF THE EARLIER SEQUENCED CHANGE
INSURE A MORE CURRENT CHANGE PACE IS NOT REPLACED WITH A LESS CURRENT PAGE.
TM 55-1520-234-23-1
C 49
CHANGE
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 7 December 1987
NO. 49
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
5-29 and 5-30
5-30C through 5-30Z
5-30AA through 5-30AZ
5-30BA through 5-30BN
----5-103 and 5-104
5-29 and 5-30
5-30C through 5-30Z
5-30AA through 5-30AZ
5-30BA through 5-30BN
5-30BP through 5-30BZ
5-30CA through 5-30CF/5-30CG
5-103 and 5-104
2. Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
CARL E. VUONO
General, United States Army
Chief of Staff
Official:
R. L. DILWORTH
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-1S
Helicopter, Attack.
URGENT
TM 55-1520-234-23-1
C 48
CHANGE
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 7 December 1987
NO. 48
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
a and b
1-11 through 1-16
1-31 and 1-32
1-32A/1-32B
1-53 and 1-54
1-67 and 1-68
1-71 and 1-72
1-72A/1-72B
2-49 and 2-50
2-57 and 2-58
--2-95 and 2-96
2-96A and 2-96B
2-133 and 2-134
2-139 and 2-140
4-49 through 4-52
5-2A and 5-2B
5-10A and 5-10B
5-14A/5-14B
5-15 and 5-16
5-25 and 5-26
5-28A and 5-28B
5-30BH.1 and 5-30BH.2
5-30BN and 5-30BP
5-37 through 5-40
5-53 and 5-54
--5-81 through 5-84
5-89 and 5-90
5-115 and 5-116
6-9 and 6-10
6-13 and 6-14
6-16C and 6-16D
6-17 and 6-18
6-20A through 6-20D
a and b
1-11 through 1-16
1-31 and 1-32
1-32A/1-32B
1-53 and 1-54
1-67 and 1-68
1-71 and 1-72
1-72A/1-72B
2-49 and 2-50
2-57 and 2-58
2-58A/2-58B
2-95 and 2-96
2-96A and 2-96B
2-133 and 2-134
2-139 and 2-140
4-49 through 4-52
5-2A and 5-2B
5-10A and 5-10B
5-14A/5-14B
5-15 and 5-16
5-25 and 5-26
5-28A and 5-23B
5-30BH.1 and 5-30BH.2
5-30BN and 5-30BP
5-37 through 5-40
5-53 and 5-54
5-54A/5-54B
5-81 through 5-84
5-89 and 5-90
5-115 and 5-116
6-9 and 6-10
6-13
6-16C and 6-16D
6-17 and 6-18
6-20A through 6-20D
TM 55-1520-234-23-1
C 48
Remove pages
Insert pages
6-43 and 6-44
6-61 and 6-62
--6-90C and 6-90D
6-101 and 6-102
6-109 and 6-110
6-113 and 6-114
6-129 and 6-130
7-9 and 7-10
7-25 and 7-26
8-7 and 8-8
8-16A and 8-16B
8-19 and 8-20
9-24A and 9-24B
9-34A/9-34B
9-57 and 9-58
9-58A/9-58B
9-59 and 9-60
9-60A/9-60B
9-85 and 9-86
6-43 and 6-44
6-61 and 6-62
6-62A/6-62B
6-90C and 6-90D
6-101 and 6-102
6-109 and 6-110
6-113 and 6-114
6-129 and 6-130
7-9 and 7-10
7-25 and 7-26
8-7 and 8-8
8-16A and 8-16B
8-19 and 8-20
9-24A and 9-24B
9-34A/9-34B
9-57 and 9-58
9-58A through 9-58D
9-59 and 9-60
9-60A through 9-60C/9-60D
9-85 and 9-86
2. Retain these sheets in front of manual for reference purposes.
By Order of the Secretary of the Army:
Official:
CARL E. VUONO
General, United States Army
Chief of Staff
R. L. DILWORTH
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-1S
Helicopter, Attack.
URGENT
TM 55-1520-234-23-1
C 47
CHANGE
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 27 May 1987
NO. 47
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
1-66A and 1-66B
1-66A and 1-66B
2. Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
JOHN A. WICKHAM, JR.
General, United States Army
Chief of Staff
Official:
R. L. DILWORTH
Brigadier General, United States Army
The A Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-1S
Helicopter, Attack.
URGENT
URGENT
NOTICE: THIS CHANCE HAS BEEN PRINTED AND DISTRIBUTED OUT OF SEQUENCE. IT SHOULD BE
INSERTED IN THE MANUAL AND USED UPON RECEIPT OF THE EARLIER SEQUENCED CHANGE
INSURE A MORE CURRENT CHANGE PACE IS NOT REPLACED WITH A LESS CURRENT PAGE.
TM 55-1520-234-23-1
C 46
CHANGE
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 6 April 1987
NO. 46
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
1-66A and 1-66B
1-71 and 1-72
1-72A/(1-72B)
1-66A and 1-66B
1-71 and 1-72
1-72A/(1-72B)
2. Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
JOHN A. WICKHAM, JR
General, United States Army
Chief of Staff
Official:
R. L. DILWORTH
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-1S
Helicopter, Attack.
URGENT
URGENT
NOTICE: THIS CHANCE HAS BEEN PRINTED AND DISTRIBUTED OUT OF SEQUENCE. IT SHOULD BE
INSERTED IN THE MANUAL AND USED UPON RECEIPT OF THE EARLIER SEQUENCED CHANGE
INSURE A MORE CURRENT CHANGE PACE IS NOT REPLACED WITH A LESS CURRENT PAGE.
TM 55-1520-234-23-1
C 45
CHANGE
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 6 March 1987
NO. 45
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
1-70C and 1-70D
1-70C and 1-70D
2. Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
JOHN A. WICKHAM, JR
General, United States Army
Chief of Staff
Official:
R. L. DILWORTH
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-1S
Helicopter, Attack (MOD).
URGENT
TM 55-1520-234-23-1
C 44
CHANGE
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 2 March 1987
NO. 44
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
iii and iv
1-17 and 1-18
--1-65 and 1-66
2-3 through 2-6
2-88A/2-88B
2-96A through 2-96D
2-97 and 2-98
2-101 through 2-104
--2-105 and 2-106
iii and iv
1-17 and 1-18
1-64A/1-64B
1-65 and 1-66
2-3 through 2-6
2-88A/2-88B
2-96A
2-97 and 2-98
2-101 and 2-102
2-104A through 2-104H
2-105 and 2-106
2. Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
Official:
CARL E. VUONO
General, United States Army
Chief of Staff
R. L. DILWORTH
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-1S
Helicopter, Attack.
URGENT
NOTICE: THIS CHANCE HAS BEEN PRINTED AND DISTRIBUTED OUT OF SEQUENCE. IT SHOULD BE
INSERTED IN THE MANUAL AND USED UPON RECEIPT OF THE EARLIER SEQUENCED CHANGE
INSURE A MORE CURRENT CHANGE PACE IS NOT REPLACED WITH A LESS CURRENT PAGE.
TM 55-1520-234-23-1
C 43
CHANGE
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 9 February 1987
NO. 43
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
1-71 and 1-72
1-71 and 1-72
2. Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
JOHN A. WICKHAM, JR
General, United States Army
Chief of Staff
Official:
R. L. DILWORTH
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-1S
Helicopter, Attack.
URGENT
TM 55-1520-234-23-1
C 42
CHANGE
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 9 February 1987
NO. 42
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
v and vi
ix and x
1-11 and 1-12
1-15 and 1-16
1-23 and 1-24
1-29 and 1-30
1-41 and 1-42
1-48.1/1-48.2
1-57 and 1-58
1-66A and 1-66B
1-69 and 1-70
1-70E/1-70F
1-71 and 1-72
2-3 through 2-6
2-79 and 2-80
2-96A/2-96B
2-131 and 2-132
--2-143 and 2-144
4-4C and 4-4D
4-53 and 4-54
5-5 and 5-6
5-30E and 5-30F
5-30Q and 5-30R
--5-30AL and 5-30AM
5-37 and 5-38
5-41 and 5-42
5-57 and 5-58
5-79 and 5-80
5-103 and 5-104
6-10A and 6-10B
6-17 and 6-18
v and vi
ix and x
1-11 and 1-12
1-15 and 1-16
1-23 and 1-24
1-29 and 1-30
1-41 and 1-42
1-48A/1-48B
1-57 and 1-58
1-66A and 1-66B
1-69 and 1-70
1-70E/1-70F
1-71 and 1-72
2-3 through 2-6
2-79 and 2-80
2-96A through 2-96D
2-131 and 2-132
2-132A and 2-132B
2-143 and 2-144
4-4C and 4-4D
4-53 and 4-54
5-5 and 5-6
5-30E and 5-30F
5-30Q and 5-30R
5-30AK.1/5-30AK.2
5-30AL and 5-30AM
5-30AM.1/5-30AM.2
5-37 and 5-38
5-41 and 5-42
5-57 and 5-58
5-79 and 5-80
5-103 and 5-104
6-10A and 6-10B
6-17 and 6-18
TM 55-1520-234-23-1
C 42
Remove pages
Insert pages
6-57 and 6-58
6-67 through 6-70
6-79 and 6-80
6-80A and 6-80B
6-90E and 6-90F
6-93 and 6-94
6-99 through 6-104
7-26A and 7-26B
7-27 and 7-28
7-41 and 7-42
7-42A/7-42B
7-43 and 7-44
7-44A/7-44B
9-33 and 9-34
8-34A/8-34B
9-85 and 9-86
6-57 and 6-58
6-67 through 6-70
6-79 and 6-80
6-80A and 6-80B
6-90E and 6-90F
6-93 and 6-94
6-99 through 6-104
7-26A and 7-26B
7-27 and 7-28
7-41 and 7-42
7-42A/7-42B
7-43 and 7-44
7-44A/7-44B
8-33 and 8-34
8-34A/8-34B
9-85 and 9-86
2. Retain these sheets in front of manual for reference purposes.
By Order of the Secretary of the Army:
Official:
JOHN A. WICKHAM, JR
General, United States Army
Chief of Staff
R. L. DILWORTH
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-1S
Helicopter, Attack (MOD).
URGENT
TM 55-1520-234-23-1
C 41
CHANGE
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 21 November 1986
NO. 41
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
1-51 and 1-52
5-23 and 5-24
---
1-51 and 1-52
5-23 and 5-24
5-24A/5-24B
2. Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
JOHN A. WICKHAM, JR
General, United States Army
Chief of Staff
Official:
R. L. DILWORTH
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-1S
Helicopter, Attack (MOD).
URGENT
URGENT
TM 55-1520-234-23-1
C 40
CHANGE
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 18 April 1986
NO. 40
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
1-70C through 1-70E/1-70F
1-70C through 1-70E/1-70F
2. Retain this sheet in front of manual for reference purposes.
By Order of the Secretary of the Army:
JOHN A. WICKHAM, JR
General, United States Army
Chief of Staff
Official:
R. L. DILWORTH
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-1S
Helicopter, Attack (MOD).
URGENT
URGENT
NOTICE: THIS CHANGE HAS BEEN PRINTED AND DISTRIBUTED OUT OF SEQUENCE. IT SHOULD BE
INSERTED IN THE MANUAL AND USED. UPON RECEIPT OF THE EARLIER SEQUENCED CHANGE
INSURE A MORE CURRENT CHANGE PAGE IS NOT REPLACED WITH A LESS CURRENT PAGE.
TM 55-1520-234-23-1
C 39
CHANGE
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 21 February 1986
NO. 39
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
c/d
1-43 and 1-44
1-67 and 1-68
5-53 and 5-54
5-73 and 5-74
c/d
1-43 and 1-44
1-67 and 1-68
5-53 and 5-54
5-73 and 5-74
By Order of the Secretary of the Army:
JOHN A. WICKHAM, JR
General, United States Army
Chief of Staff
Official:
MILDRED E. HEDBERG
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-1S
Helicopter, Attack (MOD).
URGENT
TM 55-1520-234-23-1
C 38
CHANGE
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D.C., 20 February 1986
NO. 38
Aviation Unit and Aviation Intermediate
Maintenance Manual
HELICOPTER, ATTACK AH-1S (MOD)
TM 55-1520-234-23-1, 30 September 1976, is changed as follows:
1. Remove and insert pages as indicated below. New or changed text material is indicated by a vertical bar in the
margin. An illustration change is indicated by a miniature pointing hand.
Remove pages
Insert pages
a through c/d
v through viii
xi and xii
P-1 and P-2
1-9 and 1-10
1-10A and 1-10B
1-11 through 1-18
1-18A/ 1-18B
1-19 through 1-22
1-29 and B30
1-33 and 1-34
1-43 and 1-44
1-57 and 1-58
1-61 and 1-62
1-66A and 1-66B
1-67 through 1-70
1-70A and 1-70B
1-71 and 1-72
2-75 and 2-76
2-93 through 2-96
2-100A/2-100B
2-131 through 2-134
2-139 and 2-140
2-173 through 2-178
4-5 and 4-6
4-15 and 4-16
4-19 and 4-20
4-47 and 4-48
4-48A/4-48B
4-51 and 4-52
5-10A and 5-10B
5-12A/5-12B
5-13 and 5-14
5-14A/5-14B
5-15 and 5-16
5-19 and 5-20
6-27 and 5-28
5-28A and 5-28B
5-30E through 5-30K
a through c/d
v through vii
xi and xii
P-1 and P-2
1-9 and 1-10
1-10A and 1-10B
1-11 through 1-18
1-18A/1-18B
1-19 through 1-22
1-29 and 1-30
1-33 and 1-34
1-43 and 1-44
1-57 and 1-58
1-61 and 1-62
1-66A and 1-66B
1-67 through 1-70
1-70A and 1-70B
1-71 and 1-72
2-75 and 2-76
2-93 through 2-96
2-100A/2-100B
2-131 through 2-134
2-139 and 2-140
2-173 through 2-178
4-5 and 4-6
4-15 and 4-16
4-19 and 4-20
4-47 and 4-48
4-48A/4-48B
4-51 and 4-52
5-10A and 5-10B
5-12A/5-12B
5-13 and 5-14
5-14A/5-14B
5-15 and 5-16
5-19 and 5-20
5-27 and 5-28
5-28A and 5-28B
5-30E through 5-30K
TM 66-1520-234-23-1
C38
Remove pages
Insert pages
5-30S and 5 30T
5-30Y and 5-30Z
5-30AA through 5-30AF
5-30AS and 5-30AT
--5-30AW and 5-30AX
--5-30AY and 5-30AZ
5-30BE through 5-30BH
--5-37 and 5-38
5-41 and 5-42
5-49 and 5-50
5-51 through 5-54
5-57 and 5-58
5-63 and 5-64
5-65 and 5-66
5-69 and 5-70
5-77 and 5-78
--5-79 and 5-80
5-94A and 5-94B
5-99 through 5-102
5-115 and 5-116
5-133 through 5-136
6-19 and 6-20
--6-23 and 6-24
6-25 and 6-26
6-39 through 6-42
6-55 and 6-56
6-56A/6-56B
6-77 and 6-78
6 79 and 6-80
6-80E/6-80F
6-91 and 6-92
6-103 and 6-104
6-107 and 6-108
6-109 and 6-110
7-15 and 7-16
7-31 and 7 32
7-32A/7-32B
7-39 and 740
7-42A/7-42B
7-43 and 744
8-31 and 8-32
8-34A/8-34B
8-35 and 8-36
9-11 and 9-12
9-24A and 9-24B
5-30S and 5-30T
5-30Y and 5-30Z
5-30AA through 5-30AF
5-30AS and 5-30AT
5-30AT.1/5-30AT.2
5-30AW and 5-30AX
5-30AX.1/5-30AX.2
5-30AY and 5-30AZ
5-30BE through 5-30BH
5-30BH.1 through 5-30BH.7
5-37 and 5-38
5-41 and 542
5-49 and 5-50
5-51 through 5-54
5-57 and 5-58
5-63 and 5-64
5-65 and 5-66
5-69 and 5-70
5-77 and 5-78
5-78A and 5-78B
5-79 and 5-80
5-94A and 5-94B
5-99 through 5-102
5-115 through 5-116
5-133 through 5-136
6-19 and 6-20
6-20A through 6-20E/6-20!
6-23 and 6-24
6-25 and 6-26
6-39 through 6-42
6-55 and 6-56
6-56A and 6-56B
6-77 and 6-78
6-79 and 6-80
6-80E and 6-80F
6-91 and 6-92
6-103 and 6-104
6-108 and 6-107
6-109 and 6-110
7-15 and 7-16
7-31 and 7-32
7-32A/7-32B
7-39 and 740
7-42A/7-42B
7-43 and 744
8-31 and 8-32
8-34A/8-34B
8-35 and 8-36
9-11 and 9-12
9-24A and 9-24B
2. Retain these sheets in front of manual for reference purposes.
TM 55-1520-234 23-1
By Order of the Secretary of the Army:
Official:
JOHN A. WICKHAM, JR
General, United States Army
Chief of Staff
MILDRED E. HEDBERG
Brigadier General, United States Army
The Adjutant General
DISTRIBUTION:
To be distributed in accordance with DA Form 12-31, AVUM and AVIM Maintenance requirements for AH-1S
Helicopter, Attack (MOD).
WARNING PAGE
TM 55-1520-234-23
WARNING
Personnel performing instructions involving operations, procedures, and practices
which are included, or implied in this technical manual,
shall observe the following instructions.
Disregard of these warnings and precautionary information can cause serious injury,
DEATH
or an aborted mission.
Starting and Operation
of the helicopter will be performed only by
authorized personnel in accordance with AR95-1.
HIGH VOLTAGE
Turn off all power switches before making electrical connections or disconnections.
Serious burns and electrical shock can result from contact
with exposed electrical wires or connectors.
RADIATION HAZARD
Self-luminous dials contain radio active materials.
If such an instrument is broken or becomes unsealed, avoid personal contact.
Use forceps or gloves made of rubber or polyethylene to pick up contaminated material.
Place material and gloves in a plastic bag.
Seal bag and dispose of it as radio active waste
in accordance with AR755-15 and TM 3-261 (Refer to TB 55-1500-314-24).
Repair procedures shall conform to requirements in AR700-52.
DANGEROUS CHEMICALS
Exposure to high concentration of fire extinguishing agents
can cause severe irritation of eyes or nose.
Corrosive Battery Electrolyte (Potassium Hydroxide).
Wear rubber gloves, apron, and face shield when handling leaking batteries.
If potassium hydroxide is spilled on clothing, or other material
wash immediately with clean water.
If spilled on personnel, immediately start flushing
the affected area with clean water.
Continue washing until medical assistance arrives.
Use solvent in a well ventilated area.
Do not inhaled vapors, or allow to come in contact with skin or eyes.
Observe proper fire prevention rules.
ARMAMENT
When working on, or near an armed helicopter, take all possible
precautions to avoid accidental firing or armament.
Personnel shall not occupy possible firing pattern in front of or up to
20 feet behind rocket pods.
Munitions shall be handled by authorized personnel only.
All weapons shall be dry-fired. Only dummy ammunition with
smooth cases like live ammunition shall be used.
Change 58
a
WARNING PAGE
TM 55-1520-234-23
HYDRAULIC FLUID, FUEL AND OIL
Lubricating oil used in engine, transmission and gear boxes
may cause a skin rash if prolonged contact is allowed.
When handling fuel, observe precautions and procedures in TM 10-1101.
Prolonged contact with hydraulic fluid will cause burns.
When handling hydraulic fluid (MIL-H-83282), table 1-3, item 73A,
observe the following:
Prolonged contact with liquid or mist can irritate eyes and skin.
After any prolonged contact with skin, immediately wash contacted area with soap and water.
If liquid contacts eyes, flush them immediately with clear water.
If liquid is swallowed, do not induce vomiting;
get immediate medical attention.
Wear rubber gloves when handling liquid.
If prolonged contact with mist is likely, wear an appropriate respirator.
When fluid is decomposed by heating, toxic gases are released.
JETTISON
All ground safety pints must be removed before night.
Failure to do so will prevent emergency jettison of stores.
JETTISON
Jettison circuit may be activated with battery switch
OFF and pilot's wing stores jettison circuit breaker
pulled. For positive deactivation of jettison circuit,
open both the pilots wing stores jettison circuit
breaker and the jettison circuit breaker located in
the aft electrical compartment. Serious injury can
result from accidental ground jettison.
SANDING DUST
Sanding on glass cloth reinforced laminated produces
fine dust that may cause skin irritations.
Observe necessary protective measures.
TRANSMISSION LEVELING
Do not attempt to level transmission with "Jacks Only"
Hoist must be used in conjunction with jacks while lifting transmission.
EXTERNAL STORES
Prior to any aircraft maintenance functions that
require external stores be removed, ejector cartridge shall be removed.
Remove jettison cartridges from pylon stores ejection
device prior to placing aircraft in a hangar, to prevent
injury to personnel and damage to equipment.
Exception: Removal is not necessary when aircraft
is to be placed in hangar for short-term, providing
both circuit breakers are open, ground safety pins installed,
jettison switches are OFF, and warning signs indicate that aircraft
has an armed jettison system.
Change 58
b
WARNING PAGE
TM 55-1520-234-23
CANOPY REMOVAL SYSTEM
Ground safety pins must be installed in
pilot's and gunner's arming/firing handles
of canopy removal system whenever the
helicopter is on the ground. Pins should
be installed by crew.
EPOXY BASED ADHESIVE
Epoxy based adhesive, P/N EA934,
contains an asbestos filler which could
be inhaled or ingested during grinding,
cutting, or sanding operations on
cured epoxy material.
TOOLS
Use only chrome plated steel or unplated
steel tools for disassembly or reassembly
procedures described in this manual.
Use of cadmium or zinc plated tools
is not permitted.
GROUNDING
All aircraft parked outside will be
grounded and bonded, in accordance
with FM 1-500, to the aerospace ground
equipment while servicing, i.e., fueling
or defueling, arming (ammunition or
explosives), oxygen, hydraulic fluids or
any flammable liquids. Grounding is
not necessary for aircraft parked
outside unless one of the above is
being accomplished.
INSPECTION OF REMOVED
COMPONENT
When components are being removed
from an aircraft, all inspections required
by the next phase maintenance
inspection must be accomplished prior
to either immediate re-use or
storage. Upon installation, the
component will be inspected in
accordance with the current phase
(either that phase the receiving aircraft
is in or if in between phase, the last
phase performed). This will ensure that
a re-used component will not overfly any
PM inspections, and that it will be
properly interfaced with the receiving
aircraft phase sequence.
Change 56
c
WARNING PAGE
TM 55-1520-234-23
CLEANING HYDRAULIC
COMPONENTS
The use of any alcohol in cleaning
components which contact hydraulic
fluids is prohibited. Formation of a
polymeric residue can result, which
could impair mechanical operation of
the component.
d
Change 56
TM 55-1520-234-23
TECHNICAL MANUAL
HEADQUARTERS
DEPARTMENT OF THE ARMY
WASHINGTON, D. C. 30 September 1976
OPERATORS MANUAL
HELICOPTER. ATTACK AH-1S
TABLE OF CONTENTS
Paragraph
Page
1-1
1-18
1-1
1-28
1-20
1-29
1-32
1-32
1-43
1-70
2-2
2-45
2-53
2-63
2-66
2-67
-
2-1
2-105
2-135
2-142
2-149
2-173
2-175
POWER PLANT
Power plant .............................................................................
Cooling system ........................................................................
Air induction system ................................................................
Exhaust system .......................................................................
Oil system ...............................................................................
Ignition system (not applicable)................................................
Power controls.........................................................................
4-1
4-6
4-7
4-11
4-14
-4-21
4-1
4-6
4-6
4-20
4-32
4-45
4-45
CHAPTER 5
Section
I
II
ROTORS
Main rotor system ....................................................................
Tail rotor system......................................................................
5-2
5.10
5-1
5-95
CHAPTER 6
Section
I
II
III
IV
V
VI
VII
VIII
DRIVE TRAIN SYSTEM
Drive train................................................................................
Main driveshaft........................................................................
Clutches (not applicable) .........................................................
Main transmission....................................................................
Tail rotor driveshaft..................................................................
Intermediate gearbox...............................................................
Tail rotor gearbox ....................................................................
Transmission oil system...........................................................
6-1
6-7
-6-9
6-23
6-25
6-27
6-30
6-1
6-8
6-21
6-21
6-75
6-80
6-91
6-105
CHAPTER 1
Section
I
II
III
IV
V
AIRCRAFT GENERAL
Servicing . ...............................................................................
Lubrication...............................................................................
Handling, jacking, mooring, hoisting, and sling
loading.....................................................................................
Inspection requirements ..........................................................
Overhaul and retirement schedule ...........................................
CHAPTER 2
Section
I
II
III
IV
V
VI
VII
AIRFRAME
Fuselage .................................................................................
Tailboom .................................................................................
Pylon . .....................................................................................
Wing........................................................................................
Extrusion chart ........................................................................
Corrosion control .....................................................................
Structural repair material .........................................................
CHAPTER 3
ALIGHTING GEAR
CHAPTER 4
Section
I
II
II
IV
V
VI
VII
i
TM 55-1520-234-23
Paragraph
CHAPTER 7
HYDRAULIC AND PNEUMATIC SYSTEMS
CHAPTER 8
Section
I
II
III
IV
V
VI
CHAPTER 9
Section
I
II
III
IV
V
VI
VII
Page
7-1
7-1
INSTRUMENTS SYSTEMS
Instrument maintenance ......................................................
Engine instruments ..............................................................
Flight Instruments ................................................................
Navigation instruments ........................................................
Miscellaneous instruments ...................................................
Panels ..................................................................................
8-1
8-3
8-25
8-35
8-39
8-53
8-1
8-7
8-20
8-27
8-28A
8-40
ELECTRICAL SYSTEMS
Electrical system maintenance ............................................
DC current power distribution system....................................
AC current power distribution system ...................................
Starting system.....................................................................
Ignition system .....................................................................
Lighting provisions................................................................
Miscellaneous equipment .....................................................
9-1
9-7
9-24
9-38
9-41
9-44
9-60
9-1
9-14
9-25
9-32
9-34
9-34A
9-61
TM 55-1520-234-23
LIST OF ILLUSTRATIONS
Figure
P-1
1-1
1-1A
1-2
1-3
1-4
1-5
1-6
1-7
1-7A
1-8
1-9
2-1
2-2
2-3
2-4
2-5
2-6
2-7
2-8
2-9
2-10
2-11
2-12
2-13
2-14
2-15
2-16
2-17
2-18
2-19
2-20
2-21
2-22
2-23
2-24
2-25
2-26
2-27
2-28
2-29
2-30
2-31
2-32
2-33
2-34
Title
Page
AH-1S Helicopter ..............................................................................................................
Servicing Points Diagram ...................................................................................................
Closed Circuit Refueling System .......................................................................................
Lubrication .........................................................................................................................
Ground Handling Diagram .................................................................................................
Work Aid For Towing Ground Handling Gear .....................................................................
Jacking and Mooring Fittings ..............................................................................................
Covers Diagram .................................................................................................................
Mooring Diagram ...............................................................................................................
AH-1 Paved Surface Mooring Configuration .......................................................................
Maintenance Hoist T101520 ...............................................................................................
Inspection Area Diagram ....................................................................................................
Fuselage, Wings, Pylon Mount and Tailboom ....................................................................
Structural Panels ...............................................................................................................
Non-structural Panels, Doors and Fairings ..........................................................................
Primary Structure Caps .................................... .................................................................
Pilot and Gunner Floor Panels ................................ ..........................................................
Bulkhead at Station 93.0 ................................... ................................................................
Bulkhead at Stations 148.5 and 171.61 ...............................................................................
Bulkheads at Station 184.5 and 213.94 ............................ ..................................................
Bulkheads at Stations 250.0 and 268.5 ...............................................................................
Right and Left Main Beam Panels at Station 148.5 to 186.25 ................ .............................
Right and Left Main Beam Panels at Station 213.94 to 250.0 ................ .............................
Panel at Forward Fuel Cell at Right Side and Gunners Floor ..............................................
Left and Right Beam Panels at Station 250.0 To Boom Station 41.32 ........... ...................
Ammo Floor, Support Panel and Forward Fuel Cell Panel at Station 213.94 .......................
Forward Fuel Cell Floor and Lower Panel at Station 186.25 to 213.94.................................
Lower Aft Fuel Cell Panel and Bottom Panel at Station 250.0 To Boom Station 41.32 ........
Engine Deck Installation at Station 213.94 to 298.75 ..........................................................
Forward Fuel Cell Panels and Main Beam Panels at Station 186.25 to 214.0 ......................
Vertical Fin Honeycomb Panels. .........................................................................................
Type A Damage-Body Panel Repairs........ ..........................................................................
Type B Damage-Body Panel Repairs ..................................................................................
Type C Damage-Body Panel Repairs ..................................................................................
Type D Damage-Body Panel Repairs .................................................................................
Edge Repair For Honeycomb Panel With Fiberglas Skin Opposite Titanium ......................
Edge Repair For Honeycomb Panel With Aluminum Alloy Skin
Opposite Aluminum Alloy With Fiberglas Edging .............................................................
Edge Repair For Honeycomb Panel With Aluminum Alloy Skin
Opposite Aluminum Alloy With Fiberglas Finish at Channel .............................................
Typical Rivet Pattern For Channel Section Repair...............................................................
Edge Repair on Vertical Fin ...............................................................................................
Injection Type Fastener (Insert) in Vertical Fin Panel ..........................................................
Potted Type, Injection Type, and Grommet Type Fastener (Inserts) ....................................
Fairing and Cowling For Pylon, Transmission, Engine and Tail Pipe ...................................
Fiberglas Honeycomb Fairing Repair .................................................................................
Ammo Floor Scuff Doubler Installation ...............................................................................
Station Diagram ................................................................................................................
Change 71
iii
P-3
1-2B
1-2C
1-29
1-32B
1-34
1-36
1-39
1-40
1-40A
1-42
1-44
2-2
2-3
2-4
2-9
2-11
2-12
2-13
2-14
2-15
2-16
2-17
2-18
2-19
2-20
2-21
2-22
2-23
2-24
2-25
2-26
2-27
2-28
2-30
2-32
2-33
2-34
2-35
2-36
2-38
2-40
2-43
2-47
2-48
2-50
TM 55-1520-234-23
LIST OF ILLUSTRATIONS (Cont)
Figure
2-35
2-36
2-37
2-38
2-39
2-40
2-41
2-42
2-43
2-43A
2-44
2-45
2-46
2-47
2-48
2-48A
2-49
2-50
2-51
2-51A
2-51B
2-52
2-53
2-54
2-54A
2-54B
2-55
2-56
2-57
2-58
2-59
2-60
2-61
2-62
2-63
2-64
2-65
2-66
2-66A
2-66B
2-67
2-68
2-68A
2-69
2-70
2-71
2-72
2-73
2-74
Title
Diagonal Brace Tube Installation ..........................................................................................
Canopy Frames and Center Window ....................................................................................
Aft Bulkhead Rivet Removal (Typical Two Places) ...............................................................
Alignment - Upper Right and Left Canopy Frames ...............................................................
Upper Center Window ...........................................................................................................
Center Window Rivnut Hole Dimensions ..............................................................................
Canopy Frame and Cross Member Installation .....................................................................
Stress Loads - Canopy Windows ..........................................................................................
Insert Repair to Edge of Windows ........................................................................................
Deleted
Pilots Door .
.....................................................................................................................
Gunners Door .......................................................................................................................
Strut Installation - Pilots and Gunners Doors ........................................................................
Door Handles - Pilots and Gunners Doors ............................................................................
Door Strut Assembly P/N 209-030-640-1, -3 and -5 ..............................................................
Vent and Drain Locations .....................................................................................................
Pilots Seat Installation ..........................................................................................................
Pilots Seat Assembly . ..........................................................................................................
Gunners Seat Installation ......................................................................................................
Deleted
Deleted
Armor Panels ........................................................................................................................
Engine Mount Installation .....................................................................................................
Deleted
Wire Strike Protection System .............................................................................................
Damage Limits - Wire Strike Deflector .................................................................................
Airframe External Components . ...........................................................................................
Work Aid - Tailboom Support Stand ......................................................................................
Tailboom Installation . ...........................................................................................................
Tailboom and Elevator Skins ................................................................................................
Tailboom and Synchronized Elevator Structure ....................................................................
Bulkhead at Boom Station 59.50 ..........................................................................................
Bulkhead at Boom Station 80.44 ..........................................................................................
Bulkhead at Boom Station 101.38 .........................................................................................
Bulkhead at Boom Station 122.33 ........................................................................................
Bulkhead at Boom Station 143.28 ........................................................................................
Bulkhead at Boom Station 164.23 ........................................................................................
Bulkhead at Boom Station 185.18 (Part No. 209-961-189-7) ................................................
Longeron Damage Limits ......................................................................................................
Typical Tailboom Damage Limits ..........................................................................................
Damage Limits - Tail Rotor Drive Support Fittings ................................................................
Elevator Installation ..............................................................................................................
Vertical Fin Honeycomb Panels Damage Limits ...................................................................
Pylon Support Installation......................................................................................................
Hydraulic Fitting Supports P/N 209-030-267-11 and 209-030-267-29 Installation .................
Pylon Fitting Supporting Structure - Repair of Elongated Holes ............................................
Damage Limits - Flight Control Power Cylinder Support ........................................................
Damage Limits - Fifth Mount Support Fitting Assembly .........................................................
Wing Installation ...................................................................................................................
iv
Change 54
Page
2-57
2-59
2-60
2-61
2-63
2-64
2-65
2-70
2-72
2-76
2-78
2-79
2-82
2-85
2-88A
2-91
2-92
2-94
2-97
2-101
2-104B
2-104G
2-106
2-107
2-108
2-110F
2-113
2-114
2-116
2-119
2-122
2-124
2-125
2-126
2-126A
2-126D
2-127
2-132A
2-134A
2-136
2-137
2-140A
2-141
2-142A
2-144
TM 55-1520-234-23
LIST OF ILLUSTRATIONS (Cont)
Figure
2-75
2-76
2-77
2-78
2-79
3-1
3-2
3-3
3-4
3-5
3-6
3-6A
3-6B
3-7
4-1
4-2
4-3
4-3A
4-3B
4-3C
4-4
4-5
4-6
4-7
4-8
4-9
4-10
4-11
4-12
4-13
4-14
4-15
4-16
4-17
4-18
4-19
4-20
4-21
4-22
4-23
4-24
4-25
5-1
5-1A
5-2
5-3
5-4
5-5
5-6
5-6A
5-6B
5-7
Title
Page
Limits Chart-Bushings .. ......................................................................................................
Wings, Skins, Doors and Doublers .....................................................................................
Wing Structure ...................................................................................................................
Wing Skin Repair ...............................................................................................................
Deleted
Alighting Gear and Support Installation ..............................................................................
Checking Deflection of Cross Tubes ...................................................................................
Damage Limits-Cross Tubes .............................................................................................
Skid Tube Repairs...............................................................................................................
Ground Handling Wheels ..................................................................................................
Ground Handling Gear Pump Assembly ............................................................................
Ground Handling Gear Pump P/N HP-9902-41-10...............................................................
Packing Nut Tool.................................................................................................................
Ground Handling Wheels Adjustment Dimensions...............................................................
Power Plant Installation.......................................................................................................
Deleted
Engine Vibration Test Equipment Cabling Tiedown .............................................................
Engine Assembly-Removal/Installation ...............................................................................
Engine Mount-Fittings (Trunnions) Installation.....................................................................
Damage Limits-Aft Engine Mount Fittings (Trunnions) and Bearings ...................................
Particle Separator ...............................................................................................................
Particle Separator-Exploded View .......................................................................................
Airflow Diagram . ...............................................................................................................
Doubler for Inlet Vane Reinforcement .................................................................................
Removal/Installation Foreign Object Damage Screen .........................................................
Procedural Steps Installing Foreign Object Damage Screen (Bottom Half) ........................
Air Induction Area (Cowling Omitted) ..................................................................................
Exhaust System Components .............................................................................................
Engine Tailpipe Repair........................................................................................................
Engine Deck and Firewall Sealing .......................................................................................
Exhaust Infrared Suppression System.................................................................................
Repair Procedures-Infrared Suppression .............................................................................
Engine Oil Supply and Cooling System with Bleed Air Driven Fan .....................................
Oil Cooler Cleaning Equipment Setup-Typical.....................................................................
Oil Cooling Turbine Fan (Janitrol)........................................................................................
Deleted
Power Lever Control System...............................................................................................
Flight Idle Stop Installation ..................................................................................................
Power Turbine Governor RPM Controls . ............................................................................
Rigging Diagram-Governor RPM Controls...........................................................................
Electrical Cable Installation-Engine Left Side ......................................................................
Electrical Cable Installation-Engine ....................................................................................
Main Rotor Installation ........................................................................................................
Main Rotor Installation Torque Values ................................................................................
Main Rotor Tracking Chart ..................................................................................................
Vertical Vibration Correction Char .......................................................................................
Lateral Vibration Correction Chart ......................................................................................
Tracking Main Rotor............................................................................................................
Pitch Link Adjustment (Prior to Accomplishment of MWO 55-1520-244A50-9) ....................
Pitch Link Adjustment (After Accomplishment of MWO 55-1520-244-50-9) .........................
Vertical and Lateral Vibration Chart for K747 Rotor Blades .................................................
Tool Application Grip Lock Installation on Pitch Horn ..........................................................
Change 73
v
2-145
2-146
2-147
2-148
3-2
3-2D
3-3
3-6
3-8
3-11
3-14A
3-14B
3-15
4-2
4-4
4-4A
4-4F
4-4H
4-7
4-8
4-10
4-14
4-17
4-18
4-21
4-22
4-24
4-25
4-27
4-31
4-35
4-36
4-39
4-46
4-48A
4-50
4-52
4-55
4-57
5-2C
5-3
5-6
5-7
5-9
5-10
5-10
5-10A
5-10D
5-12
TM 55-1520-234-23
LIST OF ILLUSTRATIONS (Cont)
Figure
5-8
5-9
5-10
5-11
5-12
5-13
5-14
5-15
5-16
5-17
5-17A
5-17B
5-17C
5-17D
5-17E
5-17F
5-17G
5-17H
5-17J
5-17K
5-17L
5-17M
5-17N
5-17P
5-17Q
5-17R
5-17S
5-17T
5-17U
5-17V
5-17W
5-17X
5-17Y
5-17Z
5-17AA
5-17AB
5-17AC
5-17AD
5-17AE
5-18
5-19
5-20
5-21
5-22
5-23
5-24
5-25
Title
Page
Tool Application-Main Rotor Mast Nut Removal/Installation .................................................
5-13
Corrosion and Damage Limits-Split Cones ...........................................................................
5-14
Pitch Link Assembly ..............................................................................................................
5-15
Main Rotor Hub and Blade Assembly ...................................................................................
5-16B
Work Aid For Main Rotor Blade Bolt Removal-Fabrication Instructions ...............................
5-17
Work Aid Application-Removal of Main Rotor Blade Retaining Bolt ......................................
5-18
Main Rotor Blade ..................................................................................................................
5-19
Main Rotor Blade Authorized Patch Area ..............................................................................
5-22
Main Rotor Blade Trim Tab Installation .................................................................................
5-26
Tool Application-Alignment of Main Rotor Hub and Blades ...................................................
5-29
K747 Main Rotor Blade ....................................................................................................... 5-30C
External Appearance Changes to K747-003-205 Blades Result from Improved Weight
Retention Features ...............................................................................................................
5-30E
Internal Modifications Incorporated in K747-003-205, -209, and -303 Blades for
Improved Weight Retention ..................................................................................................
5-30F
Damage Limits-Root Fittings (K747 Blade) ........................................................................... 5-30Q
Damage Limit Drag Strut (K747 Blade) ................................................................................ 5-30R
Inspection of K747-003-205, -209, and -303 Blade for Loss of Blade Weight
Retention Integrity ................................................................................................................
5-30S
Proximity Limits for Patches-K747 Main Rotor Blades .......................................................... 5-30U
Balance Adjustment for Patches (K747 Blade) ..................................................................
5-30X
Root Fitting Assembly ..........................................................................................................
5-30Z
Application of Skin Patch (K747 Blade) 5-30AB
Curing Patch with Blade Repair Fixture (K747 Blade) .......................................................... 5-30AE
Installation of Plug Patch (K747 Blade) ............................................................................... 5-30AH
Typical Double Plug Patch Repair (K747 Blade) .................................................................. 5-30AM
Application of Trailing Edge Doubler Patch (K747 Blade) . ................................................... 5-30AR
Spline Repair (K747 Blade)
.............................................................................................. 5-30AV
Rebonding Delaminated Leading Edge Erosion Guard (K747 Blade) .................................... 5-30AW
Typical Repair of Fluorocarbon Erosion Guard Nicks and Cuts using Kit, P/N K747-207 ....... 5-30AZ
Application of Leading Edge Erosion Guard Patch (K747 Blade) ......................................... 5-30BA
Placement of Erosion Guard Replacement Part (Kit K747-206-1) and Method for
Determining Current Boot Material and Thickness (K747 Blade Series) ................................ 5-30BC
K747 Blade Uralite Repair..................................................................................................... 5-30BF
Application of Vacuum Bagging Materials and Placement of Erosion Guard.......................... 5-30BH
Repair Parts and Specimen Orientation (Kit, P/N K747-206) ................................................. 5-30BM
Vacuum Bagging for Installation of Erosion Guard Repair Kit, P/N K747-206 ........................ 5-30BP
Improvised Mold for Casting a Small Section of Leading Edge Filler..................................... 5-30BS
Repair Parts Orientation Use for Removal and Installation of Stainless Steel Erosion Guard
K747-003-303 and -303 Field Modified Blades ...................................................................... . 5-30BU
Preparation of K747-003-303 Blade for Application of Sealant ............................................ 5-30BW
Application of Sealant to Stainless Steel Guard for K747-003-303 Blade .............................. 5-30BY
Stainless Steel Erosion Guard Holding Fixture ..................................................................... 5-30BZ
Masking for Paint Touch-up After Installation of Kit K747-206 .............................................. 5-30CD
Main Rotor Hub Yoke Extension and Grip Assembly .............................................................
5-32
Main Rotor Hub Yoke and Trunnion for Main Rotor Hub........................................................
5-35
Main Rotor Hub Teflon Bearing Wear Pattern .......................................................................
5-36
Damage Limits-Main Rotor Yoke Extension ..........................................................................
5-37
Damage Limits-Main Rotor Hub Bolt Holes ...........................................................................
5-39
Damage Limits-Main Rotor Hub Yoke ...................................................................................
5-41
Damage Limit-Main Rotor Hub Grip ......................................................................................
5-42
Damage Limit-Main Rotor Hub Yoke Extension and Strap Fitting .........................................
5-43
vi Change 73
TM 55-1520-234-23
LIST OF ILLUSTRATIONS (Cont)
Figure
5-26
5-26A
5-27
5-28
5-29
5-30
5-31
5-32
5-33
5-34
5-35
5-36
5-37
5-38
5-39
5-40
5-41
5-42
5-43
5-44
5-45
5-46
5-47
5-47A
5-47B
5-48
5-49
5-50
5-51
5-52
5-52A
5-53
5-54
5-55
5-56
5-57
5-58
5-59
5-60
5-61
5-62
5-63
5-64
5-65
5-66
5-67
5-68
5-69
Title
Page
Damage Limits-Main Rotor Hub Pitch Horn and Inboard Bearing Housing
(Prior to Accomplishment of MWO 55-1520-244-50-6) and Pitch Horn Bushing ...................
Damage Limits-Main Rotor Hub Pitch Horn and Inboard Bearing Housing
(After Accomplishment of MWO 55-1520-244-50-6 and MWO 55-1520-244-50-9) .... .........
Damage limits-Main Rotor Drag Brace ........................... ........................................................
Damage Limits-Main Rotor Hub Trunnion ...............................................................................
Damage Limits-Main Rotor Hub Elastomeric Bearing ................... ..........................................
Main Rotor Hub Yoke Chafing Pad Installation Dimensions ...................................................
Tool Application-Bearing Removal From Housing ..................... .............................................
Tool Application-Bearing Removal From Grip ....................... ................................................
Tool Application-Bearing and Seal Installation in Grip .................. ..........................................
Main Rotor Hub Trunnion Centering Measurement ................... ............................................
Tool Application-Grip spacing Adjustment ..................... ...... ................................................
Main Rotor Hub-Grip Dust Seal To Radius Ring ...................... ...............................................
Rotor Balancing Kit P/N 7A050 ................................ ..............................................................
Tool Application-Main Rotor Hub Balancing ........................ ..................................................
Interpretation of Balancer Indication ................... .......... ........................................................
Mast Controls Installation
..................................................................................................
Tool Application-Drive Plate Spline Wear Measurement ................. .....................................
Collective Lever Thrust Washer Wear Limits ........................ .................................................
Scissors and Sleeve Assembly ...............................................................................................
Damage Limits-Hub, Sleeve, Scissors and Link ...................... ..............................................
Damage Limits-Spline Plate ................................. ..................................................................
Damage Limits-Collective Lever ............................... .............................................................
Damage Limits-Collective Lever Idler Link ......................... ....................................................
Installation of Shims ...............................................................................................................
Collective Lever Idler Link Assembly ......................................................................................
Swashplate and Support Assembly .............................. .........................................................
Damage Limits-Swashplate and Support Assembly ...............................................................
Damage Limits-Swashplate Anti-drive Link, Bellcrank and Support . .......................................
Deleted
Damage Limits-Pitch Link Assembly
(Prior to Accomplishment of MWO 55-1520-244-50-9) ......................................................
Damage Limits-Pitch Link Assembly
(After Accomplishment of MWO 55-1520-244-50-9) . .........................................................
Tracking Tail Rotor..................................................................................................................
Tail Rotor Installation ....... .............................. ......................................................................
Tool Application-Tail Rotor Hub and Blade Assembly Balancing ............. ...............................
Tail Rotor Hub and Blade Assembly ......................................................................................
Tail Rotor Blade Assembly
................................................................................................
Tail Rotor Blade Station Diagram and Scratch Type Damage Area Locations ....... .................
Tail Rotor Blade Area Authorized For Patch-Type Repair .......................................................
Tail Rotor Blade Butt Area Repair ...........................................................................................
Damage Limits-Tail Rotor Blade Pitch Horn ............................................................................
Tail Rotor Hub Assembly .......................................................................................................
Damage Limits-Tail Rotor Hub Yoke .......................... ...........................................................
Damage Limits-Tail Rotor Hub Trunnion Set ...................... ...................................................
Bearing Staking Tool P/N 101577
............................ .........................................................
Tool Application-Bearing Installation (Striking) in Tail Rotor Yoke ............ ..............................
Tail Rotor Controls Crosshead, Weights, Links & Control Tube .............. ................................
Damage Limits-Tail Rotor Control Crosshead ........................................................................
Damage Limits-Tail Rotor Control Counterweight Bellcrank .......... .........................................
Change 73
vii
5-44
5-44B
5-45
5-46
5-47
5-49
5-50
5-51
5-52
5-53
5-54
5-56
5-58
5-59
5-60
5-61
5-65
5-66
5-68
5-70
5-75
5-76
5-77
5-78
5-78
5-82
5-84
5-87
5-94B
5-94D
5-97
5-98
5-100
5-104
5-106
5-106
5-108A
5-109
5-110
5-115
5-117
5-118
5-119
5-121
5-123
5-126
5-127
TM 55-1520-234-23
LIST OF ILLUSTRATIONS (Cont)
Figure
5-70
5-71
5-72
5-73
5-74
5-75
5-76
6-1
6-2
6-3
6-4
6-5
6-6
6-6A
6-7
6-8
6-8A
6-8B
6-8C
6-8D
6-9
6-10
6-11
6-11A
6-12
6-13
6-14
6-15
6-16
6-17
6-17A
6-18
6-19
6-20
6-21
6-22
6-23
6-24
6-25
6-26
6-27
6-28
6-28A
6-29
6-30
6-31
6-32
6-33
6-33A
6-33B
6-34
6-35
6-36
Title
Damage Limits-Tail Rotor Controls Pitch Link .........................................................................
Damage Limits-Tail Rotor Control Counterweight Link ............................................................
Damage Limits-Tail Rotor Active Counterweight Support ........................................................
Damage Limits-Tail Rotor Control Tube . ................................................................................
Damage Limits-Link Assembly ................................................................................................
Damage Limit-Idler..................................................................................................................
Damage Limits-Lever..............................................................................................................
Power Train Diagram. .............................................................................................................
Metal Particles Contamination of Gearbox Oil.........................................................................
Main Driveshaft Installation . ...................................................................................................
Main Driveshaft Assembly.......................................................................................................
Inspection and Lubrication of Main Driveshaft .........................................................................
Coupling Wear Criteria for Driveshaft .....................................................................................
Limits Chart-Main Driveshaft Assembly...................................................................................
Positioning Pylon with T101440 Jacks ....................................................................................
checking Driveshaft Alignment ................................................................................................
Main Driveshaft SKCP 2281 ..................................................................................................
Main Driveshaft Installation and Removal Tool .......................................................................
Work Aid Tool Installation on Main Driveshaft .......................................................................
Main Driveshaft Damage Limits . ...........................................................................................
Transmission-Left Side ..........................................................................................................
Transmission-Right Side .........................................................................................................
Transmission Drive Quill Diagram...........................................................................................
Damage Limits-Transmission ..................................................................................................
Transmission Installation ........................................................................................................
Pylon Lift Link and Main Mounts .............................................................................................
Transducer Bracket Installation ..............................................................................................
Transmission Buildup .............................................................................................................
Transmission Shipping Caps, Covers and Plugs .....................................................................
Pylon Mounts and Dampers . ..................................................................................................
Pylon Mount Bolt Inspection Criteria .......................................................................................
Pylon Support Access Hole ....................................................................................................
Pylon Fifth Mount Assembly ...................................................................................................
Pylon Lift Link Attachment.......................................................................................................
Pylon Dampers ......................................................................................................................
Pylon Dampers Seal Installation Tool P/N 5120-EG-007 .........................................................
Shim Replacement-Damper Barrel Assembly..........................................................................
Shim Replacement-Damper Upper Shim.................................................................................
Shim Replacement-Damper Lower Shim.................................................................................
Main Rotor Mast......................................................................................................................
Deleted
Main Rotor Mast Damage Limits .............................................................................................
Damage Limits-Main Rotor Mast Assembly .............................................................................
Mast Sleeve Taper and Out of Round .....................................................................................
Main Rotor Mast Wear Limits . ................................................................................................
Upper Mast Bearing-Shim Adjustment.....................................................................................
Transmission Top Case-Dimension Check . ............................................................................
Input Drive Quill Assembly ......................................................................................................
Transmission Input Quill Workaid Application .........................................................................
Input Drive Quill Wear Sleeve Replacement ...........................................................................
Input Drive Quill Installation-Work Aid ...................................................................................
Hydraulic Pump and Tachometer Drive Assembly ..................................................................
Cockpit Air Blower Quill-Seal Replacement ............................................................................
viii Change 73
Page
5-128
5-129
5-130
5-131
5-133
5-134
5-135
6-1
6-7
6-9
6-10B
6-12
6-15
6-16A
6-19
6-20
6-20C
6-20D
6-20D
6-20E
6-22
6-23
6-24
6-24B
6-25
6-26
6-27
6-28
6-32
6-38
6-38A
6-39
6-42
6-44A
6-46
6-47
6-49
6-50
6-51
6-53
6-56
6-56A
6-57
6-58
6-59
6-61
6-64A
6-64B
6-64D
6-65
6-68
6-70A
TM 55-1520-234-23
LIST OF ILLUSTRATIONS (Cont)
Figure
6-37
6-38
6-39
6-40
6-40A
6-40B
6-41
6-42
6-43
6-44
6-45
6-46
6-47
6-47A
6-48
6-49
6-50
6-51
6-52
6-53
6-54
6-54A
6-54B
6-54C
6-55
6-56
6-57
6-57A
6-58
6-59
6-60
6-61
6-62
6-63
6-64
6-65
6-66
6-67
6-68
6-69
6-70
6-71
6-72
6-73
6-74
6-75
7-1
7-1A
7-2
7-3
7-4
Title
Work Aid for Quill Installation .................................................................................................
Tail Rotor Drive Quill ...............................................................................................................
Tail Rotor Driveshaft Installation ............................... ..............................................................
Tail Rotor Driveshaft Inspection Diagram .......................... .....................................................
Tail Rotor Driveshaft Hanger ...................................................................................................
Work Aid for Driveshaft Hanger Support Alignment ................................................................
Intermediate Gearbox Installation ...........................................................................................
Damage Limits-Intermediate Gearbox ....................................................................................
Intermediate Gearbox Quills, Seals, and Couplings (Typical) ..................................................
Coupling Teeth Wear Patterns ................................................................................................
Oil Filter Cap Assembly............................................................................................................
Deleted
Tail Rotor Gearbox Installation ................................................................................................
Damage Limits-Tail Rotor Drive Gearbox.................................................................................
Tail Rotor Gearbox Assembly ...............................................................................................
Roller Bearing Wear Pattern ..................................................................................................
Gear Patterns-Ninety Degree Gearbox ...................................................................................
Tail Rotor Gearbox Input Quill ................................................................................................
Tool Application-Retainer Bolt Removal & Installation..............................................................
Tool Application-Retainer Nut Removal & Installation .............................................................. .
Tail Rotor Gearbox Output Seal Replacement..........................................................................
Driveshaft/Spherical Couplings . ............................................................................................
Coupling Inspection (Sheet 1 of 2) ........................... ...............................................................
Coupling Inspection (Sheet 2 of 2) ...........................................................................................
Transmission Oil System Schematic............................ ..........................................................
Transmission Oil System Installation........................................................................................
Oil Filter, Transmission Sump ................................ ................................................................
Transmission Oil Pressure Relief Valve ........................... ......................................................
Oil Cooler Bypass Valve Assembly ..........................................................................................
Damage Limits-Inlet Fitting................................ . ...................................................................
Damage Limits-Return Bypass Fitting ........................... .........................................................
Damage Limits-Inlet Bypass Piston ............................. ............................................................
Damage Limits-Piston..................................... .........................................................................
Damage Limits-Nozzles ...........................................................................................................
Damage Limits-Sleeve.............................................................................................................
Damage Limits-Universal Fitting Bolt ........................... ......................................................
Damage Limits--Housing Assembly ........................................................................................
Damage Limits-Plunger ..........................................................................................................
Damage Limits-Elbow Fitting ................... ............. ...............................................................
Repair-Fittings .................... ...................................................................................................
Repair-Nozzles.........................................................................................................................
Repair-Plunger.........................................................................................................................
Work Aid Application-Plunger and Piston ......................... ......................................................
Bench Test-Oil Cooler Bypass Valve ......................................................................................
Transmission Oil Pump ................... ................. ...................................................................
Test Setup-Transmission Oil Pump ............................. ...........................................................
Hydraulic Schematic ...............................................................................................................
Hydraulic System No. 1...........................................................................................................
Hydraulic System No. 2.................................... .....................................................................
Dual Hydraulic Servo Cylinders Installation ......................... ...................................................
Hydraulic Reservoirs and Modules (Prior to Serial No. 69-16447) ..........................................
Change 73
viii A/(viii B blank)
Page
6-71
6-73
6-76
6-77
6-80B
6-80C
6-82
6-83
6-90
6-90B
6-90C
6-92
6-92B
6-94A
6-95
6-96
6-100
6-101
6-102
6-104
6-104B
6-104B
6-104C
6-106
6-107
6-108A
6-110A
6-114
6-117
6-118
6-119
6-119
6-120
6-120
6-121
6-122
6-124
6-125
6-126
6-127
6-128
6-128
6-129
6-132
6-133
7-2
7-4
7-4A
7-15
7-19
TM 55-1520-234-23
LIST OF ILLUSTRATIONS (Cont)
Figure
7-5
7-6
7-7
7-8
7-8A
7-9
7-10
7-11
7-11A
7-12
7-13
7-13A
7-14
7-15
7-16
7-17
7-18
7-19
7-19A
7-19B
7-19C
7-20
7-21
7-22
7-23
7-24
8-1
8-1A
8-2
8-2A
8-2B
8-3
8-4
8-5
8-6
9-1
9-2
9-3
9-4
9-5
9-5A
9-5B
9-5C
9-5D
9-5E
9-5F
9-5G
9-5H
9-6
9-6A
9-6B
Title
Page
Hydraulic Reservoirs and Modules (Serial No.70-15936 & Sub.) .............................................
Replacing Hydraulic Module Filters ..........................................................................................
Cleanable and Non-Cleanable Filter Elements ........................................................................
Deleted
Hydraulic Reservoir Assembly..................................................................................................
Emergency Hydraulic Accumulator Charging Diagram .............................................................
Hydraulic Accumulator Assembly (Parker-Hannifin) .................................................................
Hydraulic Accumulator Assembly (Sprague) ............................................................................
Lockout Valve and Accumulator ...............................................................................................
Dual Hydraulic Cylinder Assembly (Typical) ............................................................................
Damage Limits for Hydraulic Cylinder Bearing Housing ...........................................................
Servo-Cylinder Removal Work Aid...........................................................................................
Hydraulic Servo Cylinder..........................................................................................................
Flight Control Hydraulic Servo Cylinder ...................................................................................
Dual Hydraulic Cylinder Valve Connections..............................................................................
Connection of Hydraulic Cylinders to Swashplate.....................................................................
SCAS Servo Actuator Assembly .............................................................................................
Tail Rotor Control Cylinder and Support Assembly-Removal and Installation ..........................
Tail Rotor Control Cylinder and Support Assembly-Disassembled ...........................................
Hydraulic Servo Cylinder, Part No.1660 Series (Typical), Exploded View ................................
Hydraulic Servo Cylinder .........................................................................................................
Armament Pylon Actuator . ......................................................................................................
Work Aid for Armament Pylon Actuator Disassembly/Assembly ...............................................
Work Aid for Seal Installation on Armament Pylon Actuator . ...................................................
Schematic Diagram of Connections To Armament Pylon Actuator for Bench Test ...................
Armament Pylon Actuator Installed in Test Fixture . .................................................................
Instrument System Equipment Location ...................................................................................
Deleted
Pitot Static System ..................................................................................................................
Connection for Pitot Leak Check (Typical)................................................................................
Connection for Static Leak Check (Typical)..............................................................................
Test Circuit Setup for Fuel Tank Unit Capacitance and Resistance Checks..............................
Test Circuit Setup for Fuel Quantity Indicator Bench Test ........................................................
Circuit Arrangement and Adapter Cable for Fuel Quantity
Adjustment Procedures on Installed System.............................................................................
Volt-Ammeter Test Setup ........................................................................................................
Simplified Electrical Bus Wiring Schematic ..............................................................................
Equipment Locations................................................................................................................
Gunners Section - Equipment Location ....................................................................................
Pilots Section - Equipment Location .........................................................................................
Standby Inverter Adjustment ...................................................................................................
Pilots Night Vision Lighting Circuitry.........................................................................................
Gunners Night Vision Lighting Circuitry . ..................................................................................
Pilots Master Caution Panel Assembly - Front View . ...............................................................
Gunners Master Caution Panel Assembly - Front View.............................................................
Pilots Master Caution Panel Schematic....................................................................................
Pilots Master Caution Panel Assembly .....................................................................................
Pilots Master Caution Panel Subassembly ...............................................................................
Control Panel Printed Circuit Board 81-0815-1 .........................................................................
Caution Panel Assembly (BHT P/N 204-075-705-43)................................................................
Gunners Master Caution Panel Assembly ................................................................................
Gunners Master Caution Panel Schematic ...............................................................................
Change 73
ix
7-20
7-22
7-23
7-26B
7-28
7-29
7-30
7-34
7-35
7-36A
7-38A
7-39
7-42
7-44
7-47
7-48
7-50
7-50A
7-50F
7-50H
7-56
7-57
7-58
7-59
7-60
8-2
8-21
8-26C
8-26D
8-31
8-32
8-33
8-38
9-2
9-3
9-5
9-6
9-29
9-36A
9-36B
9-36D
9-37
9-42
9-42F
9-42J
9-42K
9-42P
9-42R
9-42T
TM 55-1520-234-23
LIST OF IILUSTRATIONS (Cont)
Figure
9-6C
9-6D
9-6E
9-7
9-8
9-9
9-10
9-11
9-12
9-13
9-14
9-15
9-16
10-1
10-1A
10-1B
10-2
10-3
10-4
11-1
11-1A
11-1B
11-1C
11-1D
11-1E
11-1F
11-1G
11-2
11-3
11-4
11-4A
11-5
11-6
11-7
11-8
11-9
11-10
11-11
11-11A
11-11B
11-12
11-13
11-14
11-14A
11-14B
11-14C
11-14D
11-15
12-1
12-2
12-3
Title
Gunners Master Caution Panel Assembly Interconnecting Wiring Diagram .....................
Removal of Finish for Electrical Bonding .........................................................................
Gunners Master Caution Panel Assembly Polarizing Keys Locator ..................................
RPM Limit Warning Detector Nameplate (Saturn) ........................... ................................
RPM Limit Warning Detector Nameplate (Bell) ................................ ...............................
Alignment of RPM Limit Detector (Saturn or Symbolic) ...................... .............................
Alignment of RPM Limit Detector (Bell) ....................................... ...................................
Bench Test Setup for Saturn and Symbolic Display Designed Detectors ............ .............
Bench Test Setup for BHT Designed Detectors ................................. ..............................
Armament Equipment Location ............................................ ...........................................
Wing Stores Armament Test Panel ........................................... ......................................
Wing Stores Armament Test Panel Wiring Diagram ........................................................
TOW Missile Launcher Positions ............................................. .......................................
Fuel System .............................................................. ....................................................
Fuel Boost Pump Cartridge ..............................................................................................
Work Air for Fuel Pump Cartridge Removal .............................. ......................................
Collapsing Forward Fuel Cell ......................................... .................................................
Installed Forward Fuel Cell ..............................................................................................
Collapsing AT Fuel Cell ...................................................................................................
Collective Controls ...................................................... ....................................................
Anti-torque System Wear and Damage Limits .................................................................
Elevator Control System Wear and Damage Limits ............................... ........................
Collective Control System Wear and Damage Limits ............................ ..........................
Cyclic Control System Wear and Damage Limits ................................ ............................
Flight Control System Bearings .............................................. .......................................
Power Cylinder Support - Lateral Cyclic and Collective Controls ................. ....................
Power Cylinder Support - Fore and Aft Cyclic Controls ....................................................
Pilot Collective Pitch Stick .................................................. ..........................................
Pilot Collective Stick Mounting Details ....................................... .....................................
Gunner Collective Pitch Stick ..........................................................................................
Gunner Collective Pitch Stick -Wear ................................................................................
Fore and Aft Cyclic Controls.............................................................................................
Lateral Cyclic Controls ..................................................... ...............................................
Work Aid for Rigging Swashplate ............................................ ........................................
Swashplate Rigging Dimensions ......................................................................................
Pilot Cyclic Stick ......................................................... ....................................................
Pilot Cyclic Stick Friction Adjustment ....................................... .......................................
Gunner Cyclic Stick .........................................................................................................
Fore and Aft Cyclic Jackshaft...........................................................................................
Fore and Aft Cyclic Jackshaft Wear and Damage Limits ........................... ......................
Tail Rotor Controls ........................................................ ..................................................
Tail Rotor Pedal Installation .............................................................................................
Elevator Controls .......................................................... ..................................................
Stability and Control Augmentation System (SCAS).........................................................
Fore and Aft Cyclic Controls.............................................................................................
Lateral Cyclic Controls ..................................................... ...............................................
Tail Rotor Controls ........................................................ ..................................................
SCAS Transducer Calibration ..........................................................................................
Fire Detector System ....................................................... ...............................................
Rain Removal Nozzle Assembly .............................................. .......................................
Rain Removal Nozzle and Cleared Area Pattern ................................. ............................
x
Change 73
Page
9-42Y
9-44
9-44
9-45
9-46
9-47
9-48
9-52
9-55
9-74
9-82
9-83
9-98
10-2A
10-6B
10-6C
10-10
10-11
10-13
11-4
11-6D
11-6L
11-6N
11-6U
11-6Z
11-7
11-8
11-8B
11-10
11-12
11-12A
11-14
11-15
11-16B
11-17
11-19
11-22A
11-24
11-26A
11-26B
11-27
11-31
11-36
11-38B
11-38C
11-38E
11-38G
11-40
12-2
12-5
12-6
TM 55-1520-234-23
LIST OF IILUSTRATIONS (Cont)
Figure
13-1
13-2
13-3
13-4
13-4A
13-5
13-6
13-7
13-8
13-9
13-10
13-11
13-12
13-12A
13-13
13-14
16-1
16-2
16-3
16-4
16-5
16-6
16-7
16-8
16-9
16-10
16-11
17-1
17-2
17-3
17-4
17-5
17-6
17-7
C-1
F-1
F-2
F-3
F-4
F-5
F-6
F-6A
F-7
F-8
F-9
F-10
F-11
F-12
F-13
Title
Page
Environmental Control System.........................................................................................
Environmental Control Unit .............................................................................................
Environmental Control System Schematic .......................................................................
Blower Impeller Assembly P/N 209-072-436-1 .................................................................
Blower Impeller Installation - Typical for Impeller
P/N 209-070-585-1 and P/N 197803-1 .............................................................................
Air Distribution Valve Assembly .....................................................................................
Limits Chart - Air Distribution Valve .................................................................................
Tool Application - Teflon Lip Seal Installation on Poppet . ................................................
Vent Air Control Valve--Installation...................................................................................
Vent Air Control Valve Assembly ..................................................................................... .
Vent Air Control Valve Setup............................................................................................
Temperature Control Valve .............................................................................................
Temperature Control Valve Schematic.............................................................................
ECU Cooling Turbine Lubrication ....................................................................................
Pressure Regulator and Shutoff Valve .............................................................................
Pressure Regulator and Shutoff Valve Schematic ............................................................
Ml 18 Wing Mounted Smoke Dispenser............................................................................
Cradle Assemblies in Stow Position for Installation on Outboard Rack .............................
Cradle Assemblies in Position for Installation on Rocket Launcher ..................................
Wing Stores Pylon . .........................................................................................................
Special Tools for Ejector Rack ........................................................................................
Outboard Ejector Rack Installation ..................................................................................
Ejector Rack Alignment ...................................................................................................
Inboard Ejector Rack Alignment ......................................................................................
Outboard Ejector Rack - Exploded View...........................................................................
Electrical Connections for Ejector Rack ...........................................................................
Inboard Rack with Ejector Disassembled .........................................................................
Canopy Removal System Interconnections .....................................................................
Installation of Window Cutting Assembly in Gunner Window . ..........................................
Installation of Window Cutting Assembly in Pilots Window ..............................................
Installation of Window Cutting Assembly in Gunners Door ..............................................
Installation of Window Cutting Assembly in Pilots Door ...................................................
Installation of Junction Manifolds and Arming/Firing Units in Canopy
Removal System .............................................................................................................
Typical Shell Installation ..................................................................................................
Inventory Items List..........................................................................................................
Wiring Identification Code ................................................................................................
Symbols Chart ................................................................................................................
AC Load Analysis Chart . .................................................................................................
DC Load Analysis Chart . .................................................................................................
AC Power System Wiring Diagram...................................................................................
Anti collision and Position Lights Systems .......................................................................
NVG Position Lights System ............................................................................................
Armament Systems Wiring Diagram ...............................................................................
Attitude Indicating System Wiring Diagram .................................................................... ..
Battery System Wiring Diagram ......................................................................................
Caution Lights System Wiring Diagram ...........................................................................
Caution Panel Internal Schematic Diagram .....................................................................
Engine Device System Wiring Diagram ...........................................................................
Environmental Control System Wiring Diagram ..............................................................
Change 73
xi
13-2
13-4
13-5
13-9
13-10B
13-12
13-13
13-15
13-19
13-21
13-22
13-27
13-28
13-28A
13-30
13-31
16-2
16-5
16-6
16-9
16-10
16-11
16-13
16-14
16-17
16-24
16-25
17-2
17-4
17-6
17-8
17-10
17-12A
17-15
C-3
F-70
F-71
F-75
F-81
F-85
F-87
F-88A
F-89
F-115
F-117
F-119
F-121
F-123
F-125
TM 55-1520-234-23
LIST OF IILUSTRATIONS (Cont)
Figure
F-14
F-15
F-16
F-17
F-18
F-19
F-20
F-21
F-22
F-23
F-24
F-25
F-26
F-27
F-28
F-29
F-30
F-31
F-32
F-33
F-34
Title
External Power System Wiring Diagram .................................................................................
Force Trim System Wiring Diagram .......................................................................................
Fuel Valve and Engine Oil Valve System Wiring Diagram.......................................................
Fuel Boost System Wiring Diagram ........................................................................................
Fuel Quantity Indicating System Wiring Diagram ....................................................................
Generator and DC Bus System Wiring Diagram......................................................................
Governor Control System Wiring Diagram ...........................................................................
Hydraulic Control System Wiring Diagram . ............................................................................
Idle Stop System Wiring Diagram ...........................................................................................
Igniter System Wiring Diagram . .............................................................................................
Interior Lights System Wiring Diagram . ..................................................................................
Pitot Heater System Wiring Diagram ......................................................................................
Pressure Indicating Systems Wiring Diagram..........................................................................
RPM Limit Warning System Wiring Diagram ..........................................................................
RPM Test Set Schematic Diagram .........................................................................................
Searchlight and Skid Landing Light Systems ..........................................................................
Starter System Wiring Diagram ..............................................................................................
Tachometer Indicator Systems Wiring Diagram ...................................................................
Temperature Indicating Systems Wiring Diagram ...................................................................
TOW Blower Cooling System Wiring Diagram .......................................................................
Turn and Slip Indicating System Wiring Diagram ...................................................................
xii Change 73
Page
F-126
F-127
F-128
F-129
F-130
F-131
F-133
F-134
F-135
F-137
F-139
F-141
F-143
F-145
F-147
F-149
F-150
F-151
F-153
F-173
F-175
TM 55-1520-234-23
LIST OF TABLES
Number
1-1
1-2
1-3
1-4
1-5
1-6
2-1
2-1A
2-2
4-1
4-2
4-3
5-1
5-1A
5-1B
5-1C
5-1D
5-2
6-1
7-1
7-2
7-3
7-4
7-5
7-6
8-1
8-2
8-3
8-4
8-5
8-6
8-6A
8-6B
8-6C
8-7
8-8
8-9
8-10
8-11
8-12
8-13
8-14
8-15
8-16
8-17
8-18
9-1
9-2
9-3
9-4
9-5
9-6
Title
Page
Engine Fuel Specifications ......................................................................................................
Approved Fuels ......................................................................................................................
Consumable Maintenance Supplies and Materials ..................................................................
Special Tools and Equipment .................................................................................................
Adhesive Mix Ratio, Pot Life, and Cure Cycles .......................................................................
Mooring Hardware Chart ........................................................................................................
Tailboom Skin and Structure Classification and Damage ........................................................
Treatment of Corroded Surfaces ............................................................................................
Structural Repair Materials......................................................................................................
Repair Methods.......................................................................................................................
Troubleshooting-Engine Oil System ........................................................................................
Dimension Tolerance-Turbine Fan ..........................................................................................
Troubleshooting-Main Rotor System .......................................................................................
Difference Between Models K747-003 Blade...........................................................................
Classification of Damage-K747 Main Rotor Blade . .................................................................
Classification of Damage-K747 Main Rotor Blades..................................................................
Plug Patch Data ......................................................................................................................
Troubleshooting-Tail Rotor System . .......................................................................................
Troubleshooting-Drive Train System .......................................................................................
Troubleshooting-Hydraulic System . ........................................................................................
Maximum Allowable Leakage for Hydraulic System ................................................................
Torque Values for Fluid Connections.......................................................................................
Inspection Criteria ..................................................................................................................
Troubleshooting Chart.............................................................................................................
Leading Particulars T/R Control Cylinder ................................................................................
Troubleshooting-Dual and Gas Producer Tachometers............................................................
Troubleshooting-Engine Oil Pressure Indicating System..........................................................
Troubleshooting-Engine Oil Temperature Indicator..................................................................
Deleted
Troubleshooting-Fuel Pressure Indicator .................................................................................
Troubleshooting-Torque Pressure Indicators ...........................................................................
Tolerance (f knots) ..................................................................................................................
Vertical Velocity Tolerance Feet Scale Accuracy.....................................................................
Altimeter Scale Error ..............................................................................................................
Troubleshooting-Airspeed Indicators ......................................................................................
Troubleshooting-Altimeters .....................................................................................................
Troubleshooting-Attitude Indicators .........................................................................................
Troubleshooting-Turn & Slip Indicator ....................................................................................
Troubleshooting-Vertical Velocity Indicator..............................................................................
Troubleshooting-Standby Compass ........................................................................................
Troubleshooting-Clock ............................................................................................................
Troubleshooting-Fuel Quantity Indicator..................................................................................
Fuel Tank Table of Limits........................................................................................................
Troubleshooting-Free Air Temperature Gage ..........................................................................
Troubleshooting-Voltmeter/Ammeter.......................................................................................
Troubleshooting-Pilots Steering Indicator (PSI) . .....................................................................
Troubleshooting-Battery System .............................................................................................
Troubleshooting-External Power System.................................................................................
Troubleshooting-Generator and DC Bus System .....................................................................
Troubleshooting-Inverter System.............................................................................................
Troubleshooting-Starting System ...........................................................................................
Troubleshooting-Ignition System .............................................................................................
Change 71
xiii
1-3
1-4
1-11
1-19
1-24
1-38A
2-109
2-173
2-175
4-28
4-32
4-42
5-1
5-30D
5-30J
5-30N
5-30AG
5-95
6-2
7-7
7-12
7-16A
7-50E
7-50H
7-50K
8-8A
8-12
8-13
8-17
8-19
8-20A
8-20B
8-20B
8-23
8-24
8-25
8-26
8-26A
8-28
8-29
8-30
8-33
8-35
8-37
8-39
9-15
9-18
9-21
9-26
9-32
9-34
TM 55-1520-234-23
LIST OF TABLES (Cont)
Number
9-7
9-8
9-8A
9-8B
9-9
9-10
9-11
9-12
9-13
9-14
9-15
9-16
9-17
9-18
9-19
9-20
9-21
9-22
9-23
9-24
9-25
9-26
9-27
9-28
10-1
11-1
12-1
13-1
13-2
16-1
17-1
F-1
F-2
F-3
Title
Troubleshooting-Cockpit Lights .............................................................................................
Troubleshooting-Instrument Panel & Console Lights ...............................................................
Troubleshooting-Pilots Master Caution Panel Assembly .........................................................
Troubleshooting-Gunners Master Caution Panel Assembly ....................................................
Troubleshooting-Rpm Limit Warning System . ........................................................................
Troubleshooting-Position Lights System .................................................................................
Troubleshooting-Anticollision Light System ............................................................................
Troubleshooting-Searchlight System .......................................................................................
Troubleshooting-Transmission Oil Level Light .......................................................................
Troubleshooting-Engine De-ice Circuitry ...............................................................................
Troubleshooting-Engine Oil Bypass Circuitry ..........................................................................
Troubleshooting-Fuel Valve Circuitry .....................................................................................
Troubleshooting-Fuel Boost Pump Circuitry ...........................................................................
Troubleshooting-Governor Control System Circuitry................................................................
Troubleshooting-Idle Stop Solenoid Circuitry...........................................................................
Troubleshooting-Force Trim System Circuitry ........................................................................
Troubleshooting-Hydraulic Control System Circuitry ...............................................................
Troubleshooting-Environmental Control System Circuitry........................................................
Troubleshooting-Pitot Heating System Circuitry .....................................................................
Troubleshooting-TOW Blower Cooling System........................................................................
Troubleshooting-Rocket Launcher Circuitry ............................................................................
Troubleshooting-XM-18 Minigun Circuitry ..............................................................................
Troubleshooting-Wing Stores Jettison Circuitry .......................................................................
Troubleshooting-Smoke Grenade Dispenser System Circuitry ...............................................
Troubleshooting-Fuel System .................................................................................................
Troubleshooting-Flight Controls ..............................................................................................
Troubleshooting-Fire Detector System ....................................................................................
Troubleshooting-Environmental Control System......................................................................
Troubleshooting-Air Distribution Valve ....................................................................................
Troubleshooting-Ejector Racks................................................................................................
Shell Assemblies ....................................................................................................................
Equipment List (Electrical) ......................................................................................................
Equipment List (Armament) ...................................................................................................
XM 65 Interconnection Wire List .............................................................................................
xiv Change 71
Page
9-34A
9-36
9-42C
9-42U
9-50
9-57
9-58A
9-59
9-61
9-62
9-63
9-64
9-65
9-66
9-67
9-68
9-70
9-71
9-72A
9-72B
9-87
9-90A
9-92A
9-95
10-2
11-1
12-1
13-6
13-17
16-7
17-14
F-3
F-14
F-17
TM 55-1520-234-23
PREFACE
boxes are mounted under covers along the top of the tail
P-1. General
boom and front of the vertical fin. A controllable elevator
is also mounted on the tail boom.
a. This manual is the official document for Aviation
Unit and Intermediate Maintenance of Army Model AH-IS
b.
Propulsion System. The propulsion system
Helicopters.
consists of a gas turbine engine, main drive shaft,
transmission and mast, main rotor, and the tail rotor with
b. The purpose of this manual is to familiarize you
its drive shafts and gear boxes. The transmission and
with the maintenance functions to be performed at the
engine are mounted on the forward fuselage aft of the
Aviation Unit and Intermediate maintenance levels. The
crew compartment, and covered by cowling and fairing.
Table of Contents for this manual is provided to assist in
The engine drives the transmission through the short
determining the chapter in the manual in which individual
main drive shaft, rotating the mast and main rotor. Power
functions are covered. This manual provides all essential
is also taken off from the transmission to drive the tail
information for personnel to accomplish Aviation Unit and
rotor, which compensates for main rotor torque to control
Intermediate maintenance on the complete airframe, its
the helicopter heading.
Fuel tanks consist of two
components, and systems, excluding armament and
interconnected cells, located in the forward fuselage.
avionics subsystem as indicated for Aviation Unit and
Intermediate maintenance activities in the Maintenance
c. Flight Controls.
Flight controls are direct
Allocation chart (MAC). (Refer to Appendix B).
mechanical linkages from sticks and pedals at pilot's and
gunner's stations, assisted by hydraulic cylinders powered
P-2. Quality Assurance/Quality Control(QA/QC).
by transmission-driven hydraulic pumps. A stabilization
and control augmentation system is also incorporated in
Information not available.
the control linkage to steady the helicopter during use of
armament.
P-3. Description.
d. Armament Provisions.
Armament provisions
include mounting, wiring and hydraulic lines for an
armament turret under the forward end of the fuselage, an
ammunition compartment immediately aft of the turret
location, sights and control panels at crew stations, and
mounting pylons for external armament pods on each
wing.
The AH-IS helicopter is a two-place assault type
helicopter with a narrow fuselage, single main and tail
rotors, short wings, and provisions for a variety of
armament. The forward fuselage is built up on two main
longitudinal beams with lateral bulkheads, floors, shear
panels, and decks of honeycomb panel construction,
forming a box beam. Crew compartments are arranged in
tandem, with the pilot seated behind and above the
gunner, and are covered by a transparent canopy with two
doors. Both doors are sections of the canopy, hinged at
top, with the gunner's door opening at left and the pilot's
door opening at right side. Both seats are protected by
armor on backs, seats, and sides of supports. The
compartment area is ventilated by forced air from a
transmission- driven blower. The short wings provide
support for armament mounting pylons and also aid
maneuverability by providing lift at higher air speeds.
e. Landing Gear. Landing gear is skid type, with
arched cross tubes attached to the fuselage. Exposed
portions of cross tubes are covered by streamlined
fairings to reduce drag. A tail skid is provided to protect
the aft end of the tail boom during a tail-low landing.
P-4. Reporting of Errors.
Every effort is made to keep this publication current
and error free. Review conferences with using personnel,
and a constant review of accident and flight test reports
assure inclusion of the latest data in this publication.
However, we cannot correct an error unless we know of
its existence. In this regard it is essential that you do your
part. Reports of errors, omissions, and recommendations
a. Tail Boom Section. The tail boom section,
attached to the forward fuselage by four bolts, is a
tapered semimonocoque structure with a vertical fin
slanting up and aft at the rear end to support the tail rotor.
Tail rotor drive shafts and gear
P-1
TM 55-5120-234-23
for improving this publication by you are encouraged.
Your letter or DA Form 2028, Recommended Changes to
Publications and Blank Forms should be mailed to
Commander, U.S. Army Aviation Systems Command,
ATTN: AMSAV-MPSD, 4300 Goodfellow Blvd., St. Louis,
MO 63120-1798.
P-5. Destruction of Army Material To Prevent Enemy
Use.
For destructions of Army materiel to prevent enemy
use, refer to TM750-244-1-5.
P-8. Special Tools and Equipment.
Aviation Unit and Intermediate maintenance special
tools and equipment will be found in TM55-1520-234-23P
(RPSTL) manual. Use of special tools and equipment for
complex tasks is described in this manual.
P-9. Calibration.
Equipment requiring calibration shall be indicated and
reference made to a publication(s) containing the
applicable procedures.
P-6. Maintenance of Forms and Records.
Maintenance of forms, records, and reports which are
to be used by maintenance personnel at all maintenance
levels are listed in and prescribed I by DA Pam 738-751.
a. Aircraft
components,
accessories,
and
instruments requiring calibration shall be specified in
Chapter 1.
b. Special tools and test equipment shall be
calibrated as specified in TB 750-25, Army Metrology and
Calibration System.
P-7. Authority For Substitution.
Substitution or interchange of items of materiel for
maintenance of Department of the Army aircraft shall not
be authorized, nor shall orders be issued for shipment.
Substitution
or interchangeability shall only be
authorized by US
Army Aviation Systems Command.
P-2
P-10. Storage.
Refer to TM740-90-1 and Appendix E for Storage of
Aircraft.
Change 38
TM 55-1520-234-23
Figure P-1. AH-1S Helicopter
P-3/(P-4 blank)
TM 55-1520-234-23
CHAPTER 1
AIRCRAFT GENERAL
Section I. SERVICING
(8) Activate flow control handle to ON or FLOW
position. Fuel flow will automatically shutoff when fuel
cell is full. Just prior to normal shutoff, fuel flow may
Servicing information and procedures are presented by
cycle several times, as maximum fuel level is reached.
systems or components in the following paragraphs.
(9) Assure flow control handle is in OFF or NO
Points used in frequent servicing and replenishment of
FLOW position and remove nozzle.
fuels, oils, hydraulic fluid and other materials are shown
(10) Replace fuel nozzle cap.
on a diagram. (See figure 1-1.)
(11) Replace fuel filler cap.
1-2. Fuel System Servicing.
(12) Disconnect fuel nozzle ground.
(13) Disconnect ground from helicopter to servicing
The fuel supply tank consists of two cells, located in the
unit.
fuselage forward and aft of the wings, interconnected by a
(14) Disconnect servicing unit ground from ground
crossover and vented through a common outlet line.
stake.
Each cell has a sump and a fuel pump with drains (1 and
(15) Return fire extinguisher to designated location.
2, figure 1-1) accessible through doors in the fuselage
lower skin. The fuel tank filler cap (12) and static ground
b. Gravity or Open-Port Refueling (Power Off).
receptacle (11) serve both tanks. TOTAL TANK
(1) Refer to figure 1-1 for fuel filler location.
CAPACITY FOR FUEL SYSTEM IS 262 U.S. GALLONS
(2) Assure that fire guard is in position with fire
AND NORMAL SERVICE CAPACITY IS 260 U.S.
extinguisher.
(3) Ground servicing unit to ground stake.
GALLONS.
(4) Ground servicing unit to helicopter.
(5) Ground fuel nozzle to ground receptacle
WARNING
located adjacent to fuel receptacle on helicopter.
(6) Remove fuel filler cap.
Servicing personnel shall comply with all safety
precautions and procedures specified in FM 10(7) Using latch tool, attached to filler cap cable
68 Aircraft Refueling field manual.
open refueling module if equipped with closed circuit
receptacle. Refer to figure 1-1A.
(8) Remove nozzle cap and insert nozzle into fuel
NOTE
receptacle.
The fuel system can be serviced by either
(9) Activate flow control handle to ON or FLOW
closed circuit or gravity system.
position.
(10) Assure flow control handle is in OFF or NO
a. Closed Circuit Refueling (Power Off).
FLOW position and remove nozzle.
(1) Refer to Figure 1-1 for fuel filler location.
(11) Replace fuel nozzle cap.
(2) Assure fire guard is in position with fire
(12) Close refueling module by pulling cable until
extinguisher.
latch is in locked position, if equipped with closed circuit
(3) Ground servicing unit to ground stake.
receptacle. Refer to figure 1-1A.
(4) Ground servicing unit to helicopter.
(13) Replace fuel filler cap.
(5) Ground fuel nozzle to ground receptacle,
(14) Disconnect fuel nozzle ground.
ocated adjacent to fuel receptacle on helicopter.
(15) Disconnect ground from helicopter to
(6) Remove fuel filler cap, and assure refueling
servicing unit.
module is in locked position. Refer to figure 1-lA.
(16) Disconnect servicing unit ground from ground
(7) Remove nozzle cap, insert nozzle into fuel
stake.
acceptable, and lock into position.
1-1. Servicing - General.
Change 15
1-1
TM 55-1520-234-23
(17) Return fire extinguisher to designated location.
b-A. RAPID (HOT) Refueling (Closed Circuit).
(1) Before RAPID Refueling.
(a) IDLE.
(b) FORCE TRIM Switch - FORCE TRIM.
(17) Return fire extinguisher to designated location.
b-B. RAPID (HOT) GRA VITY Refueling.
(1) Before RAPID Refueling.
(a) Throttle - IDLE.
(b) FORCE TRIM Switch - FORCE TRIM.
WARNING
WARNING
In case of helicopter fire, observe fire
emergency procedures in Chapter 9 of TM551520-234-10.
(2) During RAPID Refueling. A crewmember shall
observe the refueling operation (performed by authorized
refueling personnel) and stand fireguard as required. One
crewmember shall remain in the helicopter to monitor
controls. Only emergency radio transmission should be
made dur- ing rapid refueling.
(3) Refer to figure 1-1 for fuel filler location.
(4) Assure fireguard is in position with fire extinguisher.
(5) Ground servicing unit to ground stake.
(6) Ground servicing unit to helicopter.
(7) Ground fuel nozzle to ground receptacle located
adjacent to fuel receptacle on helicopter.
(8) Remove fuel filler cap, and assure refueling module
is in closed position. Refer to figure 1-1A.
(9) Remove nozzle cap and insert nozzle into fuel
receptacle and lock into position.
(10) Activate flow control handle to ON or FLOW
position. Fuel flow will automatically shutoff when fuel
cell is full. Just prior to normal shutoff, fuel flow may
cycle several times, as maximum fuel level is reached.
(11) Assure flow control handle is in OFF or NO FLOW
position and remove nozzle.
(12) Replace fuel nozzle cap.
('13)
Replace fuel filler cap.
(14) Disconnect fuel nozzle ground.
(15) Disconnect ground from helicopter to servicing
unit. AFTER RAPID FUELING. The pilot shall be
advised, by the refueling crew, that fuel cap is secure and
grounding cables have been removed.
(16) Disconnect servicing unit ground from ground
stake.
In case of helicopter fire, observe fire
emergency procedures in Chapter 9 of TM551520-234-10.
(2) During RAPID Refueling. A crewmember, shall
observe refueling operation (performed by authorized
refueling personnel) and stand fireguard as required.
One crewmember shall remain in the helicopter to
monitor controls. Only emergency radio transmission
should be made during rapid refueling.
(3) Refer to figure 1-1 for fuel filler location.
(4) Assure fireguard is in position with fire extinguisher.
(5) Ground servicing unit to ground stake.
(6) Ground servicing unit to helicopter.
(7) Ground fuel nozzle to ground receptacle located
adjacent to fuel receptacle on helicopter.
(8) Remove fuel filler cap.
(9) Using latch tool attached to filler cap cable, open
refueling module if equipped with closed circuit rapid
refueling receptacle. Refer to figure 1-1A.
WARNING
During RAPID GRAVITY Refueling, exercise
extreme caution to prevent fuel splashing from
fuel cell or fuel nozzle. Any fuel leakage could
be extremely hazardous if ingested into engine
air intake.
(10)
Remove nozzle cap and insert nozzle into fuel
receptacle.
(11)
Activate flow control handle to ON o' FLOW
position. Fuel flow will automatically shutoff when cell is
full.
(12)
Assure flow control handle is in OFF or NO
FLOW position and remove nozzle. Close refueling
module by pulling cable until latch is in locked position, if
equipped with closed circuit receptacle. Refer to figure 11A.
1-2 Change 15
TM 55-1520-234-23
c. Defueling.
(13) Replace fuel nozzle cap.
(1) Remove drain valve from aft fuel cell.
(14) Replace fuel filler cap.
(15) Disconnect fuel nozzle ground.
NOTE
(16) Disconnect ground from helicopter to servicing
unit. AFTER RAPID FUELING. The pilot shall be
advised, by the refueling crew, that fuel cap is secure and
grounding cables have been removed.
Aft fuel drain valve assembly is a two piece valve
which will automatically close valve opening when
lower valve is removed. Refer to FM 10-68 for
defueling procedures.
(17) Disconnect servicing unit ground from ground
stake.
(2) Install MS24392D12 fitting, with flexible hose
installed, in valve assembly in bottom of cell. Valve will
open as fitting is being installed.
(18) Return fire extinguisher to designated location.
(3) After defueling, remove MS24392D12 fitting. Install
lower section of valve and lockwire.
Change 71
1-2A
TM 55-1520-234-23
Figure 1-1. Servicing points diagram
1-2B
Change 15
TM 55-1520-234-23
Figure 1-1A. Closed Circuit Refueling System
Change 15
1-2C/(1-2D blank)
TM 55-1520-234-23
worldwide use, and will be the only fuels readily available
in the Army support system.
d. Fuel Requirements. Fuel requirements for the
engine are listed in Table 1-1. A general listing of
acceptable fuels is provided in Table 1-2. The fuels listed
in Table 1-2 for each type have nearly identical
characteristics. All of the fuels are compatible and may
be mixed in aircraft fuel tanks. The use of fuels shall be
in accordance with TB 55-9150-200-25.
(2) Alternate Fuels: These are fuels which can be
used continuously when Army Standard fuel is not
available, without reduction of power output. Power
setting adjustments may be required when an alternate
fuel is used.
WARNING
(3) Emergency Fuels: These are fuels which can
be used if Army Standard and approved Alternate fuels
are not available. Their use is subject to a specific time
limit. (Refer to TM 55- 1520-234-10.)
Turbine engine fuels, as well as gasoline, form
explosive mixtures readily. To ensure safety of
personnel, aircraft handling and filling
operations shall conform to TM 10-1101.
f. Use of Fuels. There is no special limitation on
the use of Army Standard fuel, but certain limitations are
imposed when Emergency fuels are used. For the
purpose of record, fuel mixtures shall be identified as to
the major component of the mixture (except when the
mixture contains leaded gasoline) and recorded on DA
Form 2408-13 (Aircraft Inspection and Maintenance
Record). A fuel mixture which contains over 10 percent
leaded gasoline shall be recorded as all leaded gasoline.
e. Fuel Types.
Fuels are classified as Army
Standard, Alternate, or Emergency.
(1) Army Standard Fuels: These are the Army
designated primary fuels adopted for
Table 1-1. Engine Fuel Specifications
ARMY STANDARD FUEL
ALTERNATE FUEL
EMERGENCY FUEL
MIL-T-5624
MIL-T-5624
MIL-G-5572
Grade JP-4
Grade JP-5
Aviation Gasoline
(NATO Code No. F-40)
(NATO Code No. F44)
NOTE: An entry shall be made on DA Forms 2408-13 if EMERGENCY fuel is used.
(1) The use of kerosene fuels (JP-5 type) in turbine
engines dictates the need for observance of special
precautions. Both ground starts and air restarts at low
temperature may be more difficult due to low vapor
pressure. Kerosene fuels having a freezing of -40°F (40°C) limit the maximum altitude of a mission to 28,000
feet under standard day conditions. Those having a
freezing point of -55°F (-48°C) limit the maximum altitude
of a mission to 33,000 feet under standard day condition.
Refer to TB 55-9150-200-25
unleaded gasoline must be discontinued pending result of
internal inspection.
NOTE
Two parts of unleaded gasoline mixed with one
part of kerosene fuel (JP-5 type) produce a fuel
which is preferred above straight unleaded
gasoline. In the fueling record, this mixture
should be identified as unleaded gasoline.
NOTE
(2) The use of straight unleaded gasoline may
shorten the operating life of combustor parts; therefore, its
use between scheduled internal (hot end) inspections is
limited. When the time limit has been reached, the use of
Unleaded gasoline leaves combustor parts
clean; therefore, no special cleaning is
required between scheduled internal (hot end)
inspections.
1-3
TM 55-1520-234-23
Table 1-2. Approved Fuels
Source
Primary or
Standard Fuel
Alternate Fuel
US Military Fuel
JP-4 (MIL-T-5624)
Commercial Fuel
(ASTM-D-1655)
JET B
JET A
America Oil Co.
Atlantic Richfield
Richfield Div
B. P. Trading
Caltex Petroleum Corp
Cities Service Co.
Continental Oil Co
. Gulf Oil
EXXON Co, USA
American JP-4
Aerojet B
American Type A
Aerojet A
Richfield A
Mobil Oil
Phillips Petroleum
Shell Oil
Sinclair
Standard Oil Co
Chevron
Texaco
Union Oil
Foreign Fuel
B.P.A.T.G
Caltex Jet B
Conoco JP-4
Gulf Jet B
EXXON Turbo Fuel
B
Mobil Jet B
Philjet JP-4
EXXON A-1
Aeroshell JP-4
Mobil Jet A-1
Chevron B
Texaco Avjet B
Union JP-4
JP-5 (MIL-T-5624)
JET A-1
NATO F-34
Aerojet A-1
Richfield A-1
B.P.A.T.K.
Caltex Jet A-1
CITGO A
Conoco Jet-50
Gulf Jet A
EXXON A
Mobil Jet A
Philjet A-50
Conoco Jet-60
Gulf Jet A-1
EXXON-A-1
MOBIL JET A
Aeroshell 640
Aeroshell 650
Superjet A-1
Jet A-1 Kerosine
Chevron A-1
Avjet A-1
Jet A Kerosine
Chevron A-50
Avjet A
76 Turbine Fuel
Superjet A
NATO F-44
(High flash type)
Belgium
Canada
Denmark
NATO F-40
(Wide cut type)
BA-PF-2B
3-6P-24c
France
Germany (West)
Greece
3GP-22F
JP-4 MIL-T-5624
Air 3407A
UTL-9130-007/UTL 9130-010
Italy
Netherlands
VTL-9130-006
JP-4 MIL-T-5624
AMC-143
Norway
AA-M-C-1421
D. Eng RD 2493
Portugal
Turkey
United Kingdom
(Britain)
JP-4 MIL-T-5624
JP-4 MIL-T-5624
JP-4 MIL-T-5624
JP-4 MIL-T-5624
D. Eng RD 2498
NOTE:
Anti-icing and Biocidal Additive for Commercial Turbine Engine Fuel - The fuel system icing inhibitor shall conform to
MIL-I-27686. The additive provides anti-icing protection and also functions as a biocide to kill microbial growths in
aircraft fuel systems. Icing inhibitor conforming to MIL-I-27686 shall be added to commercial fuel, not containing an icing
inhibitor, during refueling operations, regardless of ambient temperatures. Refueling operations shall be accomplished in
accordance with accepted commercial procedures. Commercial product "PRIST" conforms to MIL-I-27686.
1-4
TM 55-1520-234-23
(3) Leaded gasoline, either straight or mixed
with other fuels in any proportion, will deposit a layer of
lead oxide on combustor parts. The lead oxide attacks
the underlying metal and also acts as an insulator which
reduces combustion efficiency and causes the formation
and deposition of carbon. Therefore, the operating time
between scheduled internal (hot end) inspections is
limited. If the permissible accumulated operating time is
exceeded, a special cleaning and inspection is
mandatory. See TM 55-2840-229-23.
NOTE
Transmission and gear boxes shall be drained and
refilled in accordance with paragraph 1-6.
(1) Oil
(C94)
used
in
engine,
main
transmission, and gearboxes, oil systems is authorized
and directed for ambient temperatures above minus
32°C (minus 25°F).
(2) Oil
(C93)
used
in
engine,
main
transmission, and gearboxes, oil systems is specified for
operation in ambient temperatures below minus 32°C
(minus 25°F). This oil may also be used when oil (C94)
is not available.
Special cleaning and inspection may be
delayed for 10 operating hours provided
that only Army Standard fuel is used
during the delay.
1-3.
CAUTION
Under no circumstances shall oil (C94)
be used at temperatures below minus
32°C (minus 25°F).
Engine Oil System Servicing.
The engine oil tank (8, figure 1-1) is located above
the engine in the aft fairing. Oil level sight gage and
filler cap are on front of tank, accessible through doors
on pylon center fairing. Tank drain valve is accessible
through the engine compartment, and has an overboard
drain line.
d. Procedure for Changing Engine Oils.
(1) When changing over from oil (C93) to oil
(C94) in engine oil system, accomplish steps below.
(a) Drain oil (C93) from system.
a. Fill engine oil tank to spill-over for normal
servicing. Sight gage is positioned to show low oil level.
When oil level is below spillover level, the tank should
be filled. Useful capacity of tank is 2.25 US gallons,
with expansion space of 1.15 gallons.
(b) Inspect, clean and reinstall all engine oil
filters and strainers.
b. Before adding oil, determine whether system
contains oil (C93) or oil (C94).
Maximum oil
consumption for T53-L-703 engine is 0.3 gal./hr. (2.4
US pints). The oil level sight gage is provided for the
purpose of determining a low oil condition. When oil
level is at sight gage level, oil supply is 2.75 ±.25 quarts
low. When servicing oil tank, fill completely to a spill
over condition. The system warning light will come ON
and the bypass valve will open when the oil is down 3.8
quarts low from spill over.
(d) Operate engine for 30 minutes to 1
hour. Shut down engine.
c. Usage of Oils. It is not advisable to mix oil
(C94) and oil (C93) except when an emergency exists
and conditions warrant. If mixing becomes necessary,
the engine oil system shall be drained within 6 hours of
operations, and refilled with the appropriate oil. (See
subparagraphs (1) and (2) below for oil usage.) If engine
oil system is to be replenished with oil (C94) proceed in
accordance with paragraph 1-3, step d. When refilling
engine oil system with oil (C93) proceed in accordance
with paragraph 1-3, step d (2), steps (a) thru (e).
Change 29
(c) Fill engine oil tank to lip of filler neck
with oil (C94). Motor engine to pump oil into cooler and
lines. Check tank level and refill. Repeat until level
does not change, indicating the cooler and lines are
refilled.
(e) Inspect, clean, and reinstall all engine
oil filters and strainers.
1. If
oil
filter
was
heavily
contaminated, accomplish all steps below.
2. If oil filter was not heavily
contaminated, omit steps (f) and (g) and accomplish
steps (h) through (i) below.
(f) Drain all oil from engine oil system, and
discard oil.
(g) Fill engine oil system with new oil (C94)
and release helicopter for use.
1-5
TM 55-1520-234-23
a. Before servicing transmission sump to upper
sight gage level, determine whether system contains oil
(C93) or oil (C94). If unable to determine type of oil
used, refer to paragraph 1-3, step c. Systems capacity
is 2.25 US gallons.
(h) After 5 hours operation, inspect and
clean all engine oil filters and strainers.
(i) 15 hours after oil change, inspect and
clean all engine oil filters and strainers.
(j) Revert
to
normal
schedule
inspections of engine oil filter and strainers.
of
1-5.
The 42° gear box (6, figure 1-1) is under a
removable cover, located between drive shaft covers of
tail boom and vertical fin. An oil level sight glass, a chip
detector and drain plug, are on the right side of the gear
box, a vented filler cap is on the top side of the gear
box. The 90°gear box (7) is covered by two removable
fairings near upper end of the vertical fin. An oil level
sight glass is provided on lower left side, a chip detector
and drain plug at bottom, and a vented filler cap on the
aft side of the gear box.
(2) When changing over from oil (C94) to oil
(C93) in engine oil system, accomplish the following.
(a) Drain oil (C94) from system.
(b) Inspect, clean, and reinstall all engine
oil strainers and filters.
(c) Fill engine oil tank with oil (C93) motor
engine to pump oil into cooler and lines. Check tank
level and add oil. Repeat until tank level does not
change, indicating that cooler and lines are filled.
(d) Operate engine until oil
operating temperature. Shut down engine.
a. Before servicing both gear boxes to indicated
sight glass level with turbine engine lubricating oil,
determine whether system contains oil (C93) or oil
(C94). If unable to determine type of oil used, refer to
paragraph 1-3, step c.
reaches
(e) Inspect, clean, and reinstall all engine
oil strainers and filter. Release helicopter for service
use.
1-6.
Procedures For Changing Transmission And
Gear Box Oils.
When changing over from oil (C93) to oil (C94) or from
oil (C94) to oil (C93) accomplish the following steps:
(f) After 5 hours of operation, inspect and
clean all engine oil strainers and filter.
a. Drain oil.
(g) After 15 hours of operation, since last oil
change, inspect and clean engine oil strainers and filter.
b. Change filter elements (transmission only).
c.
(h) Revert to normal interval of inspection
for engine oil strainers and filter.
1-4.
42°and 90°Gearbox Servicing.
Perform normal periodic inspection.
d Refill with appropriate oil (see paragraph 1-3,
step c.)
Transmission Oil System Servicing.
1-7.
The transmission oil supply is contained in the sump
case (9, figure 1-1.) A double sight gage on the sump
can be viewed through a small transparent plastic
window in the right-hand pylon cowling door, using a
light controlled by a push-button below the door. Before
servicing oil, determine whether system contained oil
(C93) or oil (C94). If unable to determine type of oil
used, refer to paragraph 1-3, step c. When filling is
required, open the cowling door for access to the filler
cap on the transmission support case. The sump drain
valve is accessible by removing an access panel on
fuselage below either wing.
Hydraulic System Servicing.
Reservoirs (3 and 10, figure 1-1) for hydraulic
systems No. 1 and No. 2 are located in a compartment
just aft of the canopy. Access doors are provided at
both sides of the fuselage. Fluid level sight glasses on
both reservoirs are visible through left side door. Each
reservoir has a filler cap accessible from nearest door.
The emergency collective accumulator (13) is
accessible by removing a panel located on the fuselage
below the right wing and has a pressure gage and filler
valve for nitrogen charging. A valve and stowed hose
connection, with instructions on a decal, are
1-6
TM 55-1520-234-23
provided in the same area for releasing hydraulic
pressure from the accumulator.
a. Bleed hydraulic pressure from emergency
collective accumulator.
Check accumulator gas
pressure gage, and charge as required with compressed
nitrogen (C89). (Refer to Chapter 7.)
When handling hydraulic fluid (MIL-H83282),
observe
the
following:
Prolonged contact with liquid or mist
can irritate eyes and skin. After any
prolonged
contact
with
skin,
immediately wash contacted area with
soap and water. If liquid contacts eyes,
flush them immediately with clear water.
If liquid is swallowed, do not induce
vomiting; get immediate medical
attention. Wear rubber gloves when
handling liquid. If prolonged contact
with mist is likely, wear an appropriate
respirator. When fluid is decomposed
by heating, toxic gases are released.
Change 21
1-6.1/(1-6.2 blank)
TM 55-1520-234-23
(5) Remove gaskets.
CAUTION
(6) Remove any foreign material from inside of
separator.
Hydraulic pressure must be released
from emergency collective accumulator
before adding fluid to avoid overfilling
hydraulic system.
(7) Remove foreign material from filter on front
face of lower portion of separator.
CAUTION
(8) Position gaskets over positioning pins on
lower separator.
Do not mix hydraulic fluid (C73) with fire
retardant fluid (C73A). Refer to TB 551500-334-25.
(9) Position upper separator on lower assembly.
Tilt top slightly forward to position assembly on four
positioning pins.
b. Service both reservoirs with hydraulic fluid (C73
or C73A) through filler caps. Each reservoir holds 3.2
US pints. Total capacity of System No. 1 is 6.0 pints:
capacity of System No. 2 is 6.6pints.
(10)
Secure upper assembly to flange with
latch at top of separator.
(11)
Engage latch assemblies on front face
of separator and lock.
NOTE
Secure front latch assemblies before
securing rear latches.
To avoid contamination, a sealed can of
fluid must be opened and used. Do not
use previously opened cans of hydraulic
fluid.
1-8.
(12)
Engage
separator and lock.
a. Cleaning.
(1) Open and secure transmission cowling.
on
rear
face
of
CAUTION
Particle Separator (Self Purging) Servicing.
The particle separator (4, figure 1-1) is an inertialtype consisting of an upper and lower assembly half.
Lower assembly half mounts air cleaner which collects
particles removed from engine air and ejects them
overboard. A foreign object damage screen consists of
two halves which fit around particle separator.
latches
Ensure that safety catch on latches is
engaged by exerting a slight pull on
release catch. Catch should not open.
(13)
Check for proper seating by appearance
of seals. Approximately 0.125 inch of rubber on gasket
assemblies between halves will be uniformly exposed.
Seal at aft edge of separator will be approximately half
compressed.
(2) Unlock the two latches on foreign object
screen.
(14)
Position top half of screen to engage aft
screen molding slot over separator split flange. Position
screen cut-out over latch on separator.
Refer to
Chapter 4, Section III.
(3) Disengage hook portions and lift top half of
screen free of particle separator.
(15)
Secure top half of screen to bottom half
by engaging and locking both latches.
(4) Release two latches on rear faces and two
latches on front face of separator halves by
simultaneously pressing safety latch up and lifting up on
release latch. Release latch on top of separator upper
half and remove upper half.
(16)
Minimize gaps which may exist between
top and bottom screen halves or between screens and
separator by repositioning or slightly reforming screens.
Gaps greater than 0.15 inch are cause for screen
replacement.
Change 2
1-7
TM 55-1520-234-23
(17)
1-9.
Close and secure transmission cowling.
(2) Fill pump to filler hole level with hydraulic
fluid (C73). Reinstall screw in filler hole.
Battery Servicing.
1-11.
The nickel-cadmium battery (14, figure 1-1) is
mounted in BATTERY compartment. It is connected to
the helicopter electrical system through a relay which is
controlled by the battery switch on the pilot's console.
Two overflow, or vent, tubes extend from the battery to
the underside of the fuselage. Access to the battery is
gained through a door in the helicopter battery
compartment. The battery is a lightweight, 24 volt 22
ampere-hour nickel-cadmium battery unit.
A 34
ampere-hour nickel-cadmium battery may be used as an
alternate.
CAUTION
CAUTION
To preclude damage to honeycomb
panels, solvents and water are to be
applied at the minimum pressure
required to maintain a constant flow
suitable for washing and rinsing. Steam
is not to be utilized.
a. General. The helicopter must be grounded prior
to any cleaning, maintenance, disassembly or
preservation.
Battery failure and explosions may be
caused by an excess of electrolyte in
the cells. The specific gravity of a
nickel-cadmium
battery
remains
constant when the battery is in either a
charged or discharged condition,
consequently the state of charge cannot
be determined by a test of the
electrolyte. Neither can the state of
charge be determined by a voltage test,
due to the fact that the voltage remains
constant over 90 percent of the
discharge time. Since the state of
charge cannot be determined by a
check of either voltage or the
electrolyte, the charging input to a
completely discharged battery must be
monitored in both current and time until
the ampere hour capacity of the battery
has been reached.
1-10.
NOTE
Additional cleaning procedures are
covered in this manual under individual
components.
Use trichlorotrifluoroethane cleaning
compound in a well ventilated area and
avoid prolonged breathing of vapors.
Do not use in an area with open flame
or high temperature as the products of
decomposition are toxic and very
irritating. Avoid contact with the skin.
Wear rubber gloves.
Ground Handling Wheels (Truck) Servicing.
a. Lubricate assemblies with grease (C67) through
fittings on wheels, actuating arms, and cradles as
frequently as operating conditions warrant.
b. Repair tires and tubes in accordance with TM
55-2620-200-24. Inflate with compressed air to 75 psi.
c.
Cleaning Helicopter.
Check and fill hydraulic pump as required.
(1) Hold handling gear assembly so that pump
is vertical with filler hole at upper end. Remove screw
from filler hole.
Change 54
b. Interior. Clean the interior of the helicopter to
prevent debris from falling into the operating
mechanism. If the seats and cushions need cleaning,
use mild soap (C125) and water. To remove grease or
oil spots use solvent (C143). Wipe dry with a clean
cloth. Finally, thoroughly clean the helicopter with a
vacuum cleaner.
c. Exterior.
Clean the exterior structure by
applying a mixture of one part cleaning compound (C41)
and three to seven parts water. Use the stronger
mixtures for exhaust outlet areas and other very dirty
surfaces. Wash a small area at a time making sure to
rinse thoroughly with water under pressure. If allowed to
dry or if not completely rinsed off, streaking will occur.
1-8
TM 55-1520-234-23
NOTE
1-11A. Removal of Snow and Ice.
Do not spray or rinse main rotor hub
with water under pressure.
CAUTION
Extreme care shall be exercised at all
times to prevent any damage to the
aircraft surfaces. Sharp instruments
such as picks, knives, or screw drivers
will not be used to loosen the ice
formation.
d. Canopy.
CAUTION
Do not use compounds other than those
specified. Avoid excessive scrubbing of
plastic panels during washing operation.
(1) Clean all transparent plastic with large
quantities of cleaning compound (C41) and water.
Cleaning compound (C39) may be used as an alternate
to clean transparent plastic.
a. Check entire helicopter for snow, frost, and ice
accumulation. Snow can be removed from airframe and
rotor blades by using a bristle brush or equivalent.
Ensure that helicopter skids are not frozen to ground.
(2) Gently free all caked mud or dirt with the
pads on the fingers. Do not use sponges or coarse
cloths. Rinse the area continuously while removing the
mud.
CAUTION
Extreme care must be exercised when
Extreme care must be exercised when
melting ice and frost with applied heat.
Water accumulation may flow into
critical areas in proximity to heat
application.
If heat gun is used,
exercise caution to prevent excessive
heat from damaging rotor blades,
bonded panels, metal surfaces, and
paint.
Do not use aliphatic naphtha (Type 1).
(3) Remove grease
naphtha, Type 2 (C88).
or
oil
with
aliphatic
(4) Allow surfaces to drip dry.
(5) Minor scratches may be
removed. Refer to TB 55-1560-276-24/1.
reduced
or
(6) Apply repellant and conditioner (C125).
Refer to TM 55-1500-204-25/1.
e. Rotor Blades.
b. Frost or moderate ice. Apply heat to ice
accumulations and dry with rags as melting occurs.
c. Severe ice accumulation. Helicopter should be
moved into a warm hangar, when possible, for natural
de-icing.
NOTE
Do not spray or rinse main rotor hub
with water under pressure.
Wash rotor blades with one part cleaning compound
(C41) and nine parts water.
Extreme caution must be exercised in
the use of Ethylene Glycol-water
solutions (including Ethylene Glycol,
technical or specification MIL-A-8243
Anti-Icing/Deicing/Defrosting Fluid) in
and around aircraft having silver or
silver-coated
electrical/electronic
circuitry. Rapid oxidation and fire can
occur when Glycol-water solutions
come in contact with and short across
bare or defectively insulated silver or
silver-coated electrical circuits such as
wiring, switches, circuit breakers, etc.,
which are carrying positive direct
current (DC).
Cleaning solvent is flammable and
toxic.
Provide adequate ventilation.
Avoid prolonged breathing of solvent
vapors and contact with skin or eyes.
f. Treatment of Aluminum and Magnesium Alloy
Corrosion. Aluminum and magnesium alloy corrosion
will be treated in accordance with TM 43-0105. Apply
the protective paint finish to the affected area
immediately after drying of chemical treatment in
accordance with TB 746-93-2.
Change 38
De-icing fluids are toxic irritants;
protective precautionary measures
apply.
1-9
TM 55-1520-234-23
NOTE
CAUTION
It is neither advisable nor necessary to
remove epoxy primer when refinishing.
Because of adverse effects of heated
de-icing fluid, precaution must be taken
to protect bearings, plastic windows,
covers and boots.
d. Apply de-icing/defrosting fluid (C58A) to remove
ice or heaving accumulations and to retard recurrence.
Fluid may be applied with a low pressure spray or a
brush, and will provide retarding protection for
approximately 10 hours.
(3) Wash all surfaces with cleaning compound
(C39) after removal of paint and rinse to ensure removal
from crevices, pockets, etc.
e. Upon completion of de-icing procedures, check
all controls for ice and freedom of movement.
Alodine is extremely dangerous.
It
contains an oxidizing ingredient which
may cause an explosion if it comes in
contact with combustible materials such
as paints, solvents, etc. Chromic acid is
extremely dangerous. Avoid breathing
fumes or contact with clothing or body.
Contact with combustible materials may
cause fire.
1-11B. De-Ice K747 Blade.
a. Use MIL-A-8243C or MIL-A-8243B ethylene
propylene glycol.
b. Do not break ice with sharp, blunt, or similar
instrument.
c.
Apply to ice by spray:
(1) Pressure low/medium.
(2) Hand pump spray (atomizer).
d. Wipe off de-ice fluid from all surfaces.
should not be dripping off of blade.
Fluid
e. Allow ice to melt off.
f. Very light scraping can be done using either a
wood, plastic, or teflon scraper.
(5) Allow treated surface to dry, and apply
primer as follows: Prime magnesium, aluminum,
fiberglass and cadmium-plated steel with primer (C102)
in accordance with TB746-93-2.
(6) Apply finish coats of acrylic lacquer, gloss or
camouflage of color to match adjoining area.
g. Do not scrape leading edge boot.
1-12.
(4) Apply chemical film treatment to bare metal
while the surface is still wet from washing rinse. Treat
aluminum with Alodine (C37). Treat magnesium with
chromate pickle solution (C38). Treat cadmium plated
steel with a ten percent solution of chromic acid (C3).
Before the solution dries, wipe with a clean dampened
cloth or rinse thoroughly with clean, fresh water to
remove all chemical film treatment solution.
Subassembly Painting (AVIM)
(a) Apply one light first coat and allow 30
minutes to one hour to air dry.
Helicopter and components will be painted in
accordance with procedures covered in TB 746-93-2.
Special painting procedures will be covered in this
manual under individual components. General painting
procedures will be in accordance with the following
paragraphs.
(b) Apply second full covering coat, and
allow one hour minimum drying time.
(c) Apply the final lacquer coat of the
proper color number, thinned approximately one part
basic lacquer to one and one-half parts thinner (C140).
a. Paint Refinish - Exterior.
(7) Apply final finish to various special areas as
(1) Mask around area to be repainted. Apply
masking tape (C133) at a skin edge, or break point, that
will allow the new paint to blend with the existing paint.
(2) Remove paint from area to be refinished
with thinner, (C140), or approved paint remover, using
caution to prevent solutions from contacting acrylic
plastics.
Change 38
follows:
(a) Paint the external exhaust blast area,
consisting of the engine exhaust pipe fairing, drive shaft
forward access cover and the two top forward skin
sections of the tailboom with two coats of high heat
resistant acrylic enamel (C60). Apply the first coat as a
thin tacky coat; follow with one full wet coat while the
first coat is still tacky.
1-10
TM 55-1520-234-23
(b) Finish the ADF sense antenna surface
with anti-static epoxy (C61).
(c) Finish surfaces within 12 inches of the
battery, and other surfaces subject to electrolyte spillage
or spray, with one coat of primer (C100). Follow with
two coats of lacquer (C75). Allow 30 minutes to one
hour air drying time between coats.
Toluene is highly flammable. Ground
container to dispense.
Avoid skin
contact. Avoid breathing vapors.
(8) Allow finish coats a minimum of six hours
air drying time at normal temperature (700 to 900F), and
humidity (30 to 75 percent) conditions prior to masking
for application of markings.
(9) Apply walkway coating (C148) to wing and
skid gear.
a1. Paint Refinish - Exterior - Low Reflective
Acrylic Lacquer.
NOTE
Apply lacquer (C75A) within four hours
after step (6).
(6) After masking, activate surface to be
painted by washing with toluene (C144). Allow to dry
thoroughly.
NOTE
(1) Remove any markings in area to be painted
by sanding with 320 grit sandpaper (C112).
Ensure that toluene did not loosen
masking tape. Re-mask as necessary.
NOTE
(7) Areas which were sanded must be primed
with epoxy polyamide primer (C100) prior to coating with
lacquer. Areas which did not require sanding do not
require priming.
Polyurethane coated surfaces must be
scuff sanded all over to remove glaze.
Use 320 grit or finer sandpaper (C112).
NOTE
(2) Where finish has been damaged or does not
adhere to base material, remove finish down to base
material. Use 320 grit or finer sandpaper (C112).
(3) Use 320 grit or finer sandpaper (C112)to
remove all rough, high, or corroded areas. Feather
rough areas. Fair edges where finish was removed in
step (2) or where finish did not adhere to primer.
Apply lacquer (C75A) within one to
eight hours after application of primer
(C100).
(8) Prepare Low Reflective Acrylic Lacquer for
Application.
(a) The lacquer base (C75A) in its original
container shall have all settling and/or caking
redispersed by hand stirring using a clean metal or wood
stirrer. After hand stirring, invert and agitate, on a paint
shaker, in original container, for an additional 30
minutes prior to thinning.
Use solvent (C124) in a well ventilated
area. Avoid prolonged breathing of
vapors and do not use in an area with
open flame or high temperature.
(4) Remove all grease, oil, grime, etc., by
washing area to be painted with solvent (C124). Dry
thoroughly.
(5) After solvent is completely dry, use masking
tape (C134) to mask area to be painted.
Change 38
(b) Immediately after shaking, pour lacquer
base (C75A) into a clean container and add thinner
(C140) one to one and one-half parts by volume to one
part lacquer base.
(c) Thinned mixture shall be agitated in the
pressure pot for thirty minutes prior to use. Agitation
shall then be continuous throughout application.
1-10A
TM 55-1520-234-23
(d) Acrylic retarder (C62A) may be added
up to thirty percent by volume when high temperatures
or humidity cause blushing.
(d) Instrument panel shrouds, stand by
compass card holder bracket.
(4) Use dark gull gray camouflage lacquer
(C81), on all cockpit surfaces except the cockpit floor
and items listed in step (3).
CAUTION
Halts over two minutes duration during
application will require flushing of paint
lines and spray gun to remove dried
particles prior to restarting application.
c.
Paint Refinishing - Cockpit Floor.
(1) Remove seats and clean cabin floor area.
Refer to Chapter 2 for removal and installation.
NOTE
(2) Mask all areas and items adjacent to floor.
A slight increase in fluid and air
pressure, approximately 5 psi over that
used for application of standard acrylic
lacquers, is required. Respraying of
partially dried areas will cause
excessive surface roughness.
(3) Remove all grease, oil, dust, etc., from floor
with aliphatic naphtha (C88), and clean cloths. Sand the
cleaned area with No. 400 or 600 grit paper (C112) and
wipe with clean cloth.
(4) Apply primer (C102) as necessary.
(9) Apply finish coats of low reflective acrylic
lacquer. When spraying, hold gun 10 to 12 inches from
the surface to prevent dry spraying. This material shall
be sprayed wet.
(5) Apply two coats of lacquer (C81), to floor
area.
NOTE
NOTE
Allow 2 hours drying time for paint.
The low reflective acrylic lacquer will
dry to touch in one hour. Full cure is
attained in 12 hours. Allow 30 minutes
minimum to 12 hours maximum dry
time between coats.
(6) Apply markings in accordance with TB74693-2.
(7) Apply walkway coating (C148).
(8) Install pilots and gunners seats.
Chapter 2.
b. Paint Refinishing - Cockpit Interior.
(1) Mask all acrylic surfaces and items of
different color.
(2) Use primer (C102) where old primer has
been removed.
Refer to
d. Dissimilar Metals. Faying surfaces of dissimilar
metals must be treated before assembling, to prevent
contact of bare metals and resulting corrosion damage.
e. Procedures for Dissimilar Metals.
(3) Refinish the following with lusterless black
lacquer (C80).
(a) Cyclic and collective control sticks.
(b) Edge lit plastic panels and metal
backing panels.
(c) Instrument glare shield and lighting
fixtures.
Change 38
(1) Mating surfaces of dissimilar metals shall
receive a minimum of two coats of zinc chromate primer
(C102) before assembly.
(2) Where magnesium is one of dissimilar
metals, install tape (C131) between mating surfaces.
Tape shall extend not less than 0.26 inch beyond edges
of joint to prevent moisture bridging between dissimilar
metals.
1-10B
TM 55-1520-234-23
When an item number is referenced in the manual, you
may locate the item through its C designator and item
number. C designators are used only with consumable
maintenance supplies and materials.
Consumable
maintenance supplies and materials tables are found
only in this chapter therefore, the table number will not
be referenced in the text. NSN's in the consumable
table list only the smallest quantity of the material. If
larger quantities are required, see TM 55-1520-234-23P.
1-13. List of Consumable Maintenance Supplies
and Materials.
Consumable maintenance supplies and materials are
listed in table 1-3 in alphabetical order.
Each
consumable also has an item number assigned for ease
of location and reference. When an item number is
unknown you may locate any consumables used within
this manual through its alphabetical arrangement.
TabIe1-3. Consumable Maintenance Supplies and Materials
The supplies and material listed in this table are required for maintenance support of this equipment and am authorized
to be requisitioned by CTA 50-970 (Common Table of Allowances).
Item
Ref. No.
No.
Description
and FSCM
NSN
01
Abrasive Pads, Nylon Wed, Scotch-Brite,
81348
Type A, Very Fine, L-P-50, Type 1, Class 1,
Size 1
7920-00-0659-9175
1
Acetone, Technical
0-A-51
6810-00-184-4796
2
Acid, Boric, Technical Grade
-C-265
6810-00-264-6535
3
Acid, Chromic
0-C-303, Type II
6810-00-264-6517
4
Acid, Hydrochloric (Muriatic)
O-H-765
6810-00-222-9641
5
Acid Nitric
O-N-350
6810-00-237-2918
6
Use C12A
7
Adhesive, A-6 with Activator A
A6 (98911)
8040-00-691-1322
7A
Adhesive, EA9330 (HYSOL)
(04347)
8040-01-089-9073
7B
Adhesive, EA9309
299-947-125, Type I (33564)
8040-01-012-8749
8
Adhesive, EC-678
MIL-A-9117
8040-00-262-9060
9
Adhesive, EC-776
MMM-A- 122
8040-00-664-0439
10
Adhesive, EC1300L
EC1300L(04633)
8040-00-514-1880
11
Adhesive, EC2216, Scotchweld
Scotchweld 2216BA
(94960)
8040-00-145-0019
11A
Adhesive (RP-1257)
(48909)
12
Adhesive, Epoxy, EA9340, Metalset A4
MMM-A-1754
8040-00-944-7292
12A
Adhesive, PS-18
MIL-A-8576
(77902)
8040-00-526-1910
13
Adhesive, Q2-0046
RTV 730 (71984)
8040-00-251-2312
13A
Adhesive Mixture
Epon 826
Versamid 125
Diethylenetramine (DTA)
(86961)
(11884)
(79741)
8030-00-144-9658
8030-00-893-4224
8040-01-194-3933
Change 73
1-11
TM 55-1520-234-23
Table 1-3. Consumable Maintenance Supplies and Materials (Cont)
Item
No.
Description
Ref. No.
and FSCM
NSN
14
Adhesive, Rubber Base
MMM-A-1617
8040-00-262-9031
14A
Adhesive,
Ethyl-2-Cyanoacrylate
MIL-A-46050 Type 2,
Class 3
8040-01-140-0954
15
Adhesive, Silicone Rubber
MIL-A-46146
8040-00-224-4655
16
Adhesive, Silicone Rubber
MILA-46146
8040-01-130-3210
16A
Adhesive, Silicone
MIL-1-46146A
8040-01-009-1562
17
Adhesive, Two-Part
EA934NA
MMM-A-132, Type 1
Class 3, (33564)
8040-01-102-2098
18
Adhesive, Two-Part
Liquid Silicone
Adhesive, Type --Class 2
Adhesive, Uralane
(Part A and B)
5738-A/BX
SR 529 and SRC18 (01139)
8040-00-097-6524
MMM-A-132
EC3549AB(04963)
8040-00-664-2912
8040-01-016-4726
21
Adhesive, Uralane 8089,
Type 1
(90039)
8040-00-828-4936
21A
Adhesive, Rubber Base
EC-2126
MMM-A-1617
8040-00-281-1617
22
Alcohol, Methyl
OM232F
6810-00-275-6010
23
Alcohol, Isopropyl
TT-I-735
6810-00-855-6160
24
Aluminum Wool
MIL-A-4864, Type 2
5350-00-286-4851
25
Ammonium Nitrate Solution
(For removing Cadmium Plate)
MIL-A-175B
6810-01-091-8676
25A
Anodic, Black
MIL-A-8625, Type H
26
Antiseize Compound
MIL-A-907
8030-00-597-5367
26A
Deleted
27
Bag, Plastic
PPP-B-26
8105-00-054-0939
28
Barrier Material,
Grease-Proofed
MIL-B-121, Grade A
8135-00-224-8885
29
Barrier Material,
Flexible, Type 1
MIL-B-131
30
Barrier Material,
Waterproof, Vaporproof
MIL-B-131, Class 1
19
20
Change 73
1-12
8135-00-282-0565
TM 55-1520-234-23
Table 1-3. Consumable Maintenance Supplies and Materials (Cont)
Item
No.
Description
Ref. No.
and FSCM
NSN
30A
Bottle, Screw Cap, 16 oz.
S-8275G (64484)
6640-00-404-0660
30B
Bottle, Screw, Cap, 32 oz
S-8275H (64484)
6640-00-404-0661
31
Bungee Cord
MIL-C-5651, Type 3
8305-00-267-3119
32
Deleted
33
Cellophane
L-C-110
8135-00-392-8971
34
UseC12A
35
Cement (EC1357)
MMM-A-121
(76381)
8040-00-165-8614
36
Cheesecloth
CCC-C-440
(81348)
8305-00-267-3015
37
Chemical Film, Alodine
No. 1200
MIL-C-81706
Class 1A, Form III
8030-00-613-3131
38
Chromic Pickle Brush-On
Solution, Type 1
MIL-M-3171
39
Cleaning and Polishing
Compound, Biodegradable
PP-560
7930-00-634-5340
40
Cleaning Compound
MIL-C-372
6850-00-224-6658
41
Cleaning Compound,
Aircraft Surface,
Alkaline, Waterbase
MIL-C-25769
6850-00-935-0995
42
Cleaning Compound,
Aluminum Non-Flame
Sustaining
MIL-C-5410
6850-00-282-6770
43
Cleaning Compound, Oil
Cooler Solvent
MIL-C-6864
6850-00-551-3694
44
Cloth, Abrasive
P-C-451
5350-00-192-5050
44A
Cloth, Cotton Flannel
CCC-C-458
44B
Cloth, Screen, Abrasive
3-738-7011-0 (27712)
45
Cloth, Crocus
P-C-458
5350-00-221-0872
46
Cloth, Fiberglass, Type
C (0.010 inch thick)
MIL-C-9084
8505-00-530-0109
47
Cloth, Fiberglass Type
181
MIL-C-9084
48
Cord, Nylon, Type III
MIL-C-5040
4020-00-841-6583
49
Corrosion Preventive,
Aircraft Engine
MIL-C-6529, Type III
6850-00-281-2031
Change 54
1-13
TM 55-1520-234-23
Table 1-3. Consumable Maintenance Supplies and Materials (Cont)
Item
No.
Description
Ref. No.
and FSCM
NSN
50
Corrosion Preventive
Compound (Clear
protective coating)
204-011-449
(97499)
51
Corrosion Preventive
Compound, Cold
Application, Solvent
Cutback
MIL-C-16173, Grade 1
8030-00-231-2345
52
Corrosion Preventive
Compound, Cold
Application, Solvent
Cutback
Corrosion Preventive
Compound, Hot
Application
MIL-C-16173, Grade 11
8030-00-244-1297
MIL-C-11796, Class 3
8030-00-231-2353
55
Corrosion Preventive
Fingerprint Remover
MIL-C-15074
8030-00-664-4017
56
Cushioning Material,
Bound Fiber
PPP-C-1120 (Type VI,
Class A)
8135-00-292-9800
57
Deleted
58
Cushioning Material,
Packaging, Water
Resistant, Type II
PPP-C-843
8135-00-664-0366
58A
De-Icing/De-frosting Fluid
Type II
MIL-A-8243C
6850-01-039-3842
59
Desiccants, Packaging
Use and Static
Dehumidification,
Bagged, Activated
MIL-D-3464
6850-00-264-6573
59A
Detergent
Joy (76188)
7930-00-764-5066
59B
Detergent, General
Purpose
MIL-D-16791
7930-00-985-6911
60
Enamel, Acrylic, Heat
Resistant, X130
(06613)
61
Epoxy, Anti-Static
XA147
(06613)
62
Etchant, Tetra-etch
(17217)
6850-00-431-8662
62A
Ethylene Glycol, Monoethyl,
Ether Acetate
MILE-7125
6810-00-265-0563
63
Fabric, Synthetic, 5066
(89616)
64
Fairing Compound
(RP-1257-3)
(48909)
53
Change 77
8030-00-891-3113
1-14
TM 55-1520-234-23
Table 1-3. Consumable Maintenance Supplies and Materials (Cont)
Item
No.
Description
Ref. No.
and FSCM
NSN
64A
Filler Urethane Compound
K747 Blade
R.P. 3265 (02684)
(DC 182-72)
64B
Film, Teflon, Unperforated
No. E3760-P2
Richmond Corp
Highland, CA
65
Flannel Outing
CCC-F-466
66
Fuel, Turbine, Grades JP-4 and JP-5
MIL-T-5624
66A
Gasket, AsbestosJM89-3-64
(08835)
66B
Gloves, Cotton
MIL-G-3866
8415-00-268-8353
67
Grease, Aircraft, Oscillating Bearing
MIL-G-25537
9150-00-478-0055
68
Grease, Extreme Pressure
(Tube Pack)
(97499)
ASN-Tech 3913-G1
or 204-040-755-5
9150-00-506-8497
69
Grease (Lubriplate)
(73219)
9150-00-068-6268
70
Grease, Aircraft Multipurpose
MIL-G-81322
9150-00-944-8953
70.1
Grease, Aircraft Multipurpose
MIL-G-81322 14 oz. Cartridge
9150-01-622-3358
71
Grease, Silicone, Dow Corning II,
or Equivalent
8490010
(96717)
9150-00-616-9212
72
Grease Silicone
(#33 Medium)
(71984)
DC33 Fluid
9150-00-823-8048
73
Hydraulic Fluid, Petroleum Base
MIL-H-5606
9150-00-180-6181
73A
Hydraulic Fluid, Fire Retardant
MIL-H-83282
9150-00-149-7431
74
Hydraulic Fluid, Preservative,
Petroleum Base, Type 1
MIL-H-6083
9150-00-935-9807
74A
Ink, Indelible
74B
Insulating Compound
MIL-I-46058, Type UR
75
Lacquer, Acrylic, Gull Gray, 36231
1T-L-20
9150-00-515-1568
75A
Lacquer, Acrylic, Low Reflective,
Olive Drab
MIL-L-46159
8010-00-083-6588
76
Lacquer, Acrylic, Olive Drab,
Camouflage, X34087
MIL-L-81352
8010-01-033-8917
77
Lacquer, Acrylic, Orange-Yellow,
33538
MIL-L-81352
9150-00-166-3158
78
Lacquer, Acrylic Resin, Black
MIL-L.46159
8010-01-211-1106
79
Use C75
80
Lacquer, Non-Acrylic,
Black, Lusterless, 37038
81
Use C75
MIL-L-46159
Change 71
1-15
8030-01-229-2659
8305-00-641-5606
TM 55-1520-234-23
Table 1-3. Consumable Maintenance Supplies end Materials (Cont)
Item
No.
Description
81A
Lacquer
Light Full Gray
82
Lead Wool
83
Ref. No.
and FSCM
NSN
MIL-L-81352
(81349)
8010-00-935-7060
Lexcote
G-3397 (51747)
8030-00-118-1022
84
Lubricant
ASU-M100 (78511)
9150-00-823-8048
85
Lubricant, Solid Film. Type 2
MIIL-L-8937A (59595)
9150-00-834-5608
86
Material, Nylon Sandwich Buna.
N5200 5187
(89616)
87
Methyl-ethyl-ketone
TT-M-261
6810-00-281-2785
88
Naphtha, Aliphatic
TT-N-95, Type 2
6810-00-238-8119
89
Nitrogen
AFPIDN06830-1
90
Oil, Corrosion Preventive, Grade B
MIL-C-8188
6850-00-273-2395
91
Oil, Lubricating, Jet Engine
(Grade 1010)
MIL-L-6081
9150-00-273-2388
92
Oil, Lubricating, Low Temperature,
General Purpose
MIL-L-7870
9150-00-263-3490
93
Oil, Lubricating. Synthetic Base
MIL-L-7808
9150-00-782-2627
94
Oil, Lubricating, Synthetic Base
MIL-L-23699
95
Oil, Penetrating
VV-P-216
9150-00-261-7899
95A
Paint Aircraft Green,
Polyurethane
MIL-C-46168
8010-01-141-2420
96
Petrolatum, Lubrication
Vacuum Bagging
VV-P-236
9150-00-250-0926
97
Plastilube, Moly No. 3
(02307)
9150-00-141-4481
98
Polyurethane. Color Black
37038
MIL-C-83286
8010-00-482-5671
98A
Potassium Dichromate
98B
Potting Resin Uralite
No. 3148 A B
(01490)
99
Primer, A934B
(76500)
8040-00-943-2502
100
Primer. Epoxy
MIL-P-23377
8010-00-082-2450
101
Primer, Silicone Adhesive, 1200
SS4004 (0 1139)
8010-00-701-9616
Change 54
1-16
(8oz)
(1qt)
9150-00-180-6266
9150-00-985-7099
TM 55-1520-234-23
Table 1-3. Consumable Maintenance Supplies and Materials (Cont)
Item
No.
Ref. No.
and FSCM
Description
NSN
102
Primer, Zinc Chromate
1T-P-1757, MIL-P-8585
8010-00-297-0593
103
Prussion Blue Paste, Bearing Surface
(Thinned with oil)
MIL-P-30501
8010-00-281-4105
104
Putty, Zinc Chromate
MII.P-8116
8030-00-664-4968
105
Remover, Paint
qTT-R-248B
8010-00-515-2258
106
Remover, Paint, Epoxy System
MIL.R-81294
8010-00-926-1488
106A
Repair Kit, Infrared Suppression System
(97499) 2057067831
1560-00-103-3459
107
Resin Epoxy, Liquid EPON 828 and
Catalyst Diethylenetramine (DTA)
MIL-R-9300
O-D-1271
8040-00-822-6430
6810-00-995-4804
108
Rubber, Silicone, RTV
(71984)
8030-00-903-6566
109
Rubber, Silicone, Type I
(71983)
110
Rubber Strip, Type II, Grade A, Soft
MIL-R-6130
111
Rust Stripper, (For cleaning metal
prior to application of brush
cadmium plate)
(13429)
112
Sandpaper
P-101
5350-00-224-7215
113
Scotchbrite
(81348)
L-P-0050
7920-00-659-9175
114
Use C115
115
Sealing Compound, Low
Adhesion (Proseal 706)
MIL-S-8784
PROSEAL 706B2 (83527)
8030-00-616-9191
115A
Sealant (Pro Seal 700)
MIL-S-38249
8030-00-723-5345
116
Sealant (Proseal 890)
MILS-8802, Class B-2
8030-00-723-2746
117
Sealer Polysulfide
MILS7124
117A
Sealant No. EC 801
MIL-S-8802D Class B
118
Sealing Compound, Locking and
Retaining, Single Component
MILS-22473
118A
Sealing Compound
Grade H
MIL-S-22473
8030-00-081-2330
119
Use C115
119A
Sealing Compound
Brushable CLA-1/2
MIL-S-8784
8030-00-291-8380
Change 54
1-17
9320-00-814-4583
8030-00-275-8117
TM 55- 1520-234-23
Table 1-3. Consumable Maintenance Supplies and Materials (Cont)
Item
No.
Description
Ref. No.
and FSCM
NSN
119B
Sealing Compound
Brushable CLA-2
MIL-S-8784
8030-00-152-0062
119C
Sealing Compound
Extrudable CLB- 1/2
MIL-S-8784
8030-00-152-0022
119D
Sealing Compound
Extrudable CLB-2
MIL-S-8784
8030-00-680-2041
119E
Sealing Compound
PR-1422 B-2
(83574)
8030-01-154-9254
120
Shellac, Type I, Grade B, Body 1
TT-S-300
8010-00-577-4816
121
Silicone Compound
MIS-8660
6850-00-880-7616
122
Deleted
123
Smoother, Aerodynamic, EA 960,
Type I. RP-1257-3
EA960A-B (38564)
8010-00-006-7089
124
Solvent, Dry Cleaner
P-D-680, Type I
6850-00-264-9038
124A
Solvent, Safety
MIL-S-18718
125
Soap, Liquid
P-S-624
8520-00-228-0598
126
Soap, Toilet, Cake
P-20
8520-00-531-6484
127
Steel Wool
FF-W-1825
5350-00-240-2920
128
Stone, Sharpening (India)
(Novaculite)
SS-8-736, Type II
Clam B, Style 1 (81348)
5345-00-144-6894
129
Synthane Sheet (0.0175 inch thick)
MILP- 15037
5970-00-115-8838
129A
Tack Rag
(57687)
4940-01-198-9333
130
Tape, Antichafe Teflon
5490 (76381)
7510-00-923-0591
131
Tape, Dissimilar Metal Separation
MIIT-23142
7510-00-472-4021
132
Tape, Electrical, Black
MIL-I-24391
5970-00-419-4290
132A
Tape. Fastener Hook,
Velcro
MIL-L-21840
(81349)
8315-00-926-4931
133
Tape, Insulation Spiral Wrap
(0.006 x 1.0 inch)
MIII-18746
5970-00-935-0098
134
Tape, Masking
PPP-T-42
7510-00-290-2026
134A
Tape, Polyurethane,
Pressure Sensitive
Y9265A (76381)
751000-145-0171
134B
Tape, Tuck
Whittaker Corp.
Normco Material Division
600 Victoria Street
Costa Mesa, CA, 92627
Change 57
1-18
TM 55-1520-234-23
Table 1-3. Consumable Maintenance Supplies and Materials (Cont)
Item
No.
Ref. No.
and FSCM
Description
NSN
135
Tape, Multi-Purpose,
Double-Faced, Cloth
P50 (99742)
135A
Tape, Nylon, Wright Lon
No. 7400PS
International Plastics
Products, Carson, CA
136
Tape, Pressure Sensitive, Waterproof
PPP-T-60, Type II
7510-00-663-0199
137
Tape, Teflon, Self-Adhesive
MIL-I-23594
5970-00-812-7387
138
Tape, Vinyl, No. 473 (0.003 inch thick)
(76381)
8030-00-514-0981
139
Tedlar, Bondable Film
(31708)
140
Thinner, Acrylic Lacquer
MIL-T-19544
140A
Thinner, Polyurethane
MIL T-81772/AS
141
Thread Compound
MIIL-T-23361
8030-00-292-1102
142
Use C143A
142A
Toluene-Methyl
Isobutyl Ketone Mixture
MIL-T-19588
6810-00-286-0458
142B
Toluene
TT-T-548
6810-00-281-2002
Change 54
1-18A/(1-18B blank)
7510-00-584-2848
8010-00-527-2897
TM 55-1520-234-23
Table 1-3. Consumable Maintenance Supplies and Materials (Cont)
Item
No.
Ref. No.
and FSCM
Description
NSN
143
Trichlorotrifluoroethane
Cleaning Compound
MIL-C-81302
6850-00-033-8851
143A
Trichlorethylene,Technical
O-T-634
6810-00-184-4800
143B
Trichloroethane
O-T-620
6810-00-930-6311
144
Use C142B
145
Varnish, AlkaliResistant
MS35637-1
8010-00-697-7856
146
Versilube F50
MIL-S-81087
9150-00-082-5616
147
Walkway Coating, Type
1 Color 37038
MIL-W-5044
5610-00-641-0429
148
Walkway Coating, Type
2, Color 37038
MIL-W-5044
5610-00-641-0427
149
Deleted
150
Wire, Steel (Lockwire)
MS20995C20
9525-00-221-2650
151
Wire, Steel (Lockwire)
MS20995C32
9505-00-293-4208
152
Wire, Steel (Lockwire)
MS20995C41
9505-00-603-4120
153
Xylene
TT-X-916
6810-00-598-6600
154
Use C134A
155
Use C143B
156
Versilock 204
1270-4300
8040-01-184-1704
157
1-14.
Accelerator No. 5
Special Tools And Equipment.
Special tools and test equipment are listed in table 1-4
in alphanumeric order. Each tool or piece of test
equipment has an item number assigned for ease of
location and reference. When an item number is
unknown, you may locate special tools and test
equipment through alphanumeric arrangement within
the table. When an item is referenced in the manual,
8040-01-174-4684
you may locate the item through its T designator and
item number. T designators are used only with special
tools and test equipment. The special tools and test
equipment table is found only within this chapter;
therefore the table number will not be referenced within
the text. A complete listing of all special tools and test
equipment authorized for use to perform maintenance
on AH-1S aircraft/accessories are contained in the
aircraft parts manuals.
Table 1-4. Special Tools and Test Equipment
Item
No.
1
Part No.
AN8514-2
Nomenclature
Spanner Wrench
2
AN8516-1
Spanner Wrench
R/IN
3
ED0899
Test Fixture
T
Change 65
Usability Code
Calibration
R/IN
1-19
Figure
Reference
7-24
TM 55-1520-234-23
Table 1-4. Special Tools and Test Equipment (Cont)
Item
No.
Part No.
Nomenclature
Usability Code
Calibration
Figure
Reference
4
ED0933
Control Box
T
5
BH112JB53
Tester, Exhaust
Gas Temperature
T
6
BH16492
Temperature Indicator
Adapter
T
6A
BH16491
Temperature Trim
Adapter
T
7
LTCT 773
Engine Sling
R/IN
7A
M1A1
Gunner Control Quadrant
T
8
PD1201
Power Wrench
R/IN/D/A
9
PD1404
Power Wrench
Multiplier
D/A
10
PD1468
Adapter
D/A
11
PD1469
Socket
D/A
11A
PD1470
Socket
Collective
D/A
11B
PD1471
Extension
Socket WRN.
D/A
12
PD2658
Adapter
D/A
13
PD2659
Socket
R/IN/D/A
5-8
14
PD2660
Reaction Torque
Adapter
R/IN
5-8
15
S22
Bushing Tool
IN
16
S135
Packing Seating
Tool
IN
17
SWE13852-40
Adapter,
Transmission
R/IN
18
SW4738-2
Tie Down, Main
Rotor
19
T100921
Jackscrew
R
20
T100929
Jackscrew
R
21
T101220
Rotor Hoisting
Sling
R/IN
22
T101306
Splined Wench
D/A
6-6
23
T101307
Wrench
D/A
6-52
24
T101308
Jackscrew
R/D
25
T101338
Jackscrew
R
7 23
5-8
1-4
Change 38
1-20
TM 55-1520-234-23
Table 1-4. Special Tools and Test Equipment (Cont)
Item No.
Part No.
Nomenclature
Usability Code
Calibration
Figure
Reference
26
T101356
Build-up Bench
R/IN/D/A/AD
27
T101369
Support Plate
D/A
28
T101382
Ram Adapter
D/A
29
T101392
Wrench
D/A
30
T101401
Alignment Scope
AD
31
T101414
Socket Wrench
R/IN
32
T101419
Alignment Tool
Set
T
6-8
33
T101420
Holding Fixture
D/A
6-5
34
T101421
Adapter Plate
R/IN/D/A/AD
5-17
35
T101424
Holding Bar
R/IN
36
T101440
Jack Set
AD
37
T101447
Holding Fixture
D/A
38
T101449
Wrench
D/A
6-53
39
T101564
Plate
D/A
6-52
39A
T101560
Plate
D/A
652
40
T101456
Wrench
R/IN
41
T101467
Alignment Scope
Supp
AD
42
T101468
Flap Stop
AD
5-35
43
T101475
Bearing Remover
RP
5-32
44
T101485
Tab Bending Gage
AD
45
T101487
Arbor
RP
46
T101488
Wrench
D/A
47
T101491
Bearing Puller
RP
48
T101493
Wrench
R/IN
49
T101520
Hoist
R/IN
Change 2
1-21
5-17
6-7
5-33
5-31
1-9
TM 55-1520-234-23
Table 1-4. Special Tools and Test Equipment (Cont)
Item
No.
Part No.
Nomenclature
Usability Code
Calibration
Figure
Reference
50
T101524
Cyclic Stick
Fixture
AD
51
T101525
Tab Bender
AD
52
T101549
Fixture
D/A
53
T101550
Disassembly Tool
D
54
T101551
Puller
D
55
T101553
Holding Fixture
A
56
T101559
Grip Spacing Gage
AD
5-35
57
T101577
Staking Tool
RP
5-67
57A
T102095
Staking Tool Set
RP
11-1E,
Detail B
57B
T101873
Staking Tool Set
RP
11-1E,
Detail B
58
T101600
Wrench
D/A
6-53
59
T101864
Grip Lock
R/IN/D/A/AD
5-7
59A
114-99194
TSU Boresight Device
60
5120-EG-007
Pylon Damper
Assembly Tool
A
6-22
61
204-011-178-1
Clevis Assy
62
204-050-200-11
Ground Handling
Gear
63
50-T
Sealing Iron (19838)
RP
64
Deleted
66
209-030-195-1
Jack Fitting, Aft
I/T
66
209-030-245-1
Jack Fitting, Fwd
I/T
67
209-030-405-1
Jack Fitting, Wing
I/T
68
AN/USM-223
Multimeter
T
69
AN/USM-303A
Multimeter
T
70
209-070-532-1
Shield, Air Inlet
71
209-071-239-1
Wrench, Rack
Release
R/IN
16-6
72
209-071-244-1
Pin, Ground Safety
R/IN
16-6
72A
209-071-275
Ejector Rack Alignment Fixture
1-4
1-7
Change 38
1-22
TM 55-1520-234-23
Table 1-4. Special Tools and Test Equipment (Cont)
Item
No.
Usability Code
Calibration
Figure
Reference
Part No.
Nomenclature
73
2516
Arbor (Part of 7HEL153)
AD
5-55
74
2529
Pilot Bushing (Part of 7HEL153)
AD
5-55
75
2532
Fixture (Part of 7HEL153)
AD
5-55
76
2539
Post Assembly (Part of 7HEL153)
AD
5-55
76A
387991-003
Calibration Unit, Field
T
8-5
77
3091
Positioning Yoke (Squaring Plate)
(Part of 7HEL153)
AD
5-55
78
42M76
Transmission Stand
R/IN
79
7A050
Rotor Balance Kit
AD
5-38
80
7HEL054
Balancing Kit
AD
5-38
81
7HEL066
Adapter Kit, Balancing
AD
5-38
82
7HEL074
Squaring Plate
AD
5-55
83
7HEL153
Small Part Balance Kit
AD
5-55
84
76008-157
Release Wrench
R/IN
85
76008-159
Ground Safety Pin
R/IN
86
94251
Seal Installation Tool
IN
13-7
87
5120-AH1-001
Swashplate Alignment Tool
I
D-384
88
K747-404-1
Blade Repair Fixture
RP
5-17L
89
K747406-1
Cable Assembly
RP
5-17L
90
K747-409-1
Router Assembly
I
5-17M
91
K747-401-1
Blade Repair Tool Set
RP
92
1733
Puller, Circuit Board
R
93
TY9CL2STA
Gap Setting Gage
Change 73
6-54B
1-22A/(1-22B blank)
TM 55-1520-234-23
Table 1-4. Special Tools and Test Equipment (Cont)
Item
No.
94
* 95
* 96
* 97
* 98
* 99
* 100
* 101
102
103
104
105
106
107
Usability Code
Calibration
Part No.
Nomenclature
LT40
LT40-]
LT40-2
LT40-3
LT40-4
LT40-5
LT40-6
LT40-7
LT40-8
LT40-9
MIL-B-43714
1031-102A-351X
P508
K747-407-11
Bushing and Bearing Tool Kit
Nest
Ram Adapter
Bushing Tool
Bushing Tool
Bearing Tool
Staking Tool
Staking Tool
Piercing Tool
Piercing Tool
Balance Trip
Vacuum Pump
End Mill
End Mill
USABILITY CODES
R D I
RP T A IN AD S/P -
Figure
Reference
A/D/R
DR
DR
RP
RP
RP
5-17M
5-17M
1-15. Torque Procedures and Requirements.
Refer to TM 55-1500-204-25/1.
Removal
Disassembly
Inspection
Repair/Replace
Testing
Assembly
Installation
Adjustment
Storage/Preservation
1-16.
*Part of LT40 Bushing and Bearing Tool Kit (Item No. 94).
Change 65
1-23
Deleted.
TM 55-1520-234-23
Table 1-5. Adhesive Mix Ratio, Pot Life, and Cure Cycles
Change 42
1-24
TM 55-1520-234-23
1-17.
Dimensions and Tolerances.
a. Dimensions in this manual are normally in
inches and decimal fractions thereof unless otherwise
specified. Common fractions are used to refer to rivets,
cables, raw stock, and other items supplied in fractional
sizes, and sometimes for an estimated or nominal
dimension which cannot or need not be more precise.
Angles are stated in degrees and common fractions.
b. Tolerances on dimensions in decimal fractions
of an inch can be determined by the number of decimal
places, unless otherwise specified, as shown below.
Tolerances on angles are in a common fraction of a
degree.
TOLERANCES ON
DECIMALS
.XXX ± 0.010
TOLERANCES ON
ANGLES
± 1/2°
.XX ± 0.03
.X ± 0.1
Section II. LUBRICATION
1-18.
NOTE
Lubrication.
The lubrication chart consists of a main drawing
which is a perspective diagram of the helicopter, with
enlarged or detail views where required to show items
clearly. (See figure 1-2). The chart shows all parts
requiring periodic lubrications, except the engine and
transmission and tail rotor gearboxes, which are
lubricated by oil in accordance with servicing
instructions and Preventive Maintenance Inspection
Checklists.
Change 29
When a change of lubricant is made, it
is important that all lubricants previously
used be purged from the item before the
replacement lubricant is applied.
1-19.
Symbols.
The lubrication chart uses symbols and
abbreviations to indicate the required lubricant, method
of application, and time interval for lubrication of each
part listed. A key on the chart defines the meanings of
symbols and abbreviations.
1-28
TM 55-1520-234-23
Figure 1-2. Lubrication Chart (Sheet 1 of 4)
Change 56
1-29
TM 55-1520-234-23
Figure 1-2. Lubrication Chart (Sheet 2 of 4)
1-30 Change 56
TM 55-1520-234-23
Figure 1-2. Lubrication Chart (Sheet 3 of 4)
Change 56
1-30A/(1-30B Blank)
TM 55-1520-234-23
NOTES:
Failure of swashplate to accept grease requires investigation and
correction prior to releasing aircraft for flight. Perform swashplate
alignment check in accordance with the procedures contained in the
Special Inspection section.
Take care not to strike side of pylon structure when oscillating cylinder.
Do not lubricate bearings PN KSP6099-1.
Nuts must be completely removed from drive link attaching pin before
rotating main rotor blade. This will avoid damage to anti-drive link horn
on non-rotating part of swashplate.
1. Lubricate droop cam slider (15) lightly. Wipe excess off and out of slots. Slide contacts should have minimum
lubrication required to prevent dry contact without contributing to grit build-up. More frequent lubrication may be
necessary depending on environment and usage factors.
2. If MIL-G-25537 grease (C67) was used previously, purge with MIL-G-81322 grease (C70).
3. Disconnect rod end from pitch control lever or swashplate horn and oscillate cylinder assembly while greasing.
4. Rotate main rotor by hand and grease at approximately 30°intervals until assembly has been rotated one full turn to
ensure thorough purging of bearings. After lubrication, clean debris from boot, cut safety wire, raise boot and inspect to
ensure no grease has fallen on uniball. Clean grease from uniball, if necessary. Reinstall boot and safety.
5. Disconnect drive links. Rotate swashplate, grease at 30°intervals through 360°. Continue to grease until old grease
is purged. See Special Inspection section.
6. The lubrication interval for flexible couplings on seven tail rotor driveshaft couplings and two main driveshaft
couplings is as follows:
a. Inspect and lubricate flexible couplings in main and tail rotor drive systems at time of installation of couplings on
helicopter.
NOTE
This inspection and lubrication requirement will be accomplished on all
couplings NEW and USED.
b. Maximum interval from last lubrication is 600 hours (aircraft flying hours) or 12 months (from last lubrication
date).
c.
Make entry on DA Form 2408-18 to indicate date and aircraft flying hour of next inspection and lubrication due.
Figure 1-2. Lubrication chart (Sheet 4 of 4)
Change 56
1-31
TM 55-1520-234-23
Section III. HANDLING, JACKING, MOORING, HOISTING AND SLING LOADING
1-20. Ground Handling.
gyros and CN 998/ASN-43 directional
gyros. If helicopter must be moved
after shutdown and before 25
minutes have elapsed, power should
be reapplied to these gyros. After
power has been applied for five
minutes, helicopter may be safely
moved.
Before any work in cockpit area of a
helicopter with explosive canopy
removal system, ensure that ground
safety pins are installed in pilot's and
gunner's arming/firing mechanisms.
The helicopter can be equipped for towing by
attachment of two ground handling gear assemblies (6,
figure 1-3) on landing skids. Each assembly consists of
two wheels on offset axles, mounted to a supporting
cradle, with a hand operated hydraulic pump and two
rams for actuating the axle to extended position. The
cradle has a fixed pin at rear end, and a spring-loaded
pin at forward end, for mounting to eyebolt fittings on
landing gear.
The structural panels shown in figure 22 must be installed prior to helicopter
ground run, flight, or ground handling.
Ground handling includes hoisting, jacking, mooring,
parking, towing, and application of external electrical
power.
Pre-maintenance requirements for ground handling.
Condition
Requirements
Model
Part No. or Serial No.
Special Tools
Test Equipment
Support Equipment
Minimum Personnel
Required
Consumable Materials
Special Environmental
Conditions
AH-1S
All
(T61) (T65) (T66) (T67)
(T49)
None
Tow Bar, Maintenance
Hoist
Two
None
None
1-21. Towing.
Do not move helicopter for 25
minutes after power has been
removed from MD 1 displacement
a. Work Aid for Ground Handling Gear. A work
aid, for moving ground handling gear assemblies to and
from parked helicopters can be locally fabricated. (See
figure 1-4.) The device is a small tow bar, with lugs to fit
on mounting pins of ground handling gear which can
then be pulled or pushed on its own wheels.
b. Installation - Ground Handling Gear.
Keep clear of area above handling
gear as much as possible when
weight of helicopter is on wheels, to
avoid injury if mounting pins are not
securely engaged.
(1) Position handling gear assembly, with a
spring loaded pin forward, over landing gear skid
between eyebolts.
(2) Release enough hydraulic pressure by
turning T-handle of pump valve to allow alignment of
cradle mounting pins with eyebolts. Insert fixed pin in aft
eyebolt, then engage spring- loaded pin securely in
forward eyebolt. (Spring- loaded pin can be moved by
means of flat-headed release pin.)
(3) Install handling gear on opposite skid.
1-32 Change 48
TM 55-1520-234-23
(4) Station personnel at tail skid to steady
helicopter and to force tail boom down as handling
wheel pumps are actuated.
(5) On both sets of handling gear, close pump
valve and operate handle to extend wheels until skids
are raised.
c.
Towing Procedure.
(1) Check that both ground handling gear
assemblies are installed and extended.
(Refer to
paragraph 1-21b.)
CAUTION
Towing the helicopter on ground handling
gear over prepared surfaces at a gross
weight in excess of 9500 pounds will
cause permanent set in the aft cross tube.
Towing on unprepared surfaces must not
exceed 7500 pounds. Caution should be
exercised when towing on unprepared
surfaces
at
any
gross
weight.
Change 71
1-32A
TM 55-1520-234-23
Figure 1-3. Ground Handling Diagram (Sheet 1of 2)
1-32B
Change 71
TM 55-1520-234-23
Figure 1-3. Ground Handling Diagram (Sheet 2 of 2)
Change 71
1-33
TM 55-1520-234-23
Figure 1-4. Work air for towing ground handling gear
1-34 Change 38
TM 55-1520-234-23
(2) Attach tow bar (8, figure 1-3) to rings (9) on
forward ends of landing skids.
(3) Station a man at aft end of tail boom to
balance helicopter on handling gear and to assist in
control while towing.
All structural panels must be installed prior
to jacking and leveling. Do not lower jacks
until panels have been reinstalled
to
prevent
possible permanent set to
helicopter structure.
d. Removal - Ground Handling Gear
(1) Station a man at tail boom to assist by
steadying helicopter.
(2) Release hydraulic pressure, by turning Thandle of pump valve on each set of handling gear,
allowing wheels to retract and landing skids to rest on
ground. Close valve.
a. Remove three screws and washers from forward
jack point, under ammunition compartment rear
bulkhead (Station 138.7) approximately 10 inches right
of center. Install jack fitting (T66), using three bolts and
washers provided with fitting.
(3) Push release pin on front of cradle to
disengage spring-loaded mounting pin from eyebolt.
Pull rear pin free of eyebolt and remove handling gear
assembly. Remove opposite assembly.
b. Select jack fitting (T65) with threaded end
approximately 1.0 inch long, from shoulder to end.
Install fitting in aft jack point socket, on bottom of left
main beam just forward of tail boom.
c. Remove outboard armament pods. (Refer to
Chapter 16.) Check that two remaining jack fittings
(T67), for wing jack points, are similar to fitting used in
step (2) except for shorter threaded ends approximately
0.49 inch long.
(4) Remove tow-bar.
1-22. Jacking.
Four jack fittings with mooring shackles attached are
provided as loose equipment for use at two jack points
on the fuselage and on two outboard wing pylons. (See
figure 1-5). The forward jack fitting is attached by bolts
under the structure of the right main beam and the
ammunition compartment rear bulkhead. The aft jack
fitting is screwed into a socket on the left main beam
ahead of the tail boom attach splice. Wing pylon jack
fittings are substituted for ejector tube assemblies in
outboard armament racks.
Ejector cartridges must be removed before
installing jack fittings in outboard wing
pylons.
d. At each outboard rack, remove lockwire and
remove complete ejector tube assembly and install jack
fitting.
Do not jack aircraft in open area during
windy or gusty conditions.
e. Place jacks under four jack point fittings. If
removing landing gear, align all jacks with inboard legs
parallel at approximately 27 degrees to axis of fuselage.
Outboard articulated pylons must be in the
stowed position (four degrees up) when the
helicopter is to be jacked for any purpose.
Jacking fitting, (T67), must be installed.
High center of gravity makes it imperative
that all jacks must be raised evenly.
f.
Raise helicopter slowly and evenly.
g. Observe the
helicopter is on jacks.
1-35
following
precautions
while
TM 55-1520-234-23
Figure 1-5. Jacking and mooring fittings
1-36 Change 62
TM 55-1520-234-23
(1) All personnel in immediate area shall
exercise extreme caution not to bump or otherwise
disturb helicopter while raised or supported on jacks.
greater ease of movement of mobile fire fighting
equipment within area.
(2) Personnel shall not climb into or onto
helicopter while raised or supported on jacks.
d. Fire lanes having a minimum width of 50 feet
should be provided to cross main fire lanes and isolate
blocks of 10 helicopters or less.
(3) Rope off area around helicopter and
prominently display warning signs stating CAUTION:
THIS HELICOPTER IS ON JACKS.
e. Helicopters parked on concrete ramps or aprons
should be placed to utilize mooring rings when
available.
(4) After necessary work, lower helicopter
slowly and evenly. Remove jacks.
f. Parked helicopter will be provided with a static
ground.
(5) Remove jack fittings from outboard pylon
ejector racks. Reinstall ejector tubes. Torque to a
range of 290 TO 310 INCH-POUNDS.
Reinstall
lockwire. Reinstall cartridges. (Refer to chapter 16.)
g. Under normal conditions park the helicopter as
follows:
(6) Remove forward jack fitting with bolts and
washers. Reinstall screws with washers in jack point
bolt holes.
(1) Park helicopter on a level surface,
whenever possible, so that load will be evenly
distributed on landing gear.
(2) Retract or remove ground handling wheels
to allow helicopter to rest on landing skids.
(7) Remove aft jack fitting. Return all fittings to
loose equipment.
h. Install outboard armament pods.
Chapter 16.)
NOTE
(Refer to
If helicopter .is to remain parked more than
14 days, use suitable blocks or shoring to
raise skids slightly off supporting surface.
1-23. Parking.
Parking as used in this manual, is defined as
condition in which helicopter will be secured while on the
ground. Direction of heading and location of helicopter
is normally determined by ease of maintenance and
servicing; to allow removal of any one helicopter from
parking area, and to permit ready access or mobile fire
fighting equipment within area.
Although parking
arrangements may vary according to local facilities, the
following general procedure should be observed.
Do not use a rope to pull rotor blades in- to
alignment for tiedown. Damage to leading
edge may result. Manually position rotor
blades.
(3) Align main rotor blades fore and aft and
horizontal, and tail rotor blades parallel to vertical fin.
a. Double-row lateral parking, with front and rear
helicopter of each double row placed tail to tail, should
be used where possible.
NOTE
If the collective stick is positioned in other
than full down, turn blades in the direction of
rotation and fully lower collective prior to
main rotor tiedown.
b. Helicopter should be parked not less than 750
feet from ends of center line of nearest runway, and not
less than 250 feet from edge of connecting taxi strips.
NOTE
c. Width of fire lanes between each double row
should be slightly greater than rotor span of parked
helicopters. This spacing will facilitate removal of any
helicopter from parking area, as well as permitting
Change 62
Use 1/2 inch polyester rope, NSN 4020-00630-4873, for blade tie-down. Reference
TM
1-1520-250-23-1
for
additional
information.
1-37
TM 55-1520-234-23
(4) Engage hook of main rotor tie-down (3,
figure 1-3) in hole of fitting on each end of rotor blade
and position blade above tail boom. Pull on tiedown to
remove the spanwise slack from the rotor system and
secure rotor by wrapping tiedown rope firmly around
tailboom (see figure 1-3). Tie forward tiedown rope to
tow rings on landing gear skid.
(5) Attach tail rotor tiedown rope (4) to tail rotor
and secure to loop provided on side of vertical fin.
Additional security of the main rotor tiedown can be
accomplished by inserting an AN416-2 safety pin
through a 0.060 inch hole drilled through the hook of the
main rotor tiedown. The hole is drilled perpendicular to
the plane of the handle 1/4 inch from the insertion end
of the hook. Secure the safety pin to the hook handle
with a six inch piece of NAS1455B30-6 chain and safety
wire. Insert the pin through the hook after inserting the
hook through the rotor blade fitting.
(1) Park helicopter.
(2) Moor
paragraph 1-24.
helicopter
in
accordance
with
(3) Fill fuel cell to capacity if time permits.
(4) Disconnect battery.
Secure all loose
equipment. Moor all ground support equipment at safe
distance from helicopter.
(5) After high winds have passed, inspect
helicopter for damage from flying objects.
1-24. Mooring.
1-24A. Mooring Hardware (Table 1-6).
1-24B. Mooring Procedure On Unpaved Surface.
Mooring is a process of securing parked helicopter
to avoid damage by high winds or turbulent weather.
Mooring fittings are provided on the four jack pad
fittings. Where properly spaced rings are not available,
mooring can be accomplished with a standard mooring
kit.
Before entry into cockpit area, ensure that
canopy removal system ground safety pins
are installed in pilots and gunners
arming/firing mechanisms.
a. Park helicopter on unpaved parking area,
headed in direction of highest winds forecast.
(6) Check that all switches are OFF and
external power disconnected, and close all doors and
access plates. Lock ignition and canopy doors. Remove
keys.
(7) Install pitot tube cover (2, figure 1-6), engine
air inlet shield (4), and exhaust cover (3).
b. Screw anchor rod (1, figure 1-7) into arrow (3).
c. Slip driving rod (2) over anchor rod and into
socket of arrow.
d. Turn cam of driving rod so that prongs of arrow
are not spread by driving.
NOTE
e. If necessary, loosen surface of ground.
If required and available, install canopy
covers (1).
f.
h. Under turbulent weather conditions park the
helicopter as follows:
Position driving rods as shown in figure 1-7.
g. Drive each arrow into ground until driving rod
handle is approximately three inches above surface.
h. Rotate driving rod handle approximately 90
degrees and give it a sharp blow to spread arrow
prongs.
Structural damage can occur from turbulent
weather conditions. Anchoring and mooring
should be accomplished when wind is
expected to exceed 45 knots per hour.
When possible, helicopter should be
evacuated to a safe weather area if a
tornado, hurricane, or wind condition above
75 knots is expected.
i. Return driving rod to driving position and
remove it from anchor rod.
j. Align squared socket of eye (4) with squared
end of anchor rod. Fit in place and tighten knurled nut.
k. Set arrow prongs by pulling up on eye
assembly.
1-38 Change 62
TM 55-1520-234-23
Do not overtighten tie-down cable or rope.
Overtightening may cause the mooring
fitting bolt to bend.
l. Secure helicopter with quarter inch cables or
one-inch manila rope.
NOTE
When anchor rods are no longer needed,
they may be removed by turning eye
assemblies counterclockwise, leaving
arrows in ground.
1-24C. Mooring Procedure on Paved Surface.
b. Place the hook-ends of the two forward chains
into the mooring clevis. Adjust the chains with the chain
adjusters provided on each chain. Chains should be
adjusted to the point where the slack has been removed.
c. Remove the fairing covering the aft jack-point.
Four flush head bolts must be removed. With the fairing
removed install the jack-point in the uncovered recess
and install the aft mooring clevis on the jack-point as
described in this TM. A Frearson head screw driver will
be required.
d. Place the hook-ends of the two aft mooring
chains into the aft mooring ring. Place the hook-end of
the two center mooring chains into the most aft of the
three D- rings provided under the aircraft wings. Adjust
the chains with the chain adjusters provided with each
chain. Chains should be adjusted to the point where the
slack has been removed.
a. Position the aircraft on the mooring pad with the
longitudinal centerline of the aircraft directly above and
parallel to the longitudinal axis of the pad as shown in
figure 1-7A. The aft mooring ring is to be located
directly opposite the center pair of mooring points on the
pad as illustrated.
NOTE
It is highly recommended that AH-I
helicopters be flown with the mooring
hardware installed at all times to permit a
rapid response to weather emergencies,
unless it is the commander's decision that to
fly without the fairings would significantly
impact the mission.
NOTE
It will be necessary to remove the fairing
which covers the forward jack-point, by
removing the flush head bolts which secure
the fairing to the aircraft. With the fairing
removed, install the stainless steel jackpoint in the uncovered recess, as described
in this TM. With the jack-point installed,
install the mooring clevis to the jack-point as
described in this TM. A mechanics tool kit
will be required.
NOTE
The mooring hardware is not considered
flyaway equipment.
All active mooring
points shall be equipped with this hardware.
MOORING HARDWARE
ITEM
DESCRIPTION
P/N
NSN
QUANTITY
1
CHAIN ADJUSTER
MB-I
1670-00-212-1149
6
2
CHAIN WITH HOOK
FOR MB-1
4010-00-516-8405
12
TABLE 1-6. MOORING HARDWARE CHART
Change 62
1-38A/(1-38B blank)
TM 55-1520-234-23
Figure 1-6. Covers diagram
1-39
TM 55-1520-234-23
Figure 1-7. Mooring diagram
1-40
TM 55-1520-234-23
Figure 1-7A. AH-1 Paved Surface Mooring Configuration
Change 62
1-40A/(1-40B blank)
TM 55-1520-234-23
1-25. Hoisting
The entire helicopter can be lifted by a suitable hoist
attached to an eye provided on the main rotor retaining
nut at top of the mast. This hoisting point can also be
used to lift out the mast assembly (with or without the
main rotor and rotating controls assemblies), or the
complete mast and transmission assembly.
(c) Insert lower end of tube assembly (1)
into hub assembly (8). Align lower fitting of hub in
support bracket and install pin (9).
(d) Raise hoist assembly. Attach upper
fitting of hub to upper supports, using two bolts and
washers instead of screws previously removed.
a. Hoisting Helicopter.
(1) Attach a hoisting cable and clevis or clevis
assembly (T61) to lifting eye of main rotor retaining nut
(figure 1-3). Connect a suitable hoist and take up slack.
(2) Station a man at tail skid to steady
helicopter against swinging or turning when hoisted. If
lifting beyond reach, attach a suitable rope for this
purpose.
(3) Hoist slowly, maintaining a steady lifting
force.
b. Maintenance Hoist.
(Figure 1-8).
A
maintenance hoist (T49), designed to mount on left side
of the fuselage, is provided for use in removing and
installing main rotor, mast, transmission, or engine
assemblies. The hoist consists of a tube assembly, a
hub assembly, and attaching parts. The tube assembly
has a hand-operated winch, with cable, pulleys and
weighted hook. The hub is a socket made from larger
diameter tubing, with attachment fittings, sleeve
bearings, and a platform to aid the operator. The tube
assembly rests on a steel ball in the hub, and can be
rotated by means of the crossbar handle to move the
hook into position.
(1) Installation. (Figure 1-8).
(a) At left side of fuselage, remove two
screws and washers from bolt holes of upper hoist
supports, located just ahead of engine forward firewall.
Remove six screws and washers from lower support bolt
holes, located in vertical rows of three ahead of and
behind landing gear aft crosstube.
Handle hoist with care to avoid personal
injury or damage to aircraft.
(2) Removal.
(a) Detach hub fitting from upper supports
by removing two bolts and washers.
(b) Carefully swing top of hoist assembly
outward and down until resting on ground. Remove
tube assembly (1) from hub assembly (8).
(c) Detach lower fitting of hub from bracket
(10) by pulling out pin (9). Remove hub assembly.
(d) Remove bracket with attaching bolts
and washers from fuselage. Bracket can be attached
with pin to hub for convenience.
(e) Reinstall screws and washers in bolt
holes of upper and lower support points.
1-26. Leveling.
Leveling lugs located in the ammunition
compartment floor are used with a bubble protractor
when it is necessary to level the helicopter. Pallet must
be removed for access. For fore-and-aft leveling, use
two lugs (10, figure 1-3), located in depression near left
side of floor. For lateral leveling, use two lugs (9),
located on top of aft ammunition pallet track. Apply
jacking procedures to correct helicopter position.
1-26A. Leveling Pads Replacement.
NOTE
Prior to installation of maintenance hoist
inspect upper hoist support visually for
cracks and other damage which may affect
function of the support. Replace upper
hoist support if damaged.
Do not attempt to replace leveling pads if
structure in pylon area of helicopter is
damaged.
Send helicopter having
damaged pylon structure to depot.
(b) Install bracket (10) above landing gear
cross- tube, using six bolts and washers instead of
screws removed in preceding step.
Change 56
1-41
TM 56-1520-234-23
Figure 1-8. Maintenance hoist T101620
1-42
TM 55-1520-234-23
NOTE
(2) Use scotch-brite (Cl13) to clean area.
Any or all leveling pads may be replaced as
necessary.
(3) Use adhesive (C12) to bond new leveling
shim assembly in place. Allow adhesive to dry.
a. Open transmission cowl assembly (17, figure 2-
(4) Place bubble protractor on leveling pads.
Check fore-and-aft level of leveling pads.
b. Check structure in pylon area including lift beam
(12, figure 2-69), pylon support (10) and fifth mount
support (14).
(5) As necessary, peel lamination from shim
until leveling in fore-and-aft direction in accomplished.
3).
c. Place bubble protractor in fore-and-aft direction
on lift beam (12).
k. If any or all lateral leveling pads are damaged or
missing, replace as follows:
(1) Remove
five
screws
and
remove
ammunition floor track having damaged or missing
leveling pads.
NOTE
Fore-and-aft member of pylon support (10)
may be used instead of lift beam (12).
(2) Place new ammunition floor track in position
and attach with five screws.
d. Jack up helicopter. Refer to paragraph 1-22.
e. Adjust jacks to level helicopter in fore-and- aft
direction. Use bubble protractor to check level.
f. Place bubble protractor in lateral direction on lift
beam (12) or fore-and-aft member of pylon support (10).
g. Adjust jacks to level helicopter in lateral
direction. Use bubble protractor to check level.
h. Repeat steps c and e through g to make sure
helicopter is level.
i.
2-3).
(3) Place bubble protractor on leveling pads of
new ammunition floor track. Check lateral level of
leveling pads.
(4) Place shim under ammunition floor track as
necessary and repeat steps (3) and (4) until leveling in
lateral direction is accomplished.
l. Check leveling in both directions to insure that it
is satisfactory.
m. Remove bubble protractor and other tools.
Lower helicopter slowly and evenly.
Open ammunition compartment doors (8, figure
n. Remove jacks (paragraph 1-22).
j. If either or both fore-and-aft leveling pads are
damaged or missing, replace as follows:
o.
Close access doors.
1-27. Retrieval of Disabled Helicopter
(1) Remove any remaining portion of leveling
shim assembly.
Change 56
(Refer to FM 55-413).
1-42A/(1-42B blank)
TM 55-1520-234-23
1-28. Application of External Power.
An external power receptacle (5, figure 1-3) for
application of external 28V DC power is located in left
side of the fuselage at station 274, covered by a springloaded access door. When the door is open, a switch
causes the EXTERNAL POWER caution panel segment
to be lighted. Battery switch should be at OFF position.
Use a 28V DC power source capable of delivering 650
to 800 amperes. When cable connector from power
source is connected to the receptacle, the external
power relay in the helicopter DC circuit will be energized
and power will be supplied to the main bus for
distribution.
NOTE
If battery charge is less than 24 volts,
external power may be required to avert hot
starts.
Section IV. INSPECTION REQUIREMENTS
1-29. General Information.
This section contains complete requirements for
special inspections, overhaul and retirement schedule,
and standards of serviceability applicable to the aircraft.
Daily inspections are contained in TM 55-1500-220PMD, Preventive Maintenance Daily.
Phased
maintenance inspections are contained in TM 55.1500220-PM, Phased Maintenance Checklists.
The
inspections prescribed in this chapter shall be
accomplished at specific periods by aviation unit
maintenance activities with the assistance of
intermediate maintenance activities when required.
airframe inspection intervals. Refer to DA PAM 738-751
for applicable forms, records, and worksheets required
for these inspection intervals. Typical of this type
inspection items are:
a. Special inspection frequencies that are based
on flight hours may be accomplished within a plus or
minus ten percent tolerance from the nominal time when
such inspections would ordinarily be due.
b. Inspection of components or in frame on a
calendar basis: first aid kits, weight and balance check,
aircraft inventory, etc.
b. Special inspections based on calendar times will
have a tolerance of plus or minus ten percent not to
exceed thirty days.
a. An inspection which is contingent upon specific
conditions or incidents that arise, and only because of
these conditions or incidents, immediate inspection is
required to ensure safe flight.
Typical of these
conditions are hard landings, overspeed, and sudden
stoppage.
c. Refer to DA PAM 738-751 for applicable forms,
records, and worksheets.
1-30. Standard of Servlceabllty.
When components are being removed from
an aircraft, all inspections required by the
next phase maintenance inspection must be
accomplished prior to either immediate reuse or storage.
Upon installation, the
component will be inspected in accordance
with the current phase (either that phase the
receiving aircraft is in or if in between
phase, the last phase performed). This will
ensure that a re-used component will not
overfly any PM inspections, and that it will
be properly interfaced with the receiving
aircraft phase sequence.
Standards of Serviceability to be utilized in the dayto-day inspection and maintenance of the aircraft can be
found as fits, tolerances, wear limits, and specifications
in the aircraft maintenance manuals. Standards of
serviceability for transfer to aircraft are contained in TM
55-1500-326-25.
1-31. Special Inspection.
This section supplements the scheduled inspections
as contained in TM 55-1500-220-PM Phased
Maintenance requirements. This section also includes
inspection of items which are required to be inspected at
intervals not compatible with airframe operating time or
Change 39
1-43
TM 55-1520-234-23
Figure 1-9. Inspection area diagram (Sheet 1 of 2)
1-44
TM 55-1520-234-23
AREA NO. 6:
Main Rotor
Area
All components of the main rotor I b and blade. Does not include the
mast.
AREA NO. 7:
Pylon Area
All surfaces, components and equipment contained in, and on the exterior of, the hydraulic and transmission compartments to the bottom
of the transmission. Includes transmission cowling, mast, mounts,
rotating controls, and main (input) drive shaft.
AREA NO. 8:
Wing Area
All surfaces, components and equipment in and on the wings. Includes
all external fittings and attachments.
AREA NO. 9:
Center
Fuselage
Area
All surfaces, components and equipment in and on the fuselage below
the engine deck (WL 65.00) and between the cabin area (Sta 186.25)
and tail boom attachment bulkhead (Sta 299.57). Includes forward and
aft fuel cells, compartment below transmission, oil cooler and compartments accessible through side doors and panels on fuselage.
AREA NO. 10:
Engine
Area
All surfaces, components and equipment associated with engine installation located above engine deck (WL 65.00) and within engine cowling,
tailpipe fairing, and aft fairing.
AREA NO. 11:
Tail Boom
Area
All surfaces, components and equipment located in and on the tail
boom and vertical fin. Includes tail rotor, synchronized elevator,
control linkages, drive shafts, gearboxes, electronic gear and cooling fan.
209900-16-2D
Figure 1-9. Inspection area diagram (Sheet 2 of 2)
1-45
TM 55-1520-234-23
1-46
TM 55-1520-234-23
1-47
TM 55-1520-234-23
1-48 Change 71
TM 55-1520-234-23
AIRCRAFT INSPECTION CHECKSHEET
AIRCRAFT AND SERIAL NO.
AREA
NO.
REQUIREMENT
EVERY
SPECIAL INSPECTION
INSPECTION NO.
u.
Inspect lower wire strike cutter for bends, cracks, and alignment. Replace
cutter if damage is found.
6
K747 MAIN ROTOR BLADE SUBJECTED TO HIGH WINDS
Blades unrestrained and/or have torn loose from their mooring when subjected to
winds of 60 MPH and higher are to be tap tested in main spar areas between
stations 70 to 90.
SUDDEN STOPPAGE
Sudden Stoppage is defined as an instantaneous shock load applied to the drive
train and rotor systems either POWER ON or POWER OFF. Shock loads result
from:
a.
Main rotor blade(s) striking a movable or immovable object.
b.
Tail rotor blade(s) striking a movable or immovable object.
c.
Seizures which occur as a result of an internal failure of a drive train/rotor
system component.
d.
Engine compressor stall.
After a Sudden Stoppage event has occurred, one of the following special
inspections shall be conducted depending on the origin of the shock load. All
components removed for overhaul as a result of these inspections must be tagged
to call attention to the nature of the incident.
Change 75
1-48A / (1-48B blank
NO. OF PAGES
DATE OF INSPECTION
STATUS
ITEM
2
6, 7,
10& 11
PAGE NO.
3A
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
AIRCRAFT INSPECTION CHECKSHEET
AIRCRAFT AND SERIAL NO.
AREA
NO.
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
INSPECTION NO.
REQUIREMENT
EVERY
a.
ITEM
Main Rotor Blade Strike.
(1)
No visible damage to either blade.
(a) Wipe upper and lower surfaces of main rotor blades with a
clean, soft cloth and inspect both surfaces for cracks,
distortion or bond separation.
(b)
Visually inspect hub assembly and mast for damage.
(c)
(2)
If no damage is found, inspection is complete. If damage is
found in either of the above inspections proceed to
paragraph 2 below:
Minor Damage to Either Blade.
NOTE
This category includes both field repairable damage and skin
tears whether repairable or not.
(a)
Inspect and replace the following items if damage is
found:
1.
Main rotor hub trunnion cap attach bolts and drag
brace jamnuts and attach bolts for security.
2.
Flight control system, from the rotor to servo cylinder,
for bent or damaged tubes.
3.
Scissors levers drive links for damage.
4.
Swashplate gimbal mounting for damage.
5.
Collective friction collet assembly for free travel.
6.
Structure at transmission mounting points (use tenpower magnifying glass) for cracks.
7.
Lift link and structure for damage, security and
distortion.
8.
Main driveshaft.
9.
Mast.
Change 2
1-49
PAGE NO.
4
NO. OF PAGES
24
DATE OF INSPECTION
STATUS
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
INSPECTION NO.
AIRCRAFT AND SERIAL NO.
AREA
NO.
REQUIREMENT
EVERY
ITEM
10. Transmission sump oil filter, external oil filter and
chip detector for metal particles.
a. Positive indications are cause for replacing
transmission.
b. If no metal particles are found continue operation
for 5 hours, then repeat inspection. If no positive
indications are found, resume normal operation.
11. 42 degree and 90 degree gearboxes for metal particles.
12. Tail rotor driveshafts and hanger assemblies for
obvious damage.
13. Tail rotor hub and blade assemblies.
(3)
(b)
Repair/replace blades as required.
(c)
Inspection complete.
Major damage to either blade.
NOTE
This category is restricted to non-repairable damage other than
skin tears. For skin damage, see Minor Damage Inspection.
(a)
Replace the following: (Disposition as noted)
1.
Main rotor hub assembly (overhaul).
2.
Main rotor blades scrap).
3.
Mast (overhaul).
4.
Swashplate (overhaul).
5.
Scissors and sleeve assembly (scrap).
6.
Control rods (rotor to scissors levers) (scrap).
7.
Transmission (overhaul).
1-50 Change 29
PAGE NO.
5
NO. OF PAGES
24
DATE OF INSPECTION
STATUS
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
INSPECTION NO.
AIRCRAFT AND SERIAL NO.
AREA
NO.
REQUIREMENT
EVERY
ITEM
8.
(b)
(c)
b.
Engine. Refer to TM 556-2840-229-23 for required
inspection.
Inspect and repair/replace the following as required:
1.
Tail rotor blades.
2.
Intermediate and tail rotor drive gearboxes (inspect
for damage to gears and input/output couplings).
Inspection complete.
Tail Rotor Blade Strike.
(1)
No visible damage to either blade.
(a)
Inspect doublers for bonding separation, attachment area
for distortion.
(b)
Inspect and replace the following items if damage is
found.
(c)
1
Tail rotor hub assembly.
2
Tail rotor rotating controls.
3
42 degree and 90 degree gearboxes (inspect for metal
particles).
4
Tail rotor driveshafts and hangers.
If no damage is found, inspection complete. If damage is
found proceed to following paragraph (2).
Change 29
1-50A
PAGE NO.
6
NO. OF PAGES
24
DATE OF INSPECTION
STATUS
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
INSPECTION NO.
AIRCRAFT AND SERIAL NO.
AREA
NO.
REQUIREMENT
EVERY
(2)
TEST
Visible damage to either blade.
(a)
Scrap both blades.
(b)
Replace 42 degree and 90 degree gearboxes and return for
overhaul.
(c)
Inspect and replace the following items if damage is
found.
1
Deleted
2
Tail rotor hub assembly.
3
Tail rotor rotating controls.
4
Tail rotor driveshafts.
5
Tail rotor hanger assemblies (inspect for internal
spline and curvic coupling damage).
6
Transmission sump oil filter, external oil filter and
chip detector for metal particles.
a. Positive indications are cause for replacing
transmission.
b. If no metal particles are found, continue operation
for 5 hours, then repeat inspection. If no positive
indications are found, resume normal operation.
7
Main driveshaft.
8
Tailboom attachment points.
9
Mast assembly.
10 Main rotor rotating controls.
11 Main rotor blades.
12 Main rotor hub trunnion cap attach bolts and drag
brace jamnuts for security.
13 Transmission tail rotor output quill.
(3)
Inspection complete.
1-50B
Change 11
PAGE NO.
6A
NO. OF PAGES
24
DATE OF INSPECTION
STATUS
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
INSPECTION NO.
AIRCRAFT AND SERIAL NO.
AREA
NO.
REQUIREMENT
EVERY
c.
(2)
Replace the following (disposition as noted):
(a)
Transmission (overhaul).
(b)
Mast assembly (overhaul).
(c)
42 degree gearbox (overhaul).
(d)
90 degree gearbox (overhaul).
(e)
Tail rotor driveshafts and hanger assemblies (overhaul).
(f)
Engine (refer to TM 55-2840-229-23).
(g)
Main rotor hub assembly (overhaul).
Inspect and repair/replace the following as required:
(a)
Main rotor blades.
(b)
Main rotor rotating controls.
(c)
Tail rotor blades.
(d)
Tail rotor hub assembly.
(e)
Main driveshaft (inspect for internal and curvic coupling
damage).
(f)
Deleted
(g)
Deleted
(h)
Helicopter structure.
Compressor Stall. Engine compressor stall surge is characterized by
a sharp rumble or a series of loud, sharp reports, severe engine
vibration and a rapid rise in exhaust gas temperature (egt) depending
on the severity of the surge.
(1)
Perform engine compressor stall inspection in accordance with
TM 55-2840-229-23.
Change 54
1-50C / (1-50D blank)
NO. OF PAGES
24
DATE OF INSPECTION
STATUS
Internal Failure of Drive Train/Rotor System Component.
(1)
d.
TEST
PAGE NO.
6B
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
INSPECTION NO.
AIRCRAFT AND SERIAL NO.
AREA
NO.
REQUIREMENT
EVERY
(2)
(3)
TEST
(a)
No damage to 90 degree gearbox. Visually inspect remaining
tail rotor drive shaft components. If no damage is found,
inspection complete.
(b)
Damage to 90 degree gearbox or other drivetrain component:
Perform inspection requirements of subparagraph (3).
Inspect and replace the following items if damage is found:
(a)
42 degree gearbox (inspect for damage to gears, unusual wear
pattern on either coast or drive side of gears and damage to
input/output coupling internal and curvic coupling splines)
(b)
Tail rotor hanger assemblies (inspect for internal spline and
curvic coupling damage.
(c)
Tail rotor drive shafts (loose rivets, bent).
(d)
Main rotor driveshaft (bent or twisted).
(e)
Transmission sump oil filter, external oil filter and chip
detector for metal particles.
1
Positive indications are cause for replacing transmission.
2
If no metal particles are found, continue operation for five
hours and then repeat inspection. If no positive indications
are found, resume normal operation.
(f)
Mast assembly.
(g)
Helicopter structure including tailboom attachment area and
vertical fin.
1-51
NO. OF PAGES
24
DATE OF INSPECTION
STATUS
Inspect 90 degree gearbox for damage to gears, unusual wear pattern
on either coast or drive side of gears and damage to input/out
coupling internal and curvic coupling splines.
Change 7
PAGE NO.
6C
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
INSPECTION NO.
AIRCRAFT AND SERIAL NO.
AREA
NO.
REQUIREMENT
EVERY
(4)
6&11
TEST
(h)
Replace main rotor hub trunnion attach bolts.
(i)
Tail rotor blades.
(j)
Tail rotor hub assembly.
Inspection complete.
AFTER MAIN ROTOR OVERSPEED
Inspection and/or replacements are required after any report that main
rotor has exceeded 339 rpm limit. When 356rpm has been exceeded,
additional requirements apply.
Main Rotor Overspeed Less Than 356RPM:
Inspect the following:
a.
Main rotor blades for damage, bond separation and distortion.
b.
Tail rotor blades for damage, bond separation and distortion.
Main Rotor Overspeed Exceeding356RPM:
a.
Inspect main rotor blades as follows:
(1)
Visually inspect blade skins. Any wrinkle or deformation is
cause for blade replacement.
(2)
Remove tip cap and inspect balance weights. Deformation of
weights and/or studs is cause for blade replacement. Loose
weights is not a cause for blade replacement. Inspect stud
retention nuts for looseness by applying 30 inch pound torque. Torque loose stud retention nuts to 130 to 145 inch
pounds.
(3)
Blades which pass these inspections are acceptable for further
service. Return faulty blades to depot activity with details of
overspeed incident.
b.
Replace main rotor hub assembly. Send removed hub to overhaul
facility, with information on overspeed incident. Bolts should
remain with hub.
c.
Visually inspect blade retention bolts and drag brace bolts for
shear offset.
1-52 Change 41
PAGE NO.
7
NO. OF PAGES
24
DATE OF INSPECTION
STATUS
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
INSPECTION NO.
AIRCRAFT AND SERIAL NO.
AREA
NO.
REQUIREMENT
EVERY
d.
TEST
(1)
Bond separation anywhere on the blade is cause for replacing
blades. Send removed blades to next higher maintenance level
for evaluation and possible repair.
(2)
If any movement of the tip or root end balance weights has
occurred, scrap the blade.
(3)
Check the retention bushings for evidence of looseness. If any
bushing is loose scrap the blade.
(4)
If blade passes the above inspection requirements and no other
discrepancies exist, then the blade is serviceable.
e.
Perform a thorough visual inspection of tail rotor hub. If no discrepancies are found, the hub may be retained in service.
f.
Visually inspect the following components, which may be considered
satisfactory for continued use if no visible damage is found:
Transmission Assembly
(2)
42 Degree Gearbox
(3)
90 Degree Gearbox
(4)
Tail Rotor Driveshafts and Hangers
(5)
Main Rotor Driveshaft
(6)
Mast
(7)
Swashplate Assembly
(8)
Scissors and Sleeve Assembly
(9)
Tail Rotor Hub
Change 71
1-53
NO. OF PAGES
24
DATE OF INSPECTION
STATUS
Inspect tail rotor blades:
(1)
PAGE NO.
8
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
INSPECTION NO.
AIRCRAFT AND SERIAL NO.
AREA
NO.
6
REQUIREMENT
EVERY
TEST
MAIN ROTOR HUB INSPECTION
The following inspection shall be performed whenever external indications
of hub problems exist, i.e., unusual noises, excessive heat, vibrations, etc.
a.
b.
6
Remove hub assembly from aircraft and disassemble to the extent
required in order to determine the serviceability of the following
components.
(1)
Feathering axis bearings.
(2)
Extension sleeves.
(3)
Radius rings.
(4)
Inboard bearing housing seals.
(5)
Outboard dust seals.
(6)
Flapping axis bearings, trunnion sleeves, and dust seals.
Replace items as required, reassemble hub assembly, and reinstall on
aircraft.
EVERY 20-25 HOURS
a.
Deleted.
_________
WARNING
Do not flap rotor hard against mast.
b.
Inspect teflon trunnion bearings (non-elastomeric only) for squeaking,
binding, and/or ratcheting while flapping hub to each limit of travel.
c.
Reconnect pitch links IAW para 5-4b (13)(c).
ULTRASONIC INSPECTION OF MAIN ROTOR BLADES.
Perform ultrasonic shear wave and thru transmission inspection on main
rotor blades (P/N 540-011-001-5 and 540-011-250;1) at 550, 650, 750, 850,
950 and 1,050 total blade hours since new.
1-54 Change 71
PAGE NO.
9
NO. OF PAGES
24
DATE OF INSPECTION
STATUS
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
INSPECTION NO.
AIRCRAFT AND SERIAL NO.
AREA
NO.
REQUIREMENT
EVERY
TEST
6
a.
K747 Rotor Blade-Visually inspect the erosion guard edge seams
from station 75 to 260 (213 to 215 if appropriate). Inspect chordwise,
spanwise, top and bottom surfaces for separation or delamination.
7
b.
Remove transmission sump oil filter (wafer disk screen), and electrical
chip detector, check for contamination, then clean and reinstall.
7
c.
Intermediate (42') gearbox filler cap for clogged vent.
7
d.
Hydraulic fluid is to receive spectrometric analysis per TB 43-0106
every 25 hours.
11
e.
Tail rotor (90') gearbox filler cap for clogged vent.
11
f:
Tail rotor and elevator control tubes for wear at bulkhead grommets.
7
IMMEDIATELY PRECEDING 25 HOUR LUBRICATION AND
AT INSTALLATION
Swashplate and Support.
(1)
Visually inspect for evidence of contact between outer ring or drive
link and stationary swashplate. Measure vertical clearance from
the bottom of both drive links, PIN 209-010-408-7 to all three
horns of stationary swashplate. The minimum clearance must
be not less than .035 inch. Replace swashplate if any contact is
evident or if clearance is below minimum.
(2)
Disconnect main rotor drive links from swashplate and secure to
prevent damage. Rotate swashplate ring, checking for roughness,
binding, or unusual noise from swashplate bearing. Check for up
and down play in bearing between inner and outer swashplate.
Replace swashplate if any roughness, binding, noise or play is
evident where rotation of the main rotor blades is not possible.
This procedure can be accomplished as follows: disconnect the
pitch change links at the upper universal and remove the 6(each)
bolts securing the collet extension and the collet spline plate. This
will allow you to rotate the swashplate as required.
(3)
Using a soft, clean, lint free cloth dampened with cleaning solvent
(C124), clean the inner ring assembly and outer ring assembly at
the dust cover. Insure all surface grit, sand and other surface
materials are removed.
Change 71
1-55
NO. OF PAGES
24
DATE OF INSPECTION
STATUS
EVERY 25 HOURS
a.
PAGE NO.
10
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
INSPECTION NO.
AIRCRAFT AND SERIAL NO.
AREA
NO.
REQUIREMENT
EVERY
b.
TEST
(4)
Using a grease gun with a flexible hose, purge lubricate the
swashplate per figure 1-2. If the swashplate fails to accept grease,
perform a swashplate bearing sleeve alignment check. Also,
replace grease fittings if clogged.
(5)
Remove old grease purged from swashplate using a wooden tongue
depressor (NSN 6515-00-324-5500). Place sample of old grease in
plastic bottle (NSN 8125-01-082-9697) ensuring that the bottle is
more than half filled. The bottle label should be filled out as
completely as possible to avoid confusion with other grease
samples. Label to show operating activity, swashplate serial
number, aircraft serial number and date of sample.
(6)
Prepare DD Form 2026, Transit Aircraft Oil Analysis Record,
and submit along with grease sample to the AOAP Laboratory
designated in TB 43-0106. Make appropriate entries on DA Form
2408-20 in accordance with DA PAM 738-751.
(7)
Reconnect drive links and perform maintenance operational check
of at least 15 minutes duration.
PAGE NO.
NO. OF PAGES
DATE OF INSPECTION
STATUS
RECORDED
ON
WORKSHEET
Scissors and Sleeve Assembly.
(1)
Visually inspect visible part of scissors hub assembly and boat for
signs of heat. Any heat discoloration or distortion of components
is cause for replacement.
(2)
Using a soft, clean, lint-free cloth dampened with cleaning solvent
(C124), remove all surface grit, sand, and other surface materials.
(3)
Using a grease gun with a flexible hose, purge/lubricate the
scissors bearing in accordance with figure 1-2, at approximately
30-degree intervals until the assembly has been lubricated
through one full turn (360').
1-56 Change 71
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
INSPECTION NO.
AIRCRAFT AND SERIAL NO.
AREA
NO.
REQUIREMENT
EVERY
TEST
Remove old grease purged using a wooden tongue depressor
(NSN 6515-00-324-5500). Place sample of old grease in plastic
bottle (NSN 8125-01-082-9697), ensuring that the bottle is
more than half filled. The bottle label should be filled out as
completely as possible to avoid confusion with other grease
samples. Label to show operating activity, scissors and
sleeve assembly serial number, aircraft serial number, and
date of sample.
(5)
Prepare DD Form 2026, Transit Aircraft Oil Analysis
Record, and submit along with grease sample to the AOAP
Laboratory designated in TB 43-0106. Make appropriate entries on DA Form 2408-20 in accordance with DA Pam
738-751.
1-56A / (1-56B blank)
NO. OF PAGES
DATE OF INSPECTION
STATUS
(4)
Change 50
PAGE NO.
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
INSPECTION NO.
AIRCRAFT AND SERIAL NO.
AREA
NO.
REQUIREMENT
EVERY
PAGE NO.
12
TEST
An engine hot-end inspection, in accordance with TM 55-2840-229-23 (T53 Engine Inspection Guide), is required when exhaust gas temperature limits have been exceeded. During transient for starting and acceleration when turbine gas temperature (TGT)
exceeds 950 degrees C at any time or when TGT exceeds 880 degrees C for more than
5 seconds, a hot-end inspection must be performed. Refer to TM 55-2840-229-23.
NOTE
If engine cannot be operated without exceeding TGT limits (of 950 degrees
C TGT at any time) as specified in TM 55-2840-229-23, Engine Operating
Limits Table, this is indication of engine malfunction or instrument error.
Refer to Troubleshooting Tables, to determine cause and correct action
as overtemperature inspection is not required.
WHEN ENGINE OIL TEMPERATURE LIMITS ARE EXCEEDED
Refer to TM 55-2840-229-23.
Change 42
DATE OF INSPECTION
STATUS
AFTER ENGINE OVERTEMPERATURE
1-57
NO. OF PAGES
24
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
INSPECTION NO.
AIRCRAFT AND SERIAL NO.
AREA
NO.
10
REQUIREMENT
EVERY
TEST
AFTER ENGINE OVERSPEED
An engine overspeed exists under the following conditions:
a.
When N1 speed exceeds 106 percent.
b.
When steady-state output shaft speed exceeds:
(1)
6900 rpm as a maximum 10 second limit.
(2)
6700 rpm when operating above 750°C TGT.
NOTE
At TGT of 750°C or less, a steady output shaft
speed of 6900 rpm is permissible.
If overspeed limits are exceeded, perform engine overspeed inspection
in accordance with TM 55-2840-229-23.
1-58 Change 38
PAGE NO.
13
NO. OF PAGES
24
DATE OF INSPECTION
STATUS
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
AIRCRAFT AND SERIAL NO.
AREA
NO.
INSPECTION NO.
REQUIREMENT
EVERY
TEST
10
AFTER HELICOPT'ER IS FLOWN IN A LOOSE GRASS ENVIRONMENT.
Any time the helicopter is flown in a loose grass environment, the
engine shall be inspected for grass blockage in accordance with TM
55-2840-229-23.
10
ENGINE POST-INSTALLATION INSPECTION
a.
Check installation of power control linkage in accordance with
Chapter 4, Section VII.
Change 29
1-59
PAGE NO.
14
NO. OF PAGES
24
DATE OF INSPECTION
STATUS
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
AIRCRAFT AND SERIAL NO.
AREA
NO.
REQUIREMENT
EVERY
INSPECTION NO.
TEST
PAGE NO.
15
DATE OF INSPECTION
STATUS
b.
Perform turbine gas temperature (TGT) System test. (Refer to TM 552840-229-23 and TM 55-4920-244-14.)
c.
Perform a Daily Inspection. (Refer to TM 551500-220PMD.)
NOTE
The following step d. need not necessarily be performed if the
engine has merely been removed and reinstalled for reasons
other than engine maintenance. However, the engine should be
inspected for leaks and security of mounting provisions, hoses
and accessories prior to flight.
d.
Perform inspection before and after initial check run. (Refer to TM 552840-229-23.)
e.
Perform a limited test flight. (Refer to TM 55-1520-234 MTF.)
f.
Perform an engine vibration test. (Refer to TM 55-2840-229-23.)
1-60 Change 54
NO. OF PAGES
24
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
AIRCRAFT AND SERIAL NO.
AREA
NO.
10
REQUIREMENT
EVERY
INSPECTION NO.
TEST
If an engine is dropped during handling, perform inspection.
(Refer to TM 55-2840-229-23).
AFTER EXCESSIVE ENGINE TORQUE
Overtorque is defined as any incident in which torsional loads are
introduced into the helicopter dynamic system in excess of 56 psi
as determined on the engine torquemeter (calibrated).
NOTE
Use calibrated torque for overtorque limits. The following table
shall be used to convert indicated torque to calibrated torque.
Calibration Factor Multiply Indicated Torque By
64
0.96
63
0.97
62
0.99
61
1.00
60
1.02
59
1.04
58
1.06
Example: For an indicated torque of 55 psi with a calibration factor of 58, the calibrated torque is 58.3 psi (55 x 1.06 = 58.3).
OVERTORQUE IN EXCESS OF 56 PSI BUT NOT EXCEEDING 58 PSI
Perform thorough visual inspection of components. If inspection reveals
no discrepancies, or damage to the following components, they may be
retained in service.
(1)
Main Rotor Blades
(2)
Main Rotor Hub
(3)
Tail Rotor Blades
Change 38
1-61
NO. OF PAGES
DATE OF INSPECTION
STATUS
ENGINES DROPPED DURING HANDLING
a.
PAGE NO.
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
AIRCRAFT AND SERIAL NO.
AREA
NO.
REQUIREMENT
EVERY
INSPECTION NO.
TEST
(4)
Tail Rotor Hub
(5)
42°Gearbox
(6)
90°Gearbox
(7)
Tail Rotor Driveshafts
(8)
Swashplate
(9)
Scissors and sleeve assembly
(10) Input driveshaft
(11) Mast
(12) Transmissions (all part numbers)
(13) Driveshaft Hanger assemblies
AFTER ENGINE TORQUE IN EXCESS OF 58 PSI
BUT NOT EXCEEDING 62 PSI
Inspect and/or replace components as follows. Records of replaced
components shall show over-torque condition as reason for removal.
a.
Replace main rotor trunnion bearing bolts.
.
b. Inspect main transmission sump filter and chip detector.
c.
(1)
If metal particles are found indicating internal failure, replace
transmission, and send it to overhaul for evaluation.
(2)
If there are no positive indications of failure continue operation
for 5 hours, then repeat inspection. If no indications of failure
are then found, resume normal operation.
Perform thorough visual inspection of the following components,
each may be kept in service if no discrepancy or obvious damage is
found. Replace any damaged component.
(1) Main Rotor Blades
(2) Main Rotor Hub
1-62
PAGE NO.
NO. OF PAGES
DATE OF INSPECTION
STATUS
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
INSPECTION NO.
AIRCRAFT AND SERIAL NO.
AREA
NO.
REQUIREMENT
EVERY
TEST
(3)
Tail Rotor Blades
(4)
Tail Rotor Hub
(5)
42°Gearbox
(6)
90°Gearbox
(7)
Tail Rotor Driveshafts
(8)
Tail Rotor Driveshaft Hangers
(9)
Swashplate Assembly
(10) Scissors and Sleeve Assembly
(11) Input Driveshaft
(12) Mast
AFTER ENGINE TORQUE IN EXCESS OF 62 PSI
a.
Replace the following components, and send to overhaul for
evaluation, with records showing over-torque as reason for
removal.
(1)
Transmission
(2)
Main (input) Driveshaft
(3)
Mast
(4)
Main Rotor Blades
(5) Main Rotor Hub
(6) Tail Rotor Blades
b.
Perform thorough inspection of the following components, each
may be kept in service if no discrepancy or obvious damage is
found.
(1)
Deleted
(2)
Tail Rotor Hub
(3)
42°Gearbox
Change 11
1-63
PAGE NO.
18
NO. OF PAGES
24
DATE OF INSPECTION
STATUS
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
AIRCRAFT AND SERIAL NO.
AREA
NO.
INSPECTION NO.
REQUIREMENT
EVERY
TEST
(4)
90°Gearbox
(5)
Tail Rotor Driveshafts
(6)
Driveshaft Hangers
(8) Scissors and Sleeve Assembly
ENGINE OVERTORQUE INSPECTION REQUIREMENTS:
When the engine has exceeded overtorque limits, perform engine
overtorque inspection in accordance with TM 55-2840-229-23.
The following limits are engine torque limits only. Pilot
monitoring is necessary to prevent the engine from
exceeding dynamic components (airframe) limits.
a.
Output shaft torque emits:
(1)
Military (30 minutes)
64 psi
(2)
Normal (continuous)
60 psi
(3)
Transient operation (2
seconds or less)
86 psi
AFTER TAIL ROTOR DRIVE SYSTEM OVER-TORQUE.
After any report of suspected over-torquing of the tail rotor drive
system during operation, preform the following inspection as soon as
possible after the incident.
1-64 Change 29
NO. OF PAGES
24
DATE OF INSPECTION
STATUS
(7) Swashplate Assembly
7
PAGE NO.
19
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
AIRCRAFT AND SERIAL NO.
AREA
NO.
REQUIREMENT
EVERY
INSPECTION NO.
TEST
a.
Remove output quill assembly from 42°gearbox, and inspect
output gear teeth for damage as described in Chapter 6. If no
scoring or scuffing is found, and if there are no other indications
of damage, reassemble gearbox in accordance with maintenance
manual, and retain in service. If gear teeth are scored or scuffed,
or if there are other indications of damage, replace gearbox and
perform inspection in step b.
b.
Remove output quill assembly from 900 gearbox, and inspect
condition of gears as in step a. If no scoring or scuffing is found,
and if there are no other indications of damage, reassemble
gearbox in accordance with maintenance manual, and retain in
service. If gear teeth are scored or scuffed, or if there are other
indications of damage, replace gearbox and perform inspection in
step c.
c.
Remove transmission tail rotor drive quill and inspect condition of
gear teeth. Evidence of scoring or scuffing is cause for replacement
of main transmission assembly. If it is necessary to replace the
transmission assembly, then the tail rotor hanger bearing
assemblies and tail rotor driveshafts must also be replaced.
d.
Tag any removed components with reason for removal before
turning in through normal supply channels for overhaul.
3
2
a.
Inspect wire strike deflector on TSU windows for bends, cracks, and
alignment. Replace if damage is found.
b.
Inspect nose wire strike cutter for bends, cracks, and alignment. Replace
if damage is found.
c.
Inspect channels, inserts, and nose deflectors for bends, cracks, and
alignment. Replace if damage is found.
d.
Inspect upper wire strike cutter for bends, cracks, and alignment.
Inspect panel for cracks and pulled inserts. Replace damaged parts.
e.
Inspect lower wire strike cutter for bends, cracks, and alignment.
Replace if damage is found.
Change 44
1-64A / (1-64B blank)
NO. OF PAGES
DATE OF INSPECTION
STATUS
AFTER KNOWN OR PROBABLE WIRE STRIKE
1
PAGE NO.
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
AIRCRAFT AND SERIAL NO.
AREA
NO.
7
REQUIREMENT
EVERY
INSPECTION NO.
TEST
a.
Visually inspect the mast area where the hub stop would contact the
mast. If no surface deformation of the mast has occurred, the inspection
is complete.
b.
If there is visual evidence of surface deformation of the mast due
to hub stop contact:
(1)
Evaluate the condition of the mast per the damage limits in
Chapter 6.
(2)
Inspect and replace the following items if damage is found:
(a) Main rotor hub trunnion cap attach bolts and drag
brace jamnuts and attach bolts for security.
1-65
NO. OF PAGES
24
DATE OF INSPECTION
STATUS
MAST BUMPING
Change 44
PAGE NO.
20
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
AIRCRAFT AND SERIAL NO.
AREA
NO.
INSPECTION NO.
REQUIREMENT
EVERY
TEST
(b) Flight control system, from rotor to servo cylinder, for
bent or damaged tubes and rod end bearings.
(c) Structure at transmission mounting points.
(d) Lift link and structure for damage, security and
distortion.
(e) Transmission sump oil filter, external oil filter, and
chip detector for metal particles.
(f) Main Drive shaft.
(g) Tail rotor driveshafts and hanger assemblies for
obvious damage.
(h) Tail rotor drive quill.
7
7
(i)
Canopy of aircraft.
(j)
Tail rotor driveshaft cover.
AFTER TRANSMISSION OIL OVER TEMP
a.
Troubleshoot transmission oil system to determine cause. (Refer to
Chapter 6.)
b.
Replace transmission, mast, oil cooler and external oil filter if cause
is due to transmission internal failure. If cause is due to oil system
external to transmission and oil temperature did not exceed 130°C
for 15 minutes, correct cause of overheating and drain and refill
transmission oil system.
c.
If temperature exceeded above limits, replace transmission and
mast. If abnormal contamination is present, also replace oil cooler
and external oil filter.
AFTER COMPLETE LOSS OF TRANSMISSION OIL
a.
Troubleshoot transmission oil system to determine cause.
b.
Replace transmission and mast, if engine power was applied after
complete loss of oil. Also replace oil cooler and external oil filter
if abnormal contamination is present.
1-66 Change 27
PAGE NO.
21
NO. OF PAGES
24
DATE OF INSPECTION
STATUS
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
INSPECTION NO.
AIRCRAFT AND SERIAL NO.
AREA
NO.
8
REQUIREMENT
EVERY
TEST
AFTER FIRTING EJECTOR CARTRIDGES. Refer to Chapter 16.
1
a.
Inspect wire strike deflector on TSU windows for bends,
cracks, and alignment. Replace if damage is found.
2
b.
Inspect nose wire strike cutter for bends, cracks, and
alignment. Replace if damage is found.
2
c.
Inspect lower wire strike cutter for bends, cracks, and
alignment.
9
AFTER OVERFLOW OF BATTERY
a.
Refer to TM 55-1500-204-25/1 for treatment of affected areas.
b.
Sheet metal surfaces and overlaps, both internal and external,
for damage.
c.
Rivets, bolts, screws, and other hardware for damage.
d.
Troubleshoot battery system to determine cause.
EVERY 30 DAYS OR 25 FLIGHT HOURS, WHICHEVER
OCCURS FIRST
Perform preventive maintenance checks and services on nickel-cadmium
battery. (Refer to TM 11-6140-203-14-2).
9
EVERY 100 HOURS OR 120 CALENDAR DAYS, WHICHEVER
OCCURS FIRST
a.
Perform preventive maintenance checks and services on
nickel-cadmium battery. (Refer to TM 11-6140-203-14-2.)
b.
Check voltage regulator setting; adjust for temperature as
required. TM 55-1500-204-25/1.
Change 71
1-66A
NO. OF PAGES
24
DATE OF INSPECTION
STATUS
AFTER KNOWN OR PROBABLE WIRE STRIKE
9
PAGE NO.
21A
RECORDED
ON
WORKSHEET
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
INSPECTION NO.
AIRCRAFT AND SERIAL NO.
AREA
NO.
3
REQUIREMENT
EVERY
Pilot Static System-Perform a functional check of system: refer to
paragraph 8-20A.
NOTE
Compute calendar time from date stamped on instrument case
of altimeter and airspeed indicator.
b.
Encoding Altimeter - After the 24 month test or replacement of
the altimeter, check mode C (altitude) test using the appropriate
transponder test procedure in TM 11-6625-667-12 or TM 11-4920296-14&P.
1-66B
Change 65
NO. OF PAGES
24
DATE OF INSPECTION
STATUS
EVERY 24 MONTHS
a.
PAGE NO.
21B
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
INSPECTION NO.
AIRCRAFT AND SERIAL NO.
AREA
NO.
3
PAGE NO.
22
REQUIREMENT
EVERY
Inspect and test OAT/FAT gage in accordance with TM 55-1500-204-25/1.
3
6
EVERY 12 MONTHS
a.
Magnetic compass for discoloration of liquid and proper calibration;
recompensate if necessary.
b.
Gyromagnetic compass system for proper calibration; recompensate
if necessary. (Refer to TM 55-1500-204-25/1.)
c.
Replace ejector rack cartridges (refer to Chapter 16).
d.
Deleted.
DAILY WHEN OPERATING IN HIGH HUMIDITY OR SALT-LADEN AIR
Wash main and tail rotor blades with mild soap (C125), rinse with clear
water, and dry.
6
30 DAYS OR 50 HOURS OF OPERATION (WHICHEVER IS FIRST)
Wash main and tail rotor blades with mild soap (C125), rinse with clear
water, and dry.
All
Areas
3
DATE OF INSPECTION
STATUS
EVERY 12 MONTHS OR NEAREST SCHEDULED PHASE INSPECTION
AFTER THE HELICOPTER HAS BEEN SUBJECTED TO SALT
WATER OR SALT WATER SPRAY
a.
Wash entire helicopter with fresh water, particularly inside of engine
compartment doors; wash all compartments which were exposed to
salt water; make a detail check of all surfaces for corrosion. Apply
corrosion preventive compound to exposed non-painted, anodized,
or cadmium plated assemblies. Water-wash engine internally.
b.
lean water wash engine daily when operated within 10 miles of salt
water or within 200 miles of volcanic activity. (IAW TM 55-2840229-23)
AFTER WASHING HELICOPTER
Check pitot-static system for moisture.
Change 71
1-67
NO. OF PAGES
24
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
AIRCRAFT INSPECTION CHECKSHEET
AIRCRAFT AND SERIAL NO.
AREA
NO.
3
All
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
PAGE NO.
23
INSPECTION NO.
DATE OF INSPECTION
REQUIREMENT
EVERY
STATUS
EVERY 6 MONTHS
a.
Weight check CF3BR fire extinguisher. Refer to TM 55-1500-204-25/1
b.
Inspect and test connector receptacle (ground). Refer to TM 551500-323-25.
AFTER PROBABLE EXPOSURE TO RADIOACTIVITY
Accomplish the following:
All
a.
Survey helicopter for level of radioactivity.
b.
Decontaminate helicopter as required. (Refer to TM 3-220.)
WHEN HELICOPTER IS TRANSFERRED,
STORAGE, OR REMOVED FROM STORAGE
RECEIVED,
PLACED
IN
Inventory helicopter for availability of inventoriable property. (Refer to
DA PAM 738-751.)
All
AFTER INSTALLATION, REMOVAL OR RELOCATION OF EQUIPMENT
OR MAJOR MODIFICATION WHICH RESULTS IN UNKNOWN
CHANGE IN BASIC WEIGHT AND BALANCE; AFTER REPORT OF
UNSATISFACTORY FLIGHT CHARACTERISTICS
Weigh helicopter and accomplish necessary entries in Weight and
Balance Data, DD Forms 365. (Refer to AR95-3 and TM 55-1500342-23.)
9
AFTER INDICATION OF UNUSUAL OVER OR UNDER FUEL
CAPACITY OR UNUSUAL OVER OR UNDER FUEL CONSUMPTION
Perform inspection of fuel cells with special attention to possible self sealing
material and/or inner liner separation.
All
UPON TRANSFER AND UPON RECEIPT OF A HELICOPTER; UPON
EXPIRATION OF 12 MONTHS ELAPSED TIME SINCE LAST INVENTORY,
UPON PLACING A HELICOPTER IN STORAGE AND UPON REMOVAL
FROM STORAGE (HELICOPTER NEED NOT BE INVENTORIED WHILE IN STORAGE)
Perform an inventory check. (Refer to DA Form 2408-17 and Appendix C.)
Change 65
1-68
NO. OF PAGES
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RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
INSPECTION NO.
AIRCRAFT AND SERIAL NO.
AREA
NO.
All
PAGE NO.
24
REQUIREMENT
EVERY
For general weight and balance information, refer to TM 55-1500-342-23,
Army Aviation Maintenance Engineering Manual, Weight and Balance.
Appendix B, Maintenance Allocation Chart should be consulted for
responsibility of weighing and balancing of the aircraft.
11
AFTER M65 TOW MISSILE HAS BEEN FIRED
Inspect tailboom skins for nicks, dents, scratches, creases and cracks.
11
AFTER FIRING 2.75 INCH ROCKETS
Visually inspect tailboom skins, synchronized elevators, and tail rotor
blades for damage.
7&
11
EVERY 12 MONTHS OR 600 HOURS (WHICHEVER IS FIRST)
Accomplish the following:
7
a.
Lubricate tail rotor drive train flexible couplings.
b.
Visually check splines for wear and nicks.
c.
Visually check flexible coupling seal for proper installation, cuts, and tears.
d.
Inspect hanger bearings for evidence of grease leakage, corrosion,
overheating (brown discoloration of green zinc chromate paint on
hanger), and notchiness.
e.
Remove, disassemble, clean, inspect, lubricate, assemble, and reinstall
main drive shaft assembly.
f.
Inspect main driveshaft for internal corrosion.
EVERY 18 MONTHS
Transmission external oil filter replaced.
All
DATE OF INSPECTION
STATUS
WHEN OVERHAULS, MAJOR MODIFICATIONS OR MAJOR AIRFRAME
REPAIRS ARE ACCOMPLISHED, ANY SPECIAL EQUIPMENT HAS BEEN
ADDED TO OR REMOVED FROM THE BASIC AIRFRAME OR WHEN
WEIGHT AND BALANCE DATA ARE SUSPECTED TO BE IN ERROR.
AFTER 7 DAYS
After the aircraft has remained inactive for 7 consecutive days, process
the aircraft into the appropriate storage category. (See Appendix E.)
Change 63
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NO. OF PAGES
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TM 55-1520-234-23
AIRCRAFT INSPECTION CHECKSHEET
AIRCRAFT AND SERIAL NO.
AREA
NO.
All
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
PAGE NO.
24A
INSPECTION NO.
REQUIREMENT
EVERY
a.
DATE OF INSPECTION
STATUS
AFTER LIGHTNING STRIKE
General Requirements:
Whenever the aircraft is struck by lightning:
(1)
Inspect the fuselage interior and exterior, the landing gear, the
rotor systems and static ground wire for burn marks, cracks,
pitting or other signs of high temperature stress, to determine
the lightning entry and exit points.
(2)
Trace the path of the lightning strike to the extent possible using
a magnetometer.
(3)
Check the magnetic compass for accuracy (the degree of
inaccuracy may serve as an indicator of the severity of the
strike).
(4) Inspect wiring in tunnel areas and exposed areas for burns.
(5) Inspect antennas for burns and pitting.
(6)
Inspect all electrically operated components and lightning
systems for damage.
(7)
Inspect communications and navigation equipment for damage.
(8)
If the preceding steps reveal major damage has occurred,
proceed as follows:
(a) Bench test all avionics and electrical systems and
components.
If damage to fuel probes is suspected, probes shall be removed
from helicopter and continuity check performed on bench. Fuel
quantity Indicator wiring shall be checked with fuel probes
removed, or with wiring disconnected from probes.
(b) Perform a Megger check and continuity check on all wiring
and cables.
Change 2
1-70
NO. OF PAGES
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RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
INSPECTION NO.
AIRCRAFT AND SERIAL NO.
AREA
NO.
PAGE NO.
24B
REQUIREMENT
EVERY
(c) Perform a Voltage Standing Wave Ratio (VSWR) check on
all antennas, antenna cables, and connectors.
(9) Perform specific inspections/replacements as required.
(10) Perform a ground run operational check on the helicopter.
Functionally check the flight control system, and all avionics,
electrical, lighting, communication, and navigation systems.
(11) Repair any damage and replace damaged components as
required, using standard maintenance practices.
b.
DATE OF INSPECTION
STATUS
Do not use VSWR on fuel probes installed in helicopter.
Specific Requirements:
(1) Whenever lightning strike is evident on main rotor system:
NOTE
If there is any evidence of lightning strike on blades or any other
parts, scrap parts locally. In case of doubt, proceed as outlined
below.
(a) Inspect blades for damage such as burns, pitting, skin
separation, etc. If any damage is evident, locally scrap
damaged blade(s).
(6) Remove hub assembly and return for overhaul.
(c) Replace all bearings (or next higher assembly if required)
in the fixed and rotating control system located above the
servo cylinders.
(d) Remove transmission assembly, and return for overhaul, if
required.
(e) Check main rotor driveshaft for residual magnetism. If
magnetized, remove, and visually inspect short shaft for
damage and remove engine and return for overhaul.
(f) Inspect swashplate assembly and mast assembly. Return
for overhaul if required.
Change 65
1-70A
NO. OF PAGES
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RECORDED
ON
WORKSHEET
TM 55-1520-234-23
AIRCRAFT INSPECTION CHECKSHEET
AIRCRAFT AND SERIAL NO.
AREA
NO.
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
PAGE NO.
24C
INSPECTION NO.
REQUIREMENT
EVERY
DATE OF INSPECTION
STATUS
(2) Whenever lightning strike is evident on tail rotor system:
NOTE
If there is any evidence of lightning strike on blades or any other
parts, scrap parts locally. In case of doubt, proceed as outlined
below.
(a) Inspect blades for damage such as burns, pitting, skin
separation, etc.
If damage is evident, locally scrap
damaged blade(s).
(b) Tail rotor hub: Scrap locally.
(c) Remove and condemn pitch change links, crosshead
bearing and control quill bearings.
(d) Inspect crosshead, control quill components, and control
rod for any indications of arcing. Replace as necessary.
(e) Remove main transmission and both the 42 degree and 90
degree gearboxes and return them for overhaul.
(f) Remove hanger assemblies and return for overhaul.
(g) Inspect tail rotor driveshafts for evidence of arcing, burns,
or other related damage. If damage is evident, locally
scrap damaged shafts.
(h) Remove, disassemble, and inspect main rotor driveshaft
for damage. If damage is evident or if residual magnetism
is found in couplings or shaft, remove engine and return for
overhaul.
Change 38
1-70B
NO. OF PAGES
24
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
AIRCRAFT INSPECTION CHECKSHEET
AIRCRAFT AND SERIAL NO.
AREA
NO.
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
PAGE NO.
24D
INSPECTION NO.
REQUIREMENT
EVERY
5
c.
Nonstandard locally manufactured heavy duty skid shoes exceeding
20 pounds each in weight, Inspect as follows:
(1)
Perform inspection using liquid fluorescent dye penetrant
method. This inspection is to be conducted by AVUM with
assistance from AVIM as required to gain access to inspection
area.
(a) Jack airframe and remove landing gear assembly in
accordance with specified procedures.
(b) Conduct visual inspection for nicks, scratches, or gouges
over entire cross tube surface. Refer to Chapter 3 for
allowable damage criteria.
(c) Remove blind rivets securing cross tube/fuselage
attachment fittings to cross tubes and remove fittings.
Save fittings for reuse.
(d) Fluorescent dye penetrant inspection is required on cross
tube surface of an area completely around cross tube at
support fitting locations and adjacent area one inch out
from each end of support fittings. A one inch band around
cross tubes at chem-milled step area adjacent to skid tube
saddle fitting will also be dye penetrant inspected.
(e) Prepare the surface and conduct a liquid fluorescent dye
penetrant inspection in accordance with TM 43-0103,
utilizing penetrant kit NSN 6850-00-782-2740.
(f) Wash and remove all excess penetrant and developer from
cross tube surfaces.
{g) Recoat surface of cross tube to be covered by the
attachment fittings with sealant (C116).
(h) Reinstall attachment fittings to cross tubes. Diameter and
length of rivets to be determined at time of installation.
(2)
DATE OF INSPECTION
STATUS
EVERY 50 HOURS
Cross tubes with crack indications will be scrapped. All cross tubes
with no crack indications are to be returned to serf Ice.
Change 65
1-70C
NO. OF PAGES
24
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
INSPECTION NO.
AIRCRAFT AND SERIAL NO.
AREA
NO.
5
PAGE NO.
24E
DATE OF INSPECTION
REQUIREMENT
EVERY
STATUS
EVERY 150 HOURS
Nonstandard locally manufactured heavy duty skid shoes weighing 20
pounds or less each, inspect same as 50 Hours inspection on same item
of more weight.
7
EVERY 300 HOURS
Remove, disassemble, clean and inspect main rotor hub assembly. (See
Chapter 5.)
NOTE
When main rotor hub assembly is removed from one aircraft
and installed on another, assure that the component's next
inspection due time is transferred to the receiving aircraft's DA
Form 2408-18.
10
EVERY 900 HOURS OR EACH TIME ENGINE IS REMOVED AND REPLACED
a.
Engine oil system, including oil cooler, drained and refilled. If
engine is replaced due to internal failure, flush all airframemounted oil lines and engine oil tank and replace engine oil cooler.
NOTE
Only to aircraft engine serviced with MIL-L-23699.
7808 is used, oil change is required every 300 hours.
b.
If MIL-L-
If engine is changed check gyromagnetic compass for proper
calibration; recompensate if necessary (TM 55-1500-204-25/1).
Change 71
1-70D
NO. OF PAGES
24
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
TYPE OF INSP (Daily,
Intermediate, etc.)
SPECIAL
AIRCRAFT INSPECTION CHECKSHEET
INSPECTION NO.
AIRCRAFT AND SERIAL NO.
AREA
NO.
7
PAGE NO.
24F
REQUIREMENT
EVERY
Transmission and transmission oil cooler drained and oil pump screen
inspected for metallic particles and other contaminants. Clean screen,
replace and refill transmission oil system to proper level. If transmission is
replaced due to internal failure, flush all airframe-mounted transmission oil
lines and replace transmission oil cooler.
NOTE
Only to aircraft transmission serviced with MIL-L-23699. If MILL-7808 is used, oil change is required every 300 hours.
9
EVERY 1100 HOURS OR AT OVERHAUL (WHICHEVER COMES FIRST)
Inspect diagonal brace fitting P/N 209-030-183-1 for cracks using
fluorescent penetrant method.
7
SWASHPLATE BEARING SLEEVE ALIGNMENT CHECK
Perform the following whenever the swashplate fails to accept grease or
upon any indication of linear misalignment (overheating; difficulty in
lubing).
a. Using pliers, remove lubrication fitting (NAS516) from one
side of swashplate.
b. Insert alignment tool (T65) into lubrication port. Any
stoppage before tool is fully seated is cause for rejecting
the swashplate.
c.
Lightly tap a new fitting (NAS516) into place.
d. Lubricate swashplate and support in accordance with figure
1-2.
3
DATE OF INSPECTION
STATUS
EVERY 900 HOURS OR EACH TIME TRANSMISSION IS REMOVED AND
REPLACED
FIRST AID KIT INSPECTION
Inspect per TM 55-1500-204-25/1.
Change 65
1-70E/(1-70F blank)
NO. OF PAGES
24
RECORDED
ON
WORKSHEET
TM 55-1520-234-23
Section V. OVERHAUL AND RETIREMENT SCHEDULE
1-32. Overhaul and Retirement Schedule.
WARNING
TM 55-1500-328-25 should be referred to
concerning mutilation destruction of items
when they have reached the established
life expectancy (finite life) before the items
are forwarded for property disposal.
AREA
FIGURE 1-9
This section lists units of operating equipment that are
to be overhauled or retired at the period specified.
Removal of equipment for overhaul may be accomplished
at the inspection nearest the time when overhaul is due
unless otherwise specified in TM 55-1500-328-25.
OVERHAUL
INTERVAL
PART NUMBER AND ITEM
T53-L-703
Engine
RETIREMENT
SCHEDULE
2400
Canopy
3
3
3
3
209-030-711-11
209-030-711-13
209-030-711-37
6106319
(821798-1)
Crew Compartment
Interconnect Line (Detonating Cord)
Interconnect Line (Detonating Cord)
Arming/Firing Mechanism
Arming/Firing Mechanism
3
3
3
3
54H1956
62B4407
57D677
0101949-1
Seat Belt
Seat Belt
Shoulder Harness
Shoulder Harness
72 Months****
72 Months****
72 Months***
72 Months****
Refer to TM 55-1500-204-25/1
Main Rotor
6
6
6
6
6
6
6
6
6
6
6
209-010-518-101
209-010-520-101
540-015-001-1
540-011-101-5
540-011-101-17
540-011-101-25
540-011-101-131
540-011-153-13
540-011-153-15
540-011-153-17
540-011-102-5
Pitch Link Assembly
Pitch Link Assembly
Main Rotor Blade Assembly
Main Rotor Hub Assembly
Main Rotor Hub Assembly
Main Rotor Hub Assembly
Main Rotor Hub Assembly
Main Rotor Hub Extension Assembly
Main Rotor Hub Extension Assembly
Main Rotor Hub Extension Assembly
Main Rotor Hub Yoke Assembly
6
6
6
6
6
6
6
6
6
6
6
6
6
6
6
6
6
6
6
204-012-112-7
204-012-122-3
204-012-122-7
540-011-101-129
540-011-101-131
540-011-112-1
540-011-112-5
540-011-112-3
540-011-112-7
K747-003-205
K747-003-205
K747-003-209
K747-003-303
K747-003-303
K747-003-309
K747-003-401
K747-003-403
K747-072-1
K747-061-5
Main Rotor Hub Retention Straps
Retention Straps
Retention Straps
Main Rotor Hub
Main Rotor Hub
Pin, Inboard
Pin, Inboard
Pin, Outboard
Pin, Outboard
Main Rotor Blade
Deviation Main Rotor Blade
Main Rotor Blade
Main Rotor Blade
Field Modified Main Rotor Blade
Main Rotor Blade
Main Rotor Blade
Main Rotor Blade
Fitting Assembly, Drag Strut
Fitting Assembly, Root
Change 71
1-71
5700
5500
1100
1200
1200
1200
1200
2200
2200
2200
4800
1200
1200
1200
1200
1200
2400
2400
2400
2400
10,000
10,000
10,000
10,000
10,000
10,000
10,000
10,000
1400
1000
TM 55-1520-234-23
AREA
FIGURE 1-9
6
6
6
6
6
6
6
6
6
6
6
6
K747-082-1
K747-083-1
NAS6618D112
NAS6210-38D
NAS464P8-36
NAS6206-76D
NAS6209-103D
540-011-113-1
540-011-177-1
540-011-154-5
209-010-109-5, -109
204-310-101-103
7
7
7
7
7
7
7
7
7
7
7
7
7
7
7
7
Mast Controls
209-010-460-3
209-010-400-1
209-010-403-1
209-010-402-1
204-076-317-1,-5, -7
100328
209-010-401-11, -5
209-010-405-7
204-076-428-1, -3, -5
41000434
209-076-124-1, -3, -5
209-001-358-13, -17, -21
209-010-408-7
209-010-407-1
209-010-450-5
540-011-456-1
Drive Link, Swashplate
Hub Assy.
Shaft, Mast Assy.
Sleeve, Collective
7
7
7
7
7
7
7
7
7
7
7
7
7
7
7
7
7
Transmission
209-011-208-101
209-011-212-101
209-011-219-101
209-310-100-105
205-040-263-105
209-961-410-1
209-961-400-3
212-040-001-39
205-040-263-3, -111
212-040-365-33
212-040-365-25
209-040-366-3
204-040-136-7
204-040-136-9
212-040-136-1
212-030-104-5
SKCP 2381-1
Plate
Support Set
Strap Assembly
Spring
Quill Assembly
Hydraulic Pump Quill
Transmission Assembly
Transmission Assembly
Main Input Quill
Hydraulic Pump Quill
Tail Rotor Output Quill
Mast Assembly
Mast Bearing
Mast Bearing
Mast Bearing
Lift Link
KFLEX, SHAFT
8
8
OVERHAUL
INTERVAL
PART NUMBER AND ITEM
Ejector Racks
ARD863-1
P7911-2
CCU-44/B
5184850
Drag Strut
Root Fitting
Bolt
Bolt
Fitting Bolt, Drag Strut
Bolt
Bolt
Fitting, Strap
Nut, Strap Fitting
Grip Assembly, Main
Pitch Horn Assy.
Retention Strap
RETIREMENT
SCHEDULE
4000
4000
4000
4000
450
4000
4000
2400
2400
5500
6600
1200
Pitch Control Tube
Swashplate & Support Assembly
Outer Ring
Inner Ring
Bearing Housing
Cylinder Barrel
Scissors & Sleeve Assembly
Scissors Lever
Rod End Bearing
Housing Assembly, Servo
Extension Tube
600
1200
3300
3300
3300
3300
1200
Extension Tube
3600
600
3300
3300
3300
11000
11000
3300
11000
5700
5700
5700
5700
1500
1200
1200
1200
1500
1200
1200
1200
1500
1500
1500
5700
1500
Cartridges
15 months**
Cartridges
36 months*
9
9
9
9
Oil Cooler Turbine
J33C32
H33C32
P9103NPPFS50160
P9101NPPFS50160
Bearing, Oil Cooler
Bearing, Oil Cooler
Bearing, Oil Cooler
Bearing, Oil Cooler
450
450
450
450
11
11
11
11
11
11
Tail Rotor and Drive System
204-040-623-5
212-010-744-5
212-010-704-5
212-010-750-11
212-040-003-13-23
212-040-004-9
Tail Rotor Driveshaft Hanger Bearing
Yoke Assembly
Yoke Assembly
Blade Assembly
42 Degree Gearbox
90 Degree Gearbox
600******
2400
2400
2400
Change 71
1-72
1200
1200
TM 55-1520-234-23
*
Not to exceed 8.5 years from date of manufacture (shelf life), or 36 months from the date of opening the sealed
cartridge container (installed life) or 10 installations and or removals from cartridge housing. Explosive life is
not additive and therefore cartridge removal is required whenever any of these conditions are reached.
**
Not to exceed 10 years from date of manufacture (shelf life) or 15 months from date of opening the sealed
cartridge container (installed life) or ten installations and or removals from cartridge housing. Explosive life is
not additive and therefore cartridge removal is required whenever any of these conditions are reached.
***
Denotes life is not to exceed 9 years from date of manufacture (shelf life) or 72 months from the date of
installation, whichever comes first.
****
Denotes life is not to exceed 8 years from the date of manufacture (shelf life) or 72 months from the date of
installation, whichever occurs first.
*****
Denotes life is not to exceed 9 years from the date of manufacture (shelf life) or 72 months from the date of
installation, whichever occurs first.
******
Denotes bearing has a shelf life of 5 years, effective 31 December 1990.
NOTE
All unserviceable explosive items (NSN 1377 Class) shall be tagged with
NSN, installation date, removal date, reason for removal, lot number
helicopter type model and serial number, aviation unit destination and
returned to supporting ammunition supply activity in the container used
to transport the replacement cartridge.
NOTE
All Retirement Life Items will have a Demil Code of "L" and will be
mutilated in accordance with DOD 4160.25-M-1, DEFENSE
DEMILITARIZATION MANUAL.
Change 71
1-72A/(1-72B blank)
TM 55-1520-234-23
CHAPTER 2
AIRFRAME
2-1. Airframe.
a. This chapter contains instructions for AVUM
(Aviation Unit Maintenance) and AVIM (Aviation
Intermediate Maintenance) on the helicopter airframe.
The airframe consists of the
fuselage, tailboom, and wings. See figure 2-1.
The structural panels shown on figure 2-2 must be installed
prior to helicopter ground run, flight, or ground handling.
NOTE
The nonstructural access panels and
doors shown on figure 2-3 may be
removed during helicopter ground run.
All fasteners shall be installed in
structural panels. Nonstructural panels
may have every third fastener missing,
however, no panel shall have more than
thirty-three percent of the total number
of fasteners missing.
b. The damage limits provided in chapter 2 on
bonded panels are not intended to red X the aircraft. The
limits are to provide guidance for scheduling repair or
replacement at a schedule maintenance interval. When
damage limits, particularly bond voids in bonded panels,
are exceeded, the responsible unit maintenance authority
will establish a reoccurring special inspection on the
damaged area until the damage to the structure is
corrected. If the damage is in the area that requires
engineering authority for repair, engineering should be
contacted in writing with a description of damage. If depot
assistance will be required, unit should contact AVSCOM,
AMSAV-MDP with your requirement.
Section I. FUSELAGE
2-2. Fuselage.
The fuselage constitutes the primary structural assembly of
the helicopter. It encloses and/or supports such major
provisions and systems as the tandem crew compartment,
the engine, the fuel and oil systems, the armament system,
the transmission and main rotor pylon, the alighting gear,
the wings and the tailboom. See figure 2-1.
the seal and burn the sample. Silicone seals burn readily
and leave a gray ash residue. Rubber-type seals are more
fire resistant and leave a black ash residue.
d.
Installation.
Provide adequate ventilation when using
methyl-ethyl-ketone. Avoid breathing solvent
vapors and avoid prolonged skin contact.
2-3. Seals on Doors, Cowlings and Fairings.
Many of the doors, cowlings and fairings have replaceable
seals. The seals may be either of rubber or silicone
composition. Seals that are subjected to fuel and/or oil
contamination are of the polysulfide or neoprene rubber
type.
Do not permit methyl-ethyl-ketone to contact
acrylic windows of canopy or canopy doors.
a. Inspect seals for failed bonding, tears, breaks
and deterioration that would affect function.
NOTE
b. Removal. Remove damaged or worn seal. Use
paint remover (C105) to remove old adhesive, paint and
primer from area where new seal will be installed.
It is necessary to thoroughly clean surfaces
prior to sanding to avoid working foreign
matter into pores of material.
c. Test to determine seal material. If the type
material from which the seal is made must be determined,
cut a small sample of material from
(1) Clean new seal and the metal where it is to
be applied with methyl-ethyl-ketone (C87) and dry with a
clean cloth. Sand the mating surfaces of
Change 54
2-1
TM 55-1520-234-23
Figure 2-1. Fuselage, wings, pylon mount and tailboom
Change 7
2-2
TM 55-1520-234-23
Figure 2-2. Structural panels
Change 42
2-3
TM 55-1520-234-23
Figure 2-3. Non-structural access panels, doors and fairings (Sheet 1 of 4)
Change 44
2-4
TM 55-1520-234-23
Figure 2-3. Non-structural access panels, doors and fairings (Sheet 2 of 4)
2-5
TM 55-1520-234-23
ITEM
1.
2.
3.
4.
5.
6.
7.
8.
ACCESS TO:
9.
10.
Deleted
Lower Fairing (left and right)
Lower Fairing (left and right)
Outer Panel (left and right)
Outer Panel (left and right)
Turret Fairing (left and right)
Turret Access Door (left and right)
Ammunition Compartment Door
(left and right)
Outer Panel (left and right)
Outer Panel (left and right)
11.
Access Door (left and right)
12.
13.
Forward Pylon Fairing
Center Pylon Fairing (left and right)
14.
15.
16.
17.
25.
Access Door (left and right)
Aft Pylon Fairing
Window
Transmission Cowl Assembly
(left and right)
Engine Cowl Assembly (left and right)
Tailpipe Fairing
driveshaft
External Power Door (left side only)
Oil Cooler Duct Panel (left side only)
Access Panels Assembly
(left and right)
Access Door
Gunners Floor Access Panel
(left side)
Pylon Access Panels
26.
Access Plate and Plug
27.
28.
Deleted
Ammunition Compartment
Aft Panel
Forward Crosstube Fairing
Lower Skin Panel
Drain Cover
*Lower Skin Panel
Lower Skin Panel
Drain Cover
Aft Crosstube Fairing
Lower Skin Panel
Lower Skin Panel
FM Sense Antenna
Jack Point Opening
Wing Inboard Access Covers
(right)
18.
19.
20.
21.
22.
23.
24.
29.
30.
31.
32.
33.
34.
35.
36.
37.
38.
39.
40.
Deleted
Armament
Telescopic sight unit
Flight controls
Flight controls
Turret exterior
Armament
Ammunition stowage, leveling
points
Flight controls
Structural fuel cell panel
(left access)
Hydraulic reservoirs and modules,
air distribution ducts
Cabin air intake, rotating controls
Rotating controls, mast rotating
beacon
Rotating controls
Engine Oil Tank
Transmission oil sight gage
Transmission, driveshafts, engine
air induction
Engine compartment
Exhaust tailpipe, tail rotor
External power receptacle
Oil coolers, turbine fan
Lower transmission, lift beam,
hydraulic units, control linkage
Tow interface unit
Armament turret
Pylon hydraulic and
electrical units
Telescopic sight unit
test connection
Deleted
SCAS transducer,
hydraulic lines
Forward crosstube supports
Forward fuel cell sump
Forward fuel drain
Deleted
Aft fuel cell sump
Aft fuel drain
Aft crosstube supports
Control linkages
Electrical cables
Antenna and SCAS control tube
Jack and mooring point
Tow hydraulic and electrical
installation
*NOTE
After incorporation of MWO 55-1520-234-50-3, lower skin panel is
replaced by relocated loop antenna.
209030-395-2
Figure 2-3. Non-structural access panels, doors and fairings (Sheet 3 of 4)
Change 44
2-6
TM 55-1520-234-23
41.
42.
43.
44.
45.
46.
47.
Wing Outboard Covers (right)
Wing Inboard Covers (left
Wing Outboard Access Covers (left)
Driveshaft Center Cover
Driveshaft Aft Cover
Gearbox Cover
Fin Driveshaft Cover
48.
49.
Gearbox Fairing and Cover
Avionics Compartment Door
(left)
Tail Skid Access Covers
Aft Fin Fairing
Lower Fin Inspection Cover
Tail Boom Access Door
Tail Boom Access Door
Fuel Cell Access Panel
Access Panel
Access Panel
Access Panel
Access Door (right side only)
Driveshaft Forward Cover
50.
51.
52.
53.
54.
55.
56.
57.
58.
59.
60.
Tow hydraulic and electrical installation
Tow hydraulic and electrical installation
Tow hydraulic and electrical installation
Tail rotor driveshaft
Tail rotor driveshaft
42 degree gearbox
Tail rotor driveshaft,
control cables
90 degree gearbox
Avionics. electronics equipment and
cooling fan
Tail skid attach point
Tail structure
Tail structure
Control linkage
Cooling fan
Fuel Cell
Telescopic sight unit wiring
Controls - gunner
Telescopic sight unit wiring
Interface control unit
Tail rotor driveshaft
209030-231-4C
Figure 2-3. Non-structural access panels, doors and fairings (Sheet 4 of 4)
both seal and metal with 180 grit sandpaper (C112). Clean
the sanded surface with methyl-ethyl-ketone (C87).
(2)
NOTE
Adhesive (C14) will cure at temperatures as
low as 50 degrees F. For each 12 degrees F
below 75 degrees F the cure time must be
doubled. No attempt to bond should be made
when temperature is below 50 degrees F.
Where possible bond surfaces should be
heated to 75 degrees F by use of heat lamp or
heat gun.
Bond rubber-type seals as follows:
(a)
Clean surfaces as outlined in step (1).
(b) Refer to step c for instructions to
identify rubber-type seals.
(c) Apply an even coat of rubber
adhesive (C14) to the mating surfaces of the seal and the
metal.
(d) Allow adhesive to air dry 10 to 15
minutes at 75 degrees F. or above. Check adhesive by
touching with finger. When adhesive will adhere to finger
but not transfer, apply a second coat of adhesive and air
dry to the same degree.
(e) When second coat of adhesive has
air dried until tacky, install seal on metal. Start at one end
and roll seal onto metal. Press down on seal to ensure that
all air is expelled and that the seal is in full contact with the
metal.
(f) Allow bond to air dry for a minimum
of four hours at 75 degrees F or above.
(3)
Bond silicone composition seals as follows:
(a)
Clean surface, as outlined in step (1).
(b) Refer to step c for instructions to
identify silicone composition seals.
Do not place a cap on the adhesive used in the following
step after it is mixed. This two-part adhesive releases
hydrogen gas
Change 54
2-7
TM 55-1520-234-23
after mixing which could result in high
pressures. The pot life on the mixed adhesive
is six hours.
and 2-68A for specific damage limits. Replace panels with
damage in excess of the limits shown on the illustrations.
Refer to paragraph 2-13 for instructions to replace panels.
(c) Mix adhesive (C18) in accordance
with instructions on the container. Apply an even coat of
adhesive to the mating surfaces of the seal and the metal.
b. Repair mechanical damage to fuselage and
tailboom fin honeycomb panels as outlined in this
paragraph. Repair damaged fasteners (inserts) as outlined
in paragraphs 2-5 through 2-10.
(d) Allow the adhesive to air dry at 75
degrees F. or above for at least one hour but not more than
eight hours. Install seal on metal. Start at one end and roll
seal onto metal. Press down on seal to ensure that all air
is expelled and that the seal is in full contact with the
metal.
NOTE
Adhesive (C 18) will cure at temperatures as
low as 50 degrees F. For each 12 degrees F
below 75 degrees F the cure time must be
doubled. No attempt to bond should be made
when temperature is below 50 degrees F.
Where possible bond surfaces should be
heated to 75 degrees F by use of heat lamp or
heat gun.
d.
(e) Allow bond to cure for a minimum of
twelve hours at 75 degrees F. or above.
(4) Functional check. Install door, cowling or
fairing and check to ensure that the new seal fits properly.
2-4. Honeycomb
Construction
Tailboom Fin Panels.
Fuselage
c. Dent, Scratch, Void and Penetration Damage
Repair. Damage to honeycomb panels varying from minor
dents to penetration completely through the panel is
classified as type A, type B, type C or type D damage. The
damage descriptions, damage limits and repair procedures
are shown in figures 2-20 through 2-23. Specific damage
limits for tailboom vertical fin panels are shown in figure 268A. When type A through type D damage is present on a
panel and is in the "repair permissible" area as shown on
figures 2-5 through 2-19, repair the damage as shown on
figures 2-20 through 2-23 as applicable.
and
The principal part of the fuselage and fin structure is
honeycomb panels. The panels have an aluminum core
that resembles honeycomb. Facings are bonded to the
honeycomb to form the panel. The facings may be
fiberglass or metal. The fuselage panels are joined
together and supported by the primary structural caps
which are shown on figure 2-4 by solid black shading.
Panels on the unshaded portion of figure 2-2 are either of
honeycomb panel construction or of conventional sheet
metal construction. Refer to TM 55-1500-204-25/1 for
repair instructions for the sheet metal construction panels.
a.
Inspection. Inspect honeycomb panels for
cracks, punctures, corrosion, delamination and damaged
inserts. Refer to figures 2-5 through 2-19
Change 2
Standard Patch Repairs for Honeycomb Panels.
(1) Use the same materials to fabricate
patches that were used in the original construction with two
exceptions: Stainless steel (item 58, table 2-2) 1/4 hard or
harder may be used to repair honeycomb panels which
have titanium skin. Other material substitutions can be
made when qualified authority approves the substitute
material.
(2) Repair damage that is within limits shown
on figures 2-5 through 2-23 and 2-68A. Materials required
and repair procedures are shown on the illustrations. The
chemicals, adhesives and compounds required are listed in
the consumable materials table. Instructions for using
these materials are on the containers.
e.
Edge Repairs for Honeycomb Panels.
Repair damage that is within limits as shown on
figure 2-5 through 2-19.
Comply with the following
additional instructions for fiberglass and metal faced panels
shown on figures 2-24 through 2-28.
(1)
Fiberglass
(a) Use only type C fiberglass cloth 0.010
inch thick (C46) when making edge repairs. The repair
must equal or exceed the number of plies lost.
2-8
TM 55-1520-234-23
(b) Remove all old finish from repair area
with varying grades of sandpaper (C112).
Provide adequate ventilation when using
methyl-ethyl-ketone. Avoid breathing solvent
vapors and avoid prolonged skin contact.
(c) Clean sanded area with clean cloth
moistened with methyl-ethyl-ketone (C87).
Change 2
2-8A/(2-8B blank)
TM 55-1520-234-23
Figure 2-4. Primary structural cap. (Sheet 1 of 2)
2-9
TM 55-1520-234-23
Figure 2-4. Primary structural caps (Sheet 2 of 2)
2-10
TM 55-1520-234-23
Figure 2-5. Pilot and gunner floor panels
Change 7
2-11
TM 55-1520-234-23
Figure 2-6. Bulkhead of station 93.0
Change 7
2-12
TM 55-1520-234-23
Figure 2-7. Bulkheads at stations 148.5 and 171.61
Change 4
2-13
TM 55-1520-234-23
Figure 2-8. Bulkheads at stations 184.5 and 213.94
Change 7
2-14
TM 55-1520-234-23
Figure 2-9. Bulkheads at stations 250.0 and 268.5
Change 7
2-15
TM 55-1520-234-23
Figure 2-10. Right and left main beam panels at station 148.5 to 186.25
Change 7
2-16
TM 55-1520-234-23
Figure 2-11. Right and left main beam panes at station 213.94 to 250.0
Change 2
2-17
TM 55-1520-234-23
Figure 2-12. Panel at forward fuel cell at right side and gunners floor
2-18
TM 561520-23423
Figure 2-13. Left and right beam panels at station 250.0 to boom station 41.32
Change 2
2-19
TM 55-1520-234-23
Figure 2-14. AmmoFloor, Support Panel and Forward Fuel Cell Panel at Station 213.94.
Change 29
2-20
TM 55-1520-234-23
Figure 2-15. Forward fuel cell floor and lower panel at station 186.256 to 213.94
Change 2
2-21
TM 55-1520-234-23
Figure 2-16. Lower aft fuel cell panel and bottom panel at station 250.0 to boom station 41.32
2-22
Change 2
TM 55-1520-234-23
Figure 2-17. Engine deck installation at stations 213.94 to 298.75
Change 2
2-23
TM 66-1520-234-23
Figure 2-18. Forward fuel cell panels and main beam panels at station 186.25 to 214.0
Change 2
2-24
TM 55-1520-234-23
Figure 2-19. Vertical fin honeycomb panels
Change 7
2-25
TM 55-1520-234-23
FIBERGLASS OR METAL FACINGS
FIBERGLASS FACED PANELS
METAL FACED PANELS
DESCRIPTION
Dents. scratches. scars. or erosion in facings with no
holes, cracks. or voids.
Smooth dents or depressions in the skins with no
holes or cracks. (See Type C damage for repairs to
penetrating damage.)
LIMITS - REPAIRABLE DAMAGE
1. Maximum depth: 25% of panel thickness.
1. Maximum diameter of damage: 0.50 inch.
2. Minimum distance from an edge bevel: 0.5 inch
2. Maximum depth: 20% of panel thickness.
3. Maximum area of all dents combined: 5% of
panel surface area.
4. Maximum of five dents in a 3.0 square inch area.
5. No voids may be present under the damage.
REPAIR PROCEDURES
1. Smooth out damaged area by lightly sanding.
2. Clean with methyl-ethyl-ketone (C87) and allow to dry.
SAME AS FIBERGLASS REPAIR
3. Brush on adhesive (C123) to level to contour
and allow to cure.
4. When cured, sand smooth and refinish if required.
209030-19G
Figure 2-20. Type A damage - body panel repairs
Change 7
2-26
TM 55-1520-234-23
Voids between the facings and core and separations between laminations of facings on metal or fiberglass
panels.
FIBERGLASS FACED PANELS
METAL FACED PANELS
LIMITS - REPAIRABLE DAMAGE
1. Maximum area of all damage: 4.0 square inches
or 5% of panel surface area whether as a single
void or combination of separate voids.
1. Maximum area of all damage: 5% of total area of
panel with aluminum or stainless steel skin and
3% with titanium skin.
2. Maximum length of a void: 4.0 inches in any direction.
2. Maximum area of single void: 1.5 square inches
for aluminum and stainless steel. 1.0 square
inch for titanium.
3. Damage is not repairable within 0.50 inch of
any beveled edge.
3. Voids within 3.0 inches of any structural
member and within 0.50 inch of a beveled edge
are not repairable.
4. Maximum length of a void 3.0 inches in any
direction for aluminum and stainless steel and
2.5 inches for titanium.
REPAIR PROCEDURES
1. Drill No. 40 ((or smaller) holes around edge of
damage a minimum of 1.0 inch apart. Use as
many holes as required to ensure complete
filling of cavity.
2. Inject epoxy resin (C107) with hypodermic
syringe until resin is forced out opposite hole.
SAME AS FIBERGLASS REPAIR
3. Cover repair with cellophane (C33) and level out
by clamping with blocks. Allow to cure.
4. Seal holes with adhesive (C123).
5. Clean up and smooth with fine sandpaper (112).
Refinish if required.
209030-20G
Figure 2-21. Type B damage - body panel repairs
Change 7
2-27
TM 55-1520-234-23
FIBERGLASS FACED PANELS
METAL FACED PANELS
DESCRIPTION
Sharp dents and dents containing holes and cracks
built not extending completely through panel. (See
Type D for through limits and damage greater than
0 50 Inch diameter, figure 2-23 )
Tears, fractures, and holes through fabric skins with
no damage to core. (See Type D damage limits for
core damage, figure 2-23 )
LIMITS - REPAIRABLE DAMAGE
1. Maximum area of damage 9.0 square Inches or
5% of total panel area whether a single area or
combination of separate areas
1. Maximum diameter of hole after cleanup:
0.50 inch. (See Type D for damage over 00.50 inch.)
2. Maximum number of repairs per panel: One.
3. Minimum distance from structural member
fitting, or insert: 1.0 inch.
4. Minimum distance from beveled edges: 0.50 inch.
209030-21-1G
Figure 2-22. Type C damage - body panel repairs (Sheet 1 of 2)
Change 7
2-28
TM 55-1520-234-23
REPAIR PROCEDURES
FIBERGLASS FACED PANELS
METAL FACED PANELS
1. Smooth damaged surface by light sanding.
1. Counterbore area to the diameter and
depth required to clean out damage.
(Maximum diameter 0.50 inch.)
2. Cut the required number of plies from
fiberglass cloth (C46 or C47).
2. Pack cavity with adhesive (C17).
3. Saturate each ply with epoxy resin (C107)
and place over damage.
3. Level out flush with skin and cure.
4. Cover patch with cellophane (C33) or
Tedlar (C139). Press down to smooth and
allow to cure.
4. Cut required number of doublers from the
some material as the akin. If titanium is not
available, use equivalent thickness
stainless steel to patch titanium skin.
5. If necessary sand repair to smooth out and refinish if
required.
5. Bevel the edges of doublers on top side.
6. Clean all surfaces with methyl-ethyl-keytone (C87).
7. Apply adhesive (C17) and center doublers
over damage. Clamp smoothly with blocks
and allow to cure. Refinish if required.
209030-21-2E
Figure 2-22. Type C damage - body panel repairs (Sheet 2 of 2)
Change 22
2-29
TM 55-1520-234-23
FIBERGLASS FACED PANELS
METAL FACED PANELS
DESCRIPTION
Damage penetrating metal skins greater
than 0.50 inch diameter. and damage
extending completely through panel.
Damage penetrating the facings and
extending into core. Same limits apply to
damage through one skin only and damage
completely through panel.
LIMITS - REPARABLE DAMAGE
1. Maximum area of damage after clean
up 6.0 square inches. whether a single
area or combination of separate areas.
1. Maximum damaged area after clean up:
Total of 9.0 square inches or 5% of panel
surface area per panel. Applies whether
a single area or combination of separate areas.
2. Maximum length: 3.5 inches in any direction.
2. Maximum length of damage: 4.0
inches in any direction
3. Minimum distance from structural
members or other repair: 3.0 inches.
3. Maximum diameter of clean up
counterbore: 3.75 inches.
4. Minimum distance from an edge bevel: 0.50 inches.
4 Minimum distance from an edge bevel 0 50 inch.
209030-22-1H
Figure 2-23. Type D damage - body panel repairs (Sheet 1 of 2)
Change 7
2-30
TM 655-1520-234-23
REPAIR PROCEDURES
FIBERGLASS FACED PANELS
METAL FACED PANELS
1. Clean up damage with counterbore or
hole cutter. If damage is limited to one
side of panel, counterbore only deep
enough for proper cleanup.
1. Clean up damage with counterbore or
hole cutter. If damage is limited to one
side of panel, counterbore only deep
enough for proper cleanup.
2. Pack hole with adhesive (C17). Level
out flush with surface of panel. Allow to cure.
2. Pack cavity with adhesive (C17).
Smooth flush with surface of panel
and allow to cure.
3. Cut required number of patch layers
from fiberglass cloth (C46 or C47) as shown.
4. Saturate each cloth layer with epoxy
resin (C107) and place over damage.
5. Cover patch with cellophane (C33) or
Tedlar (C139). Press down smoothly
and allow to cure.
6. If necessary, sand smooth and refinish.
3. Cut required number of doublers from
the same material as the skin. If
titanium is not available use equivalent
thickness stainless steel to patch titanium skin.
4. Clean all surfaces with methyl-ethyl-ketone (C87).
5. Apply adhesive (C17) and center
doublers over damage. Clamp
smoothly with blocks and allow to
cure. Refinish if required.
209030-22
Figure 2-23. Type D damage - body panel repairs (Sheet 2 of 2)
Change 22
2-31
TM 55-1520-234-23
EDGE REPAIR OF HONEYCOMB PANELS
NOTE
Typical edge repair of
applicable panels, for cross
hatched area as illustrated on
referenced figure. Maximum
repairable damage is 1.25 inch
diameter after cleanup. Repair
of damage exceeding 1.25 inch
diameter must be approved by
qualified Engineering authority.
APPLICABLE PANELS
209-030-119
209-030-120
209-030-122
209-030-206
209-030-127
209-030-108
209-030-123
209-030-124
209-030-204
209-030-209
209-030-801
209-030-130
209-030-138
REF. FIG.
Left Fuel Cell
Right Fuel Cell
Bulkhead
Bottom of Fuel Tank
Lower Bulkhead Sta. 186.25
Bulkhead Sta. 213.94
Upper Bulkhead Sta. 148.56
Lower Bulkhead Sta. 148.50
Forward Fuel Cell Floor
Engine Deck Sta. 213.94 and Sta. 298.75
Tail Fin
Right Main Beam
Left Main Beam
Figure 2-24. Edge repair for honeycomb panel with fiberglass skin opposite titanium
Change 7
2-32
2-11
2-11
2-9
2-16
2-8
2-8
2-7
2-7
2-15
2-17
2-19
2-10
2-10
TM 55-1520-234-23
EDGE REPAIR OF HONEYCOMB PANELS
APPLICABLE PANELS
NOTE
Typical edge repair of applicable panels
for cross hatched area as illustrated on
reference Figure. Maximum repairable
damage 1.25 in. dia. Repair of
damage greater than 1.25 in. dia. must
be approved by qualified Engineering
authority.
209-030-201
209-030-202
209-030-125
209-030-137
209-030-112
209-030-102
209-030-207
209-030-117
209-030-118
209-030-126
209-030-106
209-030-135
209-030-269
209-030-270
209-030-205
209-030-213
209-030-219
Gunners Floor
Pilots Floor
Upper Fuel Cell
Main Beam
Bulkhead Sta. 268.65
Bulkhead Sta. 93.0
Floor
Left Beam
Right Beam
Upper Bulkhead Sta. 186.25
Center Bulkhead Sta. 164 to 171.61
Gunners Seat
Panel Sta. 155.97
Fwd. Fuel Tank Support
Panel (Sta. 186.25 to 213.94)
Pylon Support
Ammo. Floor
Figure 2-25. Edge repair for honeycomb panel with aluminum
alloy skin opposite aluminum alloy with fiberglass edging
Change 7
2-33
REF. FIG.
2-5
2-5
2-18
2-18
2-9
2-6
2-16
2-13
2-13
2-8
2-7
2-12
2-12
2-12
2-15
2-14
2-14
TM 55-1520-234-23
REPAIR OF CHANNEL SECTION
OF HONEYCOMB PANELS
NOTE
APPLICABLE PANELS
Typical repair at channel section of
applicable panels, cross hatched on
reference figure. Maximum repairable
damage 1.25 inch diameter after
cleanup. Repair damage greater than
1.25 inch diameter must be approved
by qualified Engineering authority.
209-030-201
209-030-202
209-030-125
209-030-137
209-030-102
209-030-117
209-030-118
209-030-124
209-030-204
209-030-213
209-030-219
209-030-206
REF. FIG.
Gunner's Floor
Pilot's Floor
Upper Fuel Cell
Main Beam
Bulkhead Sta. 268.65
Left Beam
Right Beam
Lower Bulkhead Sta. 148.50
Fwd. Fuel Cell Floor
Pylon Support
Ammo Floor
Lower Aft Fuel Cell Panel
Figure 2-26. Edge repair for honeycomb panels with aluminum alloy skin
opposite aluminum alloy with fiberglass finish at channel
Change 7
2-34
2-5
2-5
2-18
2-18
2-6
2-13
2-13
2-7
2-15
2-14
2-14
2-16
TM 55-1520-234-23
REPAIR OF CHANNEL SECTION OF HONEYCOMB PANELS
APPLICABLE PANELS
209-030-201
209-030-202
209-030-125
209-030-137
209-030-102
209-030-117
209-030-118
209-030-124
209-030-204
209-030-213
209-030-219
209-030-206
REF. FIG.
Gunner's Floor
Pilot's Floor
Upper Fuel Cell
Main Beam
Bulkhead Sta. 268.65
Left Beam
Right Beam
Lower Bulkhead Sta. 148.50
Forward Fuel Cell Floor
Pylon Support
Ammo Floor
Lower Aft Fuel Cell Panel
Figure 2-27. Typical rivet pattern for channel section repair
Change 2
2-35
2-5
2-5
2-18
2-18
2-6
2-13
2-13
2-7
2-15
2-14
2-14
2-16
TM 55-1520-234-23
Figure 2-28. Edge repair on vertical fin
Change 7
2-36
TM 55-1520-234-23
(d) Cut fiberglass cloth (C46) to correct size
and saturate with epoxy resin (C107) and apply as a patch.
(e) If multiple layers of fiberglass are
required, overlap each successive patch for a
minimum distance of one inch.
Provide adequate ventilation when
using methyl-ethyl-ketone. Avoid
breathing solvent vapors and avoid
prolonged skin contact.
(2) Metal (Aluminum alloy, titanium or stainless steel).
(a) Stainless steel (item 58, Table 2-2) 1/4
hard or harder may be substituted for titanium.
Use stainless steel of same thickness as that
specified for tatanium patch.
(b) The minimum thickness of patches are
specified on figures 2-24 through 2-26.
d. Remove excessive adhesive with cloth moistened
with methyl-ethyl-ketone (C87) or naphtha (C88).
2-6. Damage Fastener (Insert) In Vertical Fin
Panel - Replacement.
a. Remove damaged fastener (insert) by
machining with counter bore of the same diameter.
If the fastener (insert) is loose and turns, drill out
two holes shown for injecting adhesive on figure 229. Use a spacer on twist drill while drilling out
holes to avoid drilling too deep and damaging
pane]. Attach a puller to fastener (insert) with selftaping screws and remove the fastener (insert)
from the panel.
Provide adequate ventilation when
using methyl-ethyl-ketone. Avoid
breathing solvent vapors and avoid
prolonged skin contact.
(c) Remove all old finish from repair area
with varying grades of sandpaper (C112). Clean
sanded area with clean cloth moistened with
methyl-ethyl-ketone (C87).
(d) Bond metal patches with adhesive(C17).
(e) After adhesive (C17) has cured, apply
sealant (C 115) to entire surface of titanium patch.
(f) Rivet patch with rivets of equator larger
size than original rivets in areas that were riveted
prior to application of patch. Use the standard
edge distance of two rivet diameters or space the
rivets the same as the original panel. If the panel
being worked is not riveted, use rivet spacing in
opposite panel.
(g) Fastener (Insert) Repairs. Repair loose
and/or damaged fasteners (inserts) as described in
applicable paragraph 2-5 through 2-10.
2-5. Loose Fastener (Insert) In Vertical Fin Panel Repair.
a. Drill one, or two where possible, 1/16 inch
diameter holes in surface of panel near edge of
fastener. Drill the hole or holes at an angle in
direction of fastener (insert).
b. Inject adhesive (C12) in drilled hole or holes
with a syringe. Fill the cavity. If there are two holes,
fill until adhesive exudes from the second hole.
c. Allow adhesive to cure for 24 hours at room
temperature.
Change 7
b. Crush back honeycomb core a minimum of
0.063 inch and maximum of 0.25 inch on figure 229. Clean all metal particles out of hole.
Provide adequate ventilation when
using methyl-ethyl-ketone. Avoid
breathing vapors and avoid prolonged
skin contact.
CAUTION
Do not use methyl-ethyl-ketone to clean
hole in panel.
c. Immediately prior to installation, clean new
fastener (insert) with methyl-ethyl-ketone (C87)
and air dry until moisture free. Handle fastener
(insert) with clean gloves after cleaning.
d. Cover threaded hole and injection holes with
masking tape then open the injection holes with a
pointed instrument. Apply adhesive (C12) to
bottom of fastener (insert) as shown on figure 2-29
and position in hole in panel. Inject adhesive (C 12)
in one injection hole with a syringe until it comes
out of the opposite injection hole as shown on
figure 2-29.
e. Remove excess adhesive with cloth
dampened with methyl-ethyl-ketone (Cb7) or
naphtha (C33).
f. Touch up paint to match surrounding area.
2-37
TM 55-1520-234-23
Figure 2-29. Injection - type fastener (insert) in vertical fin panel
Change 7
2-38
TM 55-1520-234-23
2-7. Damaged Fastener (Insert) In Fuselage
Honeycomb Panels - Replacement.
a. Determine whether the fastener (insert) is a
potted - type, Injection - type or Grommet - type. See
figure 2-30 for view of the fasteners (inserts).
b. Remove damaged fastener (insert) by
machining with a counterbore of the same
diameter as the fastener. Note that the grommet type fastener (insert) flanges overlap the skin of
the honeycomb panel.
c. Install new fastener (insert) as outlined in
paragraphs 2-8 through 2-10.
2-8. Potted - Type Fastener (Insert) Installation in
Honeycomb Panel.
Provide adequate ventilation when
using methyl-ethyl-ketone. Avoid
breathing solvent vapors and avoid
prolonged skin contact.
d. Remove excess adhesive from honeycomb
panel before adhesive sets up. Use cheese cloth
(C36) dampened with methyl-ethyl-ketone (C87).
Exercise caution to prevent the methyl-ethylketone from diluting the adhesive in the potted
areas.
e. Touch up paint to match surrounding area.
2-9. Injection - Type Fastener (Insert) Installation
in Honeycomb Panel.
Provide adequate ventilation when
using methyl-ethyl-ketone. Avoid
breathing solvent vapors and avoid
prolonged skin contact.
a. Immediately prior to installation, clean new
fastener (insert) by soaking in methyl-ethyl-ketone
(C87). Air dry until moisture free.
b. Place masking tape (C134) over threads of
fastener (insert) to prevent entry of adhesive.
CAUTION
An insufficient amount of adhesive will
allow moisture or other fluids to enter
the honeysomb panel core. This will
result in ultimate failure of the panel.
c. Fill cavity approximately two-thirds full of
adhesive as shown on figure 2-30. Use metal set A4
adhesive (C 12) in areas where temperature will not
exceed 180 degrees F (82 degrees C). Use adhesive
(C17) in areas where panel will be subjected to
higher temperatures but not exceeding 300 degrees
F (149 degrees C). Install fastener (insert) while
adhesive is in tacky state. Ensure that there are no
pin holes in the adhesive, and that the fastener
(insert) is properly aligned and is a snug fit where
the fastener (insert) flange mates with the
honeycomb panel face.
Provide adequate ventilation when
using methyl-ethyl-ketone. Avoid
breathing vapors and avoid prolonged
skin contact.
a. Immediately prior to installation, clean new
fastener (insert) by soaking in methyl-ethyl-ketone
(C87). Air dry until moisture free.
b. Place masking tape (C134) over threads and
injection holes of fastener (insert) to prevent entry
of adhesive. Open holes in the tape at the injection
holes with a pointed instrument to permit injection
of adhesive compound.
c. Apply a layer of adhesive to bottom of
fastener as shown in figure 2-29. Use adhesive
(C12) in areas where temperature will not exceed
180 degrees F (82 degrees C). Use adhesive (C17) in
areas where panel will be subjected to higher
temperatures but not exceeding 300 degrees F (149
degrees C). Position the fastener (insert) in the
hole.
CAUTION
An insufficient amount of adhesive will
allow moisture or other fluids to enter
the honeycomb panel core. This will
result in ultimate failure of the panel.
d. Inject same adhesive used in the preceding
step into one injection hole until a steady stream of
adhesive, without air bubbles, flows out of the
Change 7
2-39
TM 55-1520-234-23
Figure 2-30. Potted - type, injection - type, and grommet - type fasteners (inserts) (Sheet 1 of 2)
Change 7
2-40
TM 55-1520-234-23
Figure 2-30. Potted - type, injection - type, and grommet - type fasteners (inserts) (Sheet 2 of 2)
2-41
TM 55-1520-234-23
opposite injection hole. Use a syringe to inject
adhesive as shown on figure 2-30.
e. Ensure that fastener (insert) is properly aligned.
panel will be subjected to higher temperatures but
not exceeding 300°F. Install the sleeve and the plug
in their correct relative position in the panel.
Lightly tap the two parts together. Ensure that the
flanges are seated and properly aligned with the
panel.
NOTE
Provide adequate ventilation when using
methyl-ethyl-ketone. Avoid breathing
solvent vapors and avoid prolonged skin
contact.
A screw and washer may be installed in
the fastener (insert) to hold it in position
and prevent adhesive from getting on
threads. Use a parting material, such as
cellophane, under the washer to prevent
it from adhering to the fastener (insert).
f. Remove excess adhesive from honeycomb
panel before adhesive sets up. Use cheese cloth (C36)
dampened with methyl-ethyl-ketone (C87). Exercise
caution to prevent the methyl-ethyl-ketone from
diluting the adhesive in the potted areas.
CAUTION
An insufficient amount of adhesive will
allow moisture or other fluids to enter the
honeycomb panel core. This will result in
ultimate failure of the panel.
g Touch up paint to match the surrounding area.
2-10. Grommet - Type Fastener (Insert)
Installation In Honeycomb Panel.
e. Inject the same adhesive used in the preceding
step into one injection hole until a steady flow of
adhesive, without air bubbles, comes out of the
opposite hole.
Provide adequate ventilation when using
methyl-ethyl-ketone. Avoid breathing
solvent vapors and avoid prolonged skin
contact.
a. Immediately prior to installation, clean new
fastener (insert) by soaking in methyl-ethyl-ketone
(C87). Air dry until moisture free. Handle fastener
(insert) with clean white gloves after cleaning.
b. Place masking tape (C134) over threads of
fastener (insert) to prevent entry of adhesive.
c. Position the sleeve half of fastener (insert) in
honeycomb panel and mark location of two injection
holes as shown in figure 2-30. Make hole centers /8
inch from edge of flange as illustrated. Remove
sleeve and drill two holes with size 42 twist drill.
Make hole through honeycomb panel face at ninety
degrees, then slant drill as illustrated. Deburr holes
and clean all debris from cavity.
d. Apply a small bead of adhesive under flanges
of sleeve and plug as illustrated on figure 2-30. Use
adhesive (C12) in areas where temperature will not
exceed 180°F. Use adhesive (C17) in areas where
Provide adequate ventilation when using
methyl-ethyl-ketone (C87). Avoid
breathing vapors and avoid prolonged
skin contact.
f. Remove excess adhesive from honeycomb
panel before adhesive sets up. Use cheese cloth (C36)
dampened with methyl-ethyl-ketone (C87). Exercise
caution to prevent the methyl-ethyl-ketone from
diluting the adhesive in the potted areas.
g. Touch up paint to match the surrounding area.
2-11. Honeycomb Construction Pylon Fairings.
The pylon fairing portion of the fairing and
cowling shown on figure 2-31 is honeycomb
construction. The honeycomb cores are either
fiberglass or aluminum. Facings are fiberglass.
These fairings are not structural in nature and do
not carry primary loads; therefore, larger size
2-42
TM 55-1520-234-23
damage may be repaired on these fairings thorn can
be repaired on the fuselage honeycomb panels. It is
necessary to maintain contours and restore the
fairings to original strength when repair is
accomplished.
a. Inspection. Inspect honeycomb construction pylon
fairings for cracks, punctures and delamination.
(1) Compare any damage that is present with
the limits shown on figures 2-20 through 2-23 for
fuselage honeycomb panel damage. Record whether
damage is within limits.
(2) If there is any damage present that is in
excess of the limits noted in step (1), inspect to
determine whether the damage affects portions of
the honeycomb fairing that contain attachment
holes and/or fasteners. Panels with damage that
affects attachment holes and/or fasteners is not
reparable.
Figure 2-31. Fairing and cowling for pylon, transmission, engine, and tail pipe (Sheet 1 of 4)
2-43
TM 55-1520-234-23
Figure 2-31. Fairing and cowling for pylon, transmission, engine, and tail pipe (Sheet 2 of 4)
2-44
TM 55-1520-234-23
Figure 2-31. Fairings and cowling for pylon, transmission, engine, and tail pipe (Sheet 3 of 4)
2-45
TM 55-1520-234-23
Figure 2-31. Fairing and cowling for pylon
2-46
TM 55-1520-234-23
NOTE
This procedure describes repairs for
damage penetrating both faces of the panel.
When damage extends through one face
and into the core with no damage to the
opposite face, leave the undamaged face
intact. Take care to avoid cutting into the
opposite facing when removing damaged
core The remainder of the repair procedure
is the same.
(b) Sand a scarf (bevel) in the top facing
around the edge of the hole. Make the width of the
bevel about 10 times the thickness of the facing.
Figure 2-32. Fiberglass honeycomb fairing repair
Change 7
2-47
TM 55-1520-234-23
(I) After curing, remove weight and finish
opposite side as was done in the preceding steps.
(m) When second side of patch is
completely cured, finish by lightly sanding exterior side
of patch to the proper contour. Refinish with paint as
necessary.
The dust given off by sanding fiberglass
may irritate the skin or cause injury to the
respiratory system.
(c) Cut a strip of fiberglass cloth (C47) and
apply to the edge of the hole with resin (C107).
2-12. Ammo Floor Honeycomb Panel.
installation of scuff doublers.
Repair by
(d) Cut an insert of new core from material
of the same thickness as the original. Make the insert fit
snugly in the hole.
a. Repair damage that is within allowable limits as
outlined in paragraph 2-4.
(e) Place repair area over a form shaped to
the contour of the fairing and hold. in place with weights
or other suitable means.
b. Fabricate scuff doublers from stainless steel
(item 58, table 2-2) as shown in figure 2-33. The
required stock size is 0.016 x 7.0 x 44.20 inches.
(f) Make a temporary shim of the same
thickness as the inner facing, place in opening against
from and cover with cellophane (C33).
c. Make the 45 degree bend shown on figure 2-33
with a 0.06 inch radius.
(g) Coat the replacement core' around the
edge and over the top surface with resin (C107) and
place into hole.
(h) Cut the required number of fiberglass
both plies (C47). Make the largest piece to fit the
outside edge of the scarf. Cut each additional piece to
match the edges of each lower ply in the facing. Use as
many patch pieces as there are plies in the panel facing.
(i) Starting with the smallest piece, saturate
each one with resin (C107) and place over the repair. As
each piece is put down, smooth and lightly work out
excess resin and air pockets. Do not squeeze out.
(j) Lay a sheet of cellophane (C33) over the
patch after the final cloth piece has been put down.
Carefully work patch to smooth out and remove excess
resin and air pockets.
(k) Apply light pressure on patch with sandbags
or other weights and allow epoxy resin to cure.
d. Remove teflon rub strips. Retain the strips and
screws for reinstallation.
e. Use a hole finder tool to locate four holes in the
stainless steel doublers to match holes for screws that
retained the teflon rub strip. Drill holes in doubler.
b. Repair.
(1) Repair damage to the honeycomb fairings
that is within the limits described in step a(1) in the
same manner prescribed for fuselage panels in
paragraph 2-4.
(2) Repair damage to the honeycomb fairings
that is within the limits described in step a(2) by cutting
out the damaged area and inserting a new section of
honeycomb
(a) Trim-out the damaged material to a
circular or oval shape. Remove all rough and irregular
edges. See figure 2-32.
Figure 2-33. Ammo floor scuff doubler installation
2-48
Change 29
TM 55-1520-234-23
f. Remove and replace rub strip (209-030-224- 13)
on track assembly (209-030-224-19) if worn or
damaged. Use adhesive (C17).
and aft, to remove load from forward fuselage.
g. Position doublers on both sides of the ammo
floor panel. Place teflon strips removed in step d on the
scuff doublers and secure with original screws.
(4) Place jacks under jack points and raise until
hand-tight against fittings.
Alternate procedure: Remove tailboom.
2-13. Honeycomb Panel Removal and Installation.
a. Replace honeycomb panels that have damage
in excess of limits specified in paragraph 2-4.
When replacing any riveted structural
honeycomb panels or the right fuel cell
panel which is installed with screws,
structural loads must be relieved to maintain
alignment of airframe.
b. Use the following procedures to ensure that
airframe alignment is maintained.
Remove and install the panels listed in the
following step one at a time unless a work
aid fixture is used to restrain the structure
and maintain alignment or damage to the
fuselage may result.
(5) Remove and install the panels listed below
one at a time. See figure 2-34 for fuselage station
locations.
Panel, Main Beam - left and right Sta 61.0-148.5
P/N 209-030-129
Panel, Main Beam - left Sta 148.5-186.25 P/N 209030-138 (Figure 2-10)
Panel, Main Beam - right Sta 148.5-186.25 P/N 209030-130 (Figure 2-10)
Do not use aircraft structure as work
platform when structural panels or engine
decks are removed.
Require that all
personnel use work stands or airframe
warpage may result.
Panel, Main Beam - left and right Sta 186.25-214.0
P/N 209-030-137 (Figure 2-18)
(1) Attach hoist to mast retaining nut and
support the main rotor by hoisting vertically until the lift
link retaining bolt can be freely rotated. Check vertical
alignment of hoist with a spirit level or a clinometer. A
free lift link retaining bolt indicates that the load has
been removed.
Panel, Main Beam - right Sta 214.0-250.0 P/N 209030-120 (Figure 2-11)
Alternate method: Remove main rotor, mast,
controls, and transmission.
(2) Attach engine sling (T7) and hoist to engine.
Loosen pillow blocks on engine mounts. Support the
engine by hoisting vertically until engine is loose in the
pillow blocks. Check vertical alignment of hoist with a
spirit level.
Alternate method: Remove engine.
Chapter 4.
Refer to
(3) Support tailboom at two locations, forward
Change 48
Panel, Main Beam - left Sta 214.0-250.0 P/N 209030-119 (Figure 2-11)
Panel, Main Beam - left Sta 250.0-Boom Sta 41.32
P/N 209-030-117 (Figure 2-13)
Panel, Main Beam - right Sta 250.0-Boom Sta 41.32
P/N 209-030-118 (Figure 2-13)
(6) If the aft engine deck panel (Figure 2-7)
must be replaced, contact, AVSCOM, AMSAV- MEA for
information.
2-14. Sheetmetal Panels, Cowlings, and Fairings.
Sheet metal panels, cowlings and fairings are used
to cover openings and give an aerodynamically clean
contour. Honeycomb-constructed panels, cowlings and
fairings, which were discussed in
2-49
TM 55-1520-234-23
Figure 2-34. Station diagram (Sheet 1 of 2)
2-50
TM 55-1520-234-23
Figure 2-34. Station diagram (Sheet 2 of 2)
2-51
TM 55-1520-234-23
preceding paragraphs, are used for the same purpose.
Sheetmetal panels, cowlings and fairings are secured in
place by rivets, screws, twist-type fasteners or latches.
Refer to paragraphs 2-15 and 2-16 for transmission and
engine cowling instructions.
(3) Cowl assembly hinges for wear and damage
that affects function.
a. Inspection.
Inspect sheet metal panels,
cowlings and fairings for holes, cracks, corrosion,
deformation, and condition of seals and/or fasteners if
applicable
(5) Cowl assembly standoff for damage that
affects function.
b. Removal.
Remove screws or twist-type
fasteners and remove, panel, cowling or fairing. Do not
remove riveted panels unless the entire panel is to be
replaced.
(1) Remove shields, if installed, from engine air
inlet ducts in transmission cowling.
c.
(4) Window in transmission cowl for damage
that affects function.
b. Removal.
(2) Remove bolts that attach hinges to fittings
on cowl frame.
Identify washers and shims for
reinstallation in the same relative location to avoid
requirement to align cowling assembly when it is reinstalled.
Repair.
(1) Replace unsatisfactory or missing seals.
Refer to paragraph 2-3.
(3) Remove cowling assembly from helicopter.
(2) Repair holes, cracks and deformation in
accordance with TM 55-1500-204-25/1.
c.
Repair
(1) Repair cracks, dents, holes and deformation
in accordance with TM 55-1500-204-25/1. See figure 231 for list of materials that cowling is constructed from.
(3) Repair corrosion damage as outlined in
paragraph 2-66.
(4) Replace worn, damaged and missing twisttype fasteners.
(2) Repair corrosion damage in accordance
with paragraph 2-68.
d. Installation. Install panels, cowlings and fairings
with screws or twist-type fasteners as applicable.
(3) Remove damaged or worn cowl assembly
hinges and latches and install new parts.
2-15. Transmission Cowl Assemblies (Figure 2-31).
(4) Remove damaged window and replace with
new window. Install new seal with the window if the old
seal is not serviceable.
The transmission is cowled on each side by cowling
assemblies which swing open fore and aft on articulated
hinges. Openings in the transmission cowl assemblies
form the engine air inlet ducts. A small window in the
right hand transmission door permits viewing the
transmission oil level. A safety indicator is located at
each cowling latch. The indicator will protrude slightly
past cowling skin when the latch is properly engaged.
(5) Remove damaged or worn cowling standoff
and replace with new standoff.
Replace standoff
retaining spring clip if the old clip is not serviceable.
d. Installation.
(1) Position cowling on helicopter and install
bolts, washers, shims, and nuts to attach hinges to
fittings on cowl frame. Install the shims in the same
location from which removed. Refer to step b above.
a. Inspection.
(1) Cowl assemblies for cracks, dents, holes,
deformation and corrosion.
(2) Close cowling and check alignment of
cowling and operation of latches, adjust shims on hinges
and/or latch bolt assemblies if required. See figure 255.
(2) Latch assemblies for wear and damage that
affects function.
2-52
TM 55-1520-234-23
screen. Carefully drill out the rivets that retain the
screen. Remove the damaged screen and retainer
strips. Deburr holes and touch up bare metal with
chemical film (C37) and primer (C102). Trim new screen
to fit. Install new screen, retainer strips and spoiler with
rivets (item 47, table 2-2). Touch up paint.
(3) Open cowling and check standoff to ensure
that it operates properly.
2-16.
Fairing.
Engine Cowl Assemblies and Tailpipe
The engine is cowled on each side by cowling
assemblies which swing open fore and aft on articulated
hinges. Each of the cowlings has a small air scoop. A
safety indicator is located at each cowling latch. The
indicator will protrude slightly past cowling skin when the
latch is properly engaged.
d. Installation.
(1) Position cowling on helicopter and install
bolts, washers and nuts to attach hinges to fittings on
cowl frame. Install the washers in the same location
from which removed. Refer to step b above.
a. Inspection. See figure 2-31.
(2) Close cowling and check alignment of
cowling and operation of latches. Adjust washers on
hinges and/or latch bolt assemblies if required. See
figure 2-55.
(1) Cowl assemblies for cracks, dents, holes,
deformation and corrosion.
(2) Latch assemblies for wear and damage that
affects function.
2-17. Tailpipe Assembly Fairing. See figure 2-31.
(3) Cowl assembly hinges for wear and damage
that affects function.
The tailpipe assembly fairing encloses the engine
tailpipe and supports the infrared suppression exhaust
duct.
(4) Screens (9, figure 2-31) for distortion and
other damage that results in significantly enlarged holes.
a. Inspection. See figure 2-31.
b. Removal.
(1) Fairings (5, 7, 8, 11 and 12) and spoiler (6)
for cracks, dents, holes, deformation and corrosion.
(1) Open the cowling assembly that is to be
removed.
(2) Fairing assembly installation turnlock and
machine screw fasteners for wear, damage and secure
installation.
(2) Remove bolts that attach hinges to fittings
on cowl frame. Identify washers for reinstallation in the
same relative location to avoid requirement to align
cowling assembly when it is reinstalled.
(3) Screens (13) for distortion and other
damage that results in significantly enlarged holes.
(4) Spoiler (6) for secure installation.
(3) Remove cowling assembly from helicopter.
c.
b. Removal. See figure 2-31.
Repair.
(1) Remove
infrared
suppression
duct
assembly (10) from aft end of tailpipe assembly fairing.
Refer to Chapter 4.
(1) Repair cracks,
dents, holes and
deformation in accordance with TM 55-1500-204-25/1.
See figure 2-31 for list of materials that cowling is
constructed from.
(2) Disconnect turnlock fasteners and remove
machine screws to release tailpipe assembly fairing
from fuselage and from pylon fairing. Remove tailpipe
assembly fairing.
(2) Repair corrosion damage in accordance
with paragraph 2-66.
(3) Remove damaged or worn cowl assembly
hinges and latches and install new parts.
c.
Repair. See figure 2-31.
(1) Repair cracks, dents, holes and deformation
in accordance with TM
(4) Replace damaged screen (9, figure 2-31).
Carefully drill out rivets and remove the aluminum
alloy spoiler to gain access to the rivets that retain the
2-53
TM 55-1520-234-23
b. Removal.
55-1500-204-25/1. See figure 2-31 for list of materials
that fairing and spoiler are constructed from.
(1) Release latches and open door.
(2) Repair corrosion damage in accordance
with paragraph 2-68.
(2) Remove bolt to separate door-holding
spring (restrainer) at the lower hinge. Remove bolts to
disconnect hinges from hinge supports and remove
door.
(3) Replace damaged and missing turnlock and
machine screw fasteners.
d. Installation. See figure 2-31.
c.
(1) Place tailpipe assembly fairing on fuselage.
Install machine screws, and secure turnlock fasteners.
Repair.
(1) Replace faulty latches.
(2) Replace damaged seals or rebond seals
with adhesive. Refer to paragraph 2-3 for procedure.
(2) Install infrared suppression duct (10). Refer
to Chapter 4.
(3) Replace faulty hinges.
2-18. Infrared Suppression Duct Assembly.
figure 2-31.
See
(4) Repair cracks, dents and holes that are
within limits shown on figures 2-20 through 2-23. Use
repair procedures shown on the illustrations.
Refer to Chapter 4 for inspection, removal, repair
and installation instructions for the infrared suppression
duct assembly.
(5) Repair corrosion damage. Refer to
paragraph 24-68.
2-19. Access Covers and Doors.
(6) Replace door if it is distorted to the degree
that it will not close properly and fit smoothly with the
fuselage.
The access covers and doors are shown on figure 23. The components which are accessible through each
cover and door are listed on the illustration. Refer to
paragraphs 2-20 and 2-21 for instructions on hydraulic
compartment doors and ammunition compartment
doors.
d. Installation.
Position door on fuselage and install bolts to attach
hinges to supports. Attach door-holding spring at the
lower hinge. Open and close door several times to
ensure that latches operate properly.
2-20. Hydraulic Compartment Doors.
The two hydraulic compartment doors (11, figure 23) are constructed of laminated fiberglass edges,
honeycomb core and aluminum skin.
2-21. Ammunition Compartment Doors.
The two ammunition compartment doors (8, figure
2-3) immediately aft of the gun turret give access to the
ammunition compartment. The door hinges are at the
bottom of the doors. The doors hinge open to the
horizontal position and are supported by cables.
Construction is aluminum frame and skin.
a. Inspection. Inspect both doors as follows:
(1) Latches for correct operation.
(2) Seals for cuts, chaffing and secure adhesion
to door surface.
a. Inspection: Inspect both doors as follows:
(3) Hinges for cracks. If cracks are suspected
for any reason, remove hinges and inspect by
fluorescent penetrant method.
(1) Catch assemblies for correct, operation and
damage. Catch assembly covers and strips for damage.
(4) Doors for cracks, dents, holes, deformation
and corrosion.
(2) Door support cables and cable fasteners for
proper safetying and condition.
2-54
TM 55-1520-234-23
(3) Doors for cracks, dents, holes, deformation
and corrosion.
(4) Door hinges for damage.
(5) Door rubber strips (seals) for cuts, chaffing
and secure adhesion to door surface.
Provide adequate ventilation when using
methyl-ethyl-ketone.
Avoid breathing
solvent vapors and avoid prolonged skin
contact.
b. Removal.
(1) Release latches and open door.
(f) Remove overlay patch and doubler
from door. Clean and deburr parts and coat outside
surface with primer (C102). Clean inside surface of
doubler, patch and mating surfaces on door with methylethyl-ketone (C87).
(2) Support door in horizontal position and
disconnect the support cables.
(3) Remove hinge pin from hinge and remove
door.
c.
(g) Apply a thin smooth layer of adhesive
(C17) on mating surfaces of door, patch and doubler.
Position the patch and doubler in the door. Install rivets
(item 51, table 2-2) in holes prepared in step (e). Install
cherry rivets (item 32, table 2-2) in remaining three
sides of patch.
Repair.
(1) Replace faulty catch assemblies, damaged
catch assembly covers, and strips.
(2) Replace damaged door support cables and
cable fasteners (attachment brackets).
(h) Install rubber strip (seal) that was
removed in step (b). Refer to paragraph 2-3 for
procedure.
(3) Repair cracks, dents, holes, deformation
and corrosion. Refer to TM 55-1500-204-25/1 for
general repair instructions. Refer to paragraph 2-68 for
corrosion damage repair instructions. Repair fatiguetype vertical cracks along aft spot weld seam on outer
skin of ammunition compartment doors as follows:
(a)
(i)
Touch up paint to match surrounding
area.
(4) Replace damaged rubber strips (seals) or
rebond with adhesive. Refer to paragraph 2-3 for
procedure.
Stop drill ends of cracks.
(5) Replace faulty hinges.
(b) Remove rubber strip (seal) from inside
aft edge of door.
d. Installation.
(c) Fabricate doubler of 2024T3 aluminum
alloy 0.040 inch thick. Make doubler to fit the width and
length of the inside edge of door.
(1) Align door hinge half with the mating hinge
half on the fuselage and install pin.
(2) Install two door support cables.
(d) Fabricate overlay patch for outside
skin of door of 2024 TO aluminum alloy 0.025 inch
thick. Make overlay patch to overlap cracks in skin by
1½ 2 inches.
(3) Check operation of door catches to ensure
that they function properly.
Check that door fits
smoothly with fuselage and that rubber strips (seals) are
in position on the door.
(e) Clamp inside doubler and overlay
patch to door. Drill holes for rivets (item 51, table 2-2)
through overlay patch, aft edge of door and doubler.
Use one inch spacing between rivets and 1/4 inch edge
distance. Countersink the doubler for installation of
these rivets. Drill holes for bulbed-type cherry rivets
(item 32, table 2-2) on remaining three edges of patch.
Use the same spacing.
2-22. Structural Members.
Three types of structural members are included in
this manual. They are the primary structural caps
(figure 2-4), structural members other than
2-55
TM 55-1520-234-23
(3) Corrosion damage in excess of 0.010 inch
deep after polishing out and/or which affects over twenty
percent of the area of the tube is not acceptable.
Replace tube.
honeycomb panels, and diagonal braces (figure 2-35).
Refer to paragraph 2-23 for maintenance information on
the diagonal brace tubes.
a. Inspection.
(4) Wear in bolt holes in fittings in ends of
diagonal brace assemblies in excess of 0.005 inch is not
acceptable. Make the inspection for bolt hole wear only
if the diagonal brace tubes are removed for some other
purpose.
(1) Inspect the primary structural caps (figure 24) for cracks, corrosion and distortion.
(2) Inspect the other structural members,
except honeycomb panels, shown on figure 2-4 for
cracks, corrosion and distortion.
(5) Distortion of the diagonal brace assemblies
that can be detected visually is not acceptable. Replace
tube.
b. Repair.
b. Removal.
NOTE
(1) Remove right and left access panels (22,
figure 2-3).
Damage that requires jigs and fixtures for
repair must be repaired at Depot Level
Maintenance.
(2) Remove bolts (10 and 18, figure 2-35) and
remove right diagonal brace tube (15).
(1) Repair limits for primary structural caps are
shown on figure 2-4. Repair requires approval by
engineering authority.
(3) Remove bolts (1 and 9, figure 2-35) and
remove left diagonal brace tube (5).
(2) Repair structural members, other than
honeycomb panels and primary structural caps, shown
on figure 2-4 in accordance with TM 55-1500-204-25/1.
c.
Repair.
(1) Polish out nicks, scratches and corrosion
damage that is within limits.
2-23. Diagonal Brace Tube Assemblies.
(2) Touch up repair area with primer (C102).
The diagonal brace tubes are located in the fuselage
adjacent to the wing attachment area.
(3) Replace adjustable rod end connector if bolt
hole is worn beyond 0.005 inch limit. Replace entire
diagonal brace tube assembly if the bolt hole in the fixed
fitting is worn beyond the 0.005 inch limit.
NOTE
Diagonal brace tube assemblies must be
installed prior to flight.
d. Installation.
a. Inspection.
Remove right and left access
panels (22, figure 2-3) and inspect both diagonal brace
tube, assemblies (figure 2-35) for damage in excess of
the following limits:
(1) Position left diagonal brace tube assembly
(5) figure 2-35 in helicopter. Install bolt (1), washers (2
and 3) and nut (4). Adjust rod end connector in opposite
end of left diagonal brace tube assembly if necessary
and install bolt (9), washers (6 and 8) and nut (7).
(1) Dents in excess of 0.010 inch. Smooth
dents up to 0.010 inch deep are acceptable without
polishing out.
(2) Install right diagonal brace tube assembly
(15) in the same manner with the exception that special
bolt (18) is used in the forward end.
(2) Nicks and scratches in excess of 0.010 inch
depth. Nick and scratch damage less than 0.010 inch
deep is acceptable if the damage is polished out.
(3) Install right and left access panels (22,
figure 2-3).
2-56
TM 55-1520-234-23
Figure 2-35. Diagonal brace tube installation
2-57
TM 55-1520-234-23
2-24. Canopy.
The canopy over the pilot and gunners stations
consists of acrylic plastic windows mounted in a
supporting framework of aluminum alloy. The pilot and
gunners doors form a portion of the canopy. The
windows in the pilot and gunners doors and the windows
in the frames opposite the doors are equipped with
linear explosive window cutting assemblies for
emergency removal of the four windows. Engine bleed
air is utilized for rain removal.
Canopy
instructions
damage to
confined to
window.
frame
removal/installation
apply only when there is
the frame.
If damage is
a window, replace only the
(1) Remove pilot and gunners doors. Refer to
paragraph 2-27.
(2) Remove bracket of gunners door strut (4,
figure 2-45) from structure.
2-25. Canopy Assembly Frames.
The aluminum canopy assembly frames support the
acrylic plastic windows and the pilot and gunners doors.
a. Inspection. Canopy frames (2 and 6) and cross
member (3) shown in figure 2-36 shall be inspected.
(1) No cracks are allowed.
(3) Remove all hinge halves from left canopy
frame (2, figure 2-36).
(4) Remove window cutting assembly junction
manifold from bracket on frame. Remove rivets to
detach bracket and interconnect lines from canopy.
Record locations for installation of parts on new canopy
frame.
(5) Remove all screws attaching center window
(1, figure 2-36).
(2) No holes are allowed.
(3) Corrosion is not acceptable, but the frames
and cross members may be cleaned up to make them
acceptable. Do not reduce frame or cross member
thickness by more than 3 percent.
(4) Distortion is acceptable, but must be
negligible enough that fit and function is not adversely
affected.
(6) Insert blade of putty knife between center
window and canopy frame to separate center window
from sealant. Force putty knife around entire mating
surface and remove window.
(7) Drill out rivets attaching left end of cross
members (3, figure 2-36) to left canopy frame (2).
(8) Remove the tees supporting the antenna.
Remove screws attaching left canopy frame to top roof
section.
NOTE
Ensure that both the pilot and gunners
arming/firing mechanism handles are
secured with safety pins prior to entry into
the cockpit area.
Keep all parts removed for use as templates
for installation of new parts on canopy
structure.
b. Removal - Left Canopy Frame
(9) Remove rivets attaching left canopy frame
to structure as follows:
(a) Drill out all rivets attaching left canopy
frame to front of bulkhead. Remove two screws from
inboard side of forward bulkhead which are secured by
rivnuts on inboard side of left canopy.
Ensure that both the pilot and gunners
arming/firing mechanism handles are
secured with safety pins prior to entry into
the cockpit area.
(b) Drill out the two vertical rivets attaching
canopy frame and tab to bulkhead at left rear corner of
rear bulkhead and leave the tab. See
NOTE
2-58
Change 48
TM 55-1520-234-23
figure 2-37. Remove the balance of rivets attaching
frame to rear bulkhead.
(c) Drill out vertical rivets attaching left side
of pilots console to left frame.
(d) Drill out vertical rivets that attach left
frame to tabs of vertical stringers of structure. Drill
these rivets out from below.
that
(e) Drill out all remaining horizontal rivets
attach lower canopy frame to structure.
Change 48
2-58A/(2-58B blank) I
TM 55-1520-234-23
4. Aft Bulkhead (Reference)
5. Forward Bulkhead (Reference)
6. Right Canopy Frame
1. Center Window
2. Left Canopy Frame
3. Cross Member (4 places)
Figure 2-36. Canopy frames and center window
(10)
Carefully separate upper part of left
canopy frame (2, figure 2-36) from upper cross
members, forward roof section, forward bulkhead and
aft bulkhead. Remove left canopy frame from structure.
Avoid distorting the left canopy frame because
dimensions on the old frame can be used to locate rivet
holes on the new frame.
c.
(1) Remove all burrs and rough spots on
helicopter structure where canopy frame is to be
installed. Give special attention to rivet holes.
(2) Position new left frame (2, figure 2-36) on
helicopter.
Installation - Left Canopy Frame
NOTE
Ensure that both the pilot and gunners
arming/firing mechanism handles are
secured with safety pins prior to entry into
the cockpit area.
Change 22
There are no horizontal pilot rivet holes in
the lower section of new canopy frame when
received for installation. The rivnuts for
door hinge attachment are installed. The
rivnuts for attachment of center window are
not installed but are supplied separately.
(3) Attach left canopy frame to structure as
follows: See figures 2-36 and 2-38.
2-59
TM 55-1520-234-23
209030-55
Figure 2-37. Aft bulkhead rivet removal (typical two places)
upper edge of left canopy frame, and back drill
corresponding rivet holes from cross members to left
canopy edge.
NOTE
Use the old left canopy frame and
remaining canopy structure as template for
rivet patterns and attachment points.
(d) Remove left canopy frame from
structure. Drill a 0.221 to 0.226 inch hole at each of two
spots marked on forward end of canopy from forward
bulkhead. At each of two holes drilled in forward canopy
edge, install a rivnut (item 62, table 2-2).
(a) Align the upper edge of left canopy
frame along left buttock line eight. See figure 2-38.
Align upper edge of right canopy frame along right
buttock line eight. Maintain 16 inch distance between
frames, measured from outboard surface. Use clamps
or other suitable method. Clamp upper left edge of
canopy frame to forward and aft bulkheads.
(e) Again place upper edge of left canopy
frame along left buttock line eight, which is eight inches
left of helicopter center line, and again anchor upper
edge of left canopy frame to structure. See figure 2-38.
(b) Use holes existing in forward bulkhead
as template and mark the centers of the two holes on
inboard side of new left canopy frame.
(f) Install two screws through existing holes
in left side of forward bulkhead and secure screws in
rivnuts installed in left canopy frame.
(c) Position left ends of cross members on
2-60
TM 55-1520-234-23
1. Left Canopy Frame
2. Rivnuts - Door Hinge
3. Right Canopy Frame
4. Cross Member Attach Angles
209030-56B
Figure 2-38. Alignment - upper right and left canopy frames
2-61
TM 55-1520-234-23
and align attachment holes in right side of window, with
the corresponding rivnuts in right frame.
(g) Align rivet holes from upper cross
members with corresponding holes drilled in canopy
frame and install rivets of same type, size and
dimension as in original installation, and attach cross
members to left canopy frame.
(b) Install several attachment screws
through attachment holes along right side of window and
thread into corresponding rivnuts in right canopy frame
to prevent movement of window.
(h) Use the rivet holes in structure as
template and, beginning at upper rivet hole over forward
bulkhead, backdrill a corresponding hole of same size in
forward edge of left canopy. Install a rivet of same size,
type and dimension as original.
(c) Use center window as template and
mark attachment hole locations from window to new left
canopy frame and new cross members, if installed, for
rivnut installation. Also mark hole locations for center
window and for roof structure aft of center window.
NOTE
NOTE
Refer to TM 55-1500-204-25/1 for riveting
procedures.
If areas of original canopy frame are too
badly damaged for use as template, see
figure 2-40.
(i) Repeat procedure as outlined in
paragraph above, at each existing rivet hole over
forward bulkhead.
(d) Remove center window. At the marked
locations on left canopy and cross members, if new, drill
and install rivnuts (item 62, table 2-2) at each marked
location for window and roof attachment. See figure 240.
(j) Install rivets at aft end of left canopy
frame to structure bulkhead.
(k) Install two corner rivets to attach clip
and aft end of canopy frame to aft bulkhead. See figure
2-37.
(e) Remove burrs and filings or anything
that might damage window or prevent it from seating.
(l) Back drill lower section of canopy left
frame using longitudinal rivet holes in structure as a
rivet pattern. Install longitudinal rivets of same type,
size and dimension as used in original structure.
(f) Apply sealant (Cl16) to mating surface
of center window and canopy mating surface as follows:
(m) Install rivets attaching left side of pilots
console to left canopy frame.
(n) Install vertical rivets to attach lower side
of left canopy frame to structure at each tab of vertical
stringers.
Provide adequate ventilation when using
methyl-ethyl-ketone.
Avoid breathing
solvent vapors and avoid prolonged skin
contact.
(4) Install previously installed center window as
follows: See figures 2-39, 2-40 and 2-41.
1 Thoroughly clean canopy mating surface
with clean cloth, saturated with methyl-ethyl-ketone
(C87).
Wipe cleaned area dry before cleaner
evaporates.
NOTE
Due to the difference in procedure for
reinstallation of a center window and of a
new, undrilled window, the two procedures
will be discussed separately. Refer to
paragraph 2-26, f if a new center window is
to be installed.
2 Apply masking tape (C134) adjacent to
inboard mating edge of center window to prevent
excessive adhesive from smearing on window beyond
mating area.
(a) Place center window in installed position
2-62
TM 55-1520-234-23
Figure 2-39. Upper center window
2-63
Foldout Figure 2-40. Center window rivnut hole dimensions
(Located in back of manual)
2-64
TM 55-1520-234-23
Figure 2-41. Canopy frame and cross member installation (Sheet 1 of 2)
2-65
TM 55-1520-234-23
Figure 2-41. Canopy frame and cross member installation (Sheet 2 of 2)
2-66
TM 55-1520-234-23
Do not use aliphatic naphtha type I in or
around cockpit. Use of this solvent can
result in damage to acrylic plastic and
window cutting assembly.
3 Clean mating area of window with clean
cloth, saturated with naphtha. Wipe dry before cleaner
evaporates.
4 Apply a thin coat of sealant (C116) to
mating area of window.
(g) Place center window in install position,
with attachment holes in window aligned with
corresponding rivnuts in canopy frame.
(h) Install center window in canopy frame,
with original hardware (short screws). Start at forward
end and work progressively aft.
NOTE
Separate the 12 longer screws which are
used for attachment of four door hinge
halves.
(i) Remove masking tape from inside
window surface..
(5) Place gunners door on canopy in closed
position.
assembly junction manifold bracket that was removed in
step b, (4). Locate the bracket in the same relative
position it was in on the old left canopy frame. Also
ensure that the bracket is located with 0.06 inch
clearance between top of bracket and surface of window
edging and is positioned for good routing of interconnect
lines.
(11)
Install the window
assembly junction manifold on the bracket.
cutting
(12)
Remove cover or cap from
transfer lines and inspect tips for damage. If the tip
surface is marred, replace the transfer line. If the tips
are satisfactory, attach the transfer lines to the manifold.
Ensure that all window cutting assembly transfer lines
for the doors and windows are connected.
(13)
Position door strut attachment
fitting on upright frame member. Secure with two bolts
and thin washers to rivnuts on aft side. Install two
countersunk screws from outboard side and install
washers and nuts at inboard side. Install a support clip
that extends across retainer of window cutting assembly
of fixed window with the upper screw.
(14)
to paragraph 2-27.
Install gunners door strut. Refer
(15)
Check operation of gunners
door latch and adjust if required. Refer to paragraph 230.
d. Removal - Right Canopy Frame
(6) Assemble a blank hinge half on each door
hinge half with a headed pin, washer and cotter pin.
(7) Use a hole finder to mark location for three
holes in hinge to match rivnuts in canopy frame. Drill
0.170 to 0.176 inch diameter holes and countersink to
match attaching screws.
(8) Place a shim (2, figure 2-39) under each
hinge half and install screws to secure hinges to center
window and to left frame.
(9) Install window cutting assembly in the
window that was installed with the left canopy frame.
Refer to Chapter 17 for instructions to install the window
cutting assembly. Also ensure that the window cutting
assembly is positioned properly for connection to the
manifold in the following steps.
(10)
Install
the
window
cutting
Change 22
Ensure that both the pilot and gunners
arming/firing mechanism handles are
secured with safety pins prior to entry into
the cockpit area.
Remove the right canopy frame (6, figure 2-36) by
the same procedure outlined for the left canopy frame in
the preceding step d with the following exceptions:
(1) The right side of the pilots console is not
attached to the right canopy frame; therefore, there are
no rivets to be removed in that area.
(2) Transpose "left" to "right" and "gunners" to
"pilots" throughout procedure.
2-67
TM 55-1520-234-23
e. Installation - Right Canopy Frame
(1) Remove all burrs and rough spots on
helicopter structure where canopy frame is to be
installed. Give special attention to rivet holes.
Ensure that both the pilot and gunners
arming/firing mechanism handles are
secured with safety pins prior to entry into
the cockpit area.
(2) Position new left canopy frame (2, figure 236) and right canopy frame (6) on helicopter and clamp
in position with top edge of canopy frames on buttock
lines 8 as shown on figure 2-38.
Install the right canopy frame (6, figure 2-36) by the
same procedure outlined for the left canopy frame in
preceding step c with the following exceptions:
(3) Position cross members (3, figure 2-36)
between top edges of left and right canopy frames and
attach to frames at same stations as old canopy. Take
dimensions from old canopy parts and see figure 2-41
for station locations.
(1) The right side of the pilots console is not
attached to the right canopy frame; therefore, there are
no rivets to be installed in that area.
(2) Transpose "left" to "right" and "gunners" to
"pilots" throughout procedure.
f. Removal - Both Left Canopy Frame and Right
Canopy Frame
(4) Install rivnuts (item 62, table 2-2), that were
furnished with left and right frame assemblies, in top
sides of canopy frames for installation of center window.
Use old center window and roof section aft of center
window as templates for rivnut locations. See figure 241 for station locations.
(5) Install center window. Refer to paragraph 226.
(6) Install gunners door, window cutting
assembly, and gunners door strut. Refer to steps c,5
through c,15.
Ensure that both the pilot and gunners
arming/firing mechanism handles are
secured with safety pins prior to entry into
the cockpit area.
(7) Install pilots door in same manner described
in preceding step. Transpose "left" to "right" and
"gunners" to "pilots" throughout procedure.
NOTE
2-26. Canopy Assembly Windows.
Do not remove both right and left canopy
frames unless both frames are damaged.
Alignment during installation is more difficult
if both frames are removed.
Ensure that both the pilot and gunners
arming/firing mechanism handles are
secured with safety pins prior to entry into
the cockpit area.
(1) Remove left canopy frame. Refer to step b.
(2) Remove right canopy frame. Refer to step
d.
g. Installation - Both Left Canopy Frame and Right
Canopy Frame
The canopy assembly windows are the upper center
window, fixed windows in the left canopy frame and the
right canopy frame. All the windows are acrylic plastic.
NOTE
The canopy and windows are subjected to
aerodynamic stress loads which apply a
negative or outward pressure to the
Ensure that both the pilots and gunners
arming/firing mechanism handles are
secured with safety pins prior to entry into
the cockpit area.
2-68
Change 22
TM 55-1520-234-23
canopy doors and windows. The windows
transfer the applied stress loads from
window edges to the mating edges of the
canopy frame. See figure 2-42.
enclosed with a 0.50 inch diameter circle after clean-up.
If more than two such damage areas fall within a 3.0
inch circle, replace the panel.
(3) Inspect the canopy windows for crazing.
Crazing is small minute cracks on the surface of the
material. No crazing is acceptable which impairs vision
of crew members.
a. Cleaning.
Do not use aliphatic naphtha Type I in or
around cockpit. Use of this solvent can
result in damage to acrylic, plastic and
window cutting assembly.
NOTE
If crazing has penetrated the plastic sheet,
classify it as cracks and refer to following
step.
(4) Inspect the canopy windows for cracks. No
cracks are allowed.
Do not use compounds that contain any
abrasive material or solutions that contain
chlorinated carbons.
Avoid excessive
scrubbing of plastic panels during washing
operation.
(1) Clean the transparent plastic windows with
cleaning compound (C41) and water. Gently free all
caked mud or dirt with fingers. Do not use sponges or
coarse cloths.
Rinse frequently with water while
removing mud.
(2) Remove any grease or oil that remains on
windows after washing as described in step (1), with
naphtha (C88) and repeat cleaning with cleaning
compound as described in step (1).
(3) Allow surfaces to drip dry.
(4) Polish out minor scratches which may
interfere with pilot or gunners vision.
(5) Apply rain repellant.
b. Application of Chemical Rain Repellant. Refer
to TM55-150)-204-25/1 for cleaning and application
instructions.
c. Inspection.
Inspect canopy windows and
classify damage according to limits stated below.
(1) Scratches up to 0.016 inch deep and not
exceeding 1.0 inch in length allowed if no other damage
occurs within 1.0 inch.
(2) Nicks, chips, and gouges are allowed if not
deeper than 0.050 inch and not larger than can be
Change 29
(5) Inspect for holes in windows. Holes that can
be cleaned up not to exceed 0.75 inches in diameter
can be patched with a tapered plug if there is no other
damage within three inches.
(6) Inspect for delamination or bond separation
between
windows
and
reinforcement
edges.
Delamination is reparable if not over three inches in
length with the following exceptions:
(a) No delamination repairs are allowed
within five inches fore and aft of door hinges. See
figure 2-42.
(b) No delamination repairs are allowed in
the first 27 inches of the upper center 'window. See
figure 2-42.
(7) Inspect for failure at attachment holes in
window and door frame edge members. Attachment
hole failures are reparable if not more than two
adjoining fasteners are involved except that no two
missing fasteners are acceptable within five inches of a
door hinge. See figure 2-42.
d. Repair.
(1) Blend out small scratches that do not
exceed damage limits. Refer to TM55-1500-204-25/1.
Refer
to
TM55-1520-234-23P
for
windshield
maintenance kit and plexiglass repair kit.
(2) Repair delamination damage that is within
limits by injecting adhesive (C34) in delaminated area
with a syringe. If delamination damage exceeds limits,
replace window or door as applicable.
2-69
TM 55-1520-234-23
Figure 2-42. Stress loads - canopy windows
2-70
TM 55-1520-234-23
pilots window.
(3) Repair attachment hole failures that are
within limits as shown on figure 2-43.
e. Removal.
NOTE
Protect window cutting assembly leadcovered linear explosive charge and backup cushion from damage when exposed.
Refer to paragraph 2-27 for instructions to
remove windows from pilot and gunners
doors.
(b) Carefully loosen window from sealant
around mating surface of canopy, frame and explosive
window cutting assembly with a putty knife or other
suitable tool. Remove window.
(1) Upper center window removal.
(a) Remove gunner and pilots doors. Refer
to paragraph 2-27.
(c) Smooth up and clean canopy window.
Remove old sealant with a sharp plastic scraper and
sandpaper (C112) of various grits. Touch up bare metal
with primer (C102).
(b) Disconnect leads from each antenna
terminal (6, figure 2-39) mounted on inside surface of
center window. Remove antenna.
f.
(c) Remove three hinge halves (1, figure 239) with shims (2) and screws (4).
Installation.
NOTE
(d) Remove remaining screws attaching
center window to canopy frames.
Refer to paragraph 2-29 for instructions to
install windows in pilot and gunners doors.
(e) Apply outward pressure on center
Window and insert a putty knife or other suitable tool
between window surface and canopy frame.
(1) Center window installation.
(a) Make sure that window mating surface
of canopy frame is clean and smooth to prevent damage
to window.
(f) Free window from frame by sliding
blade of tool around window edge to break adhesive
seal.
(b) Place upper center window in position
and check fit.
(g) Remove upper center window.
NOTE
(h) Smooth up and clean canopy mating
surface. Remove old sealant with a sharp plastic
scraper and sandpaper (C112) of various grits. Touch up
bare metal with primer (C102).
Maximum clearance between edges of
window and adjacent structure is to be no
more than 0.030 inch.
Replacement
windows are purposely oversize and usually
require trimming.
(2) Fixed Side Window Removal.
NOTE
(c) Trim edges of window to fit canopy
frame.
These instructions are for either of two fixed
canopy windows, located at right side of
gunners and at left side of pilots station.
(d) Place window again in position. Using a
hole finder, drill holes around window edge to match
existing rivnuts in canopy frame. Make all holes 0.184
to 0.190 inch diameter, except nine hinge attachment
holes. Make hinge attachment holes 0.170 to 0.176 inch
diameter.
(a) Remove screws attaching window
edges to canopy frame. This will include screw passing
through window frame and window cutting assembly
flange and secured by nuts and washers at lower aft
corner of gunners window and at lower forward corner of
2-71
TM 55-1520-234-23
Figure 2-43. Insert repair to edge of windows (Sheet 1 of 2)
2-72
TM 5 15620-234-23
NOTES
4. Fabricate two plates from aluminum or stainless steel n shown on View B. Use aluminum
for repair of canopy side windows. Use stainless steel for repair of upper center window
and for door windows.
5. Fabricate insert as shown on View B.
6. Fit plates and insert as shown on View C.
209030-45-2B
Figure 2-43. Insert repair to edge of windows (Sheet 2 of 2)
2-73
(i) Tighten screws evenly. Remove sealant
squeeze-out and masking material
CAUTION
Use caution to avoid making countersink
for screw heads too deep.
(e) Countersink drilled holes, except at
hinge locations, with 100 degree countersink.
Remove window, deburr holes and round all window
edges.
(j) Install heat shield as follows:
1 Cut shield material 5x13.25 inches,
from asbestos sheet (C66A) and fit to front end
of center window. Trim to fit contour of nose
section where it joins windows.
2 Mask window above heat shield.
(f) Prepare mating surfaces of window and
canopy as follows:
3 Remove heat shield. Brush a coating
of adhesive (C12A) on exposed window surface
below masking.
4 Fit heat shield over prepared area
and roll out all air bubbles with roller. Remove
excess adhesive and masking, and allow to cure.
Provide adequate ventilation when
using methyl-ethyl-ketone. Avoid
breathing solvent vapors and avoid
prolonged skin contact.
(k) Install antenna and connect antenna
terminal (6, figure 2-39).
1 Thoroughly clean canopy mating
surface with a clean cloth saturated with
methyl-ethyl-ketone (C87). Wipe cleaned area dry
with clean cloth before solvent evaporates.
2 Mask inside surface of window adjacent to mating
area.
CAUTION
Do not use aliphatic naphtha, Type I, in
or around cockpit. Use of this solvent can
result in damage to acrylic plastic and
window cutting assembly.
3 Clean mating surface of window with
naphtha (C88). Wipe dry with clean cloth before
cleaner evaporates:
(l) Install gunner and pilots doors. Refer to
paragraph 2-27.
(2) Fixed side windows installation.
(a) Make sure that window mating surface
of canopy frame is clean and smooth to prevent
damage to window.
(b) Position new window in canopy
frame. Maintain equal clearances between window and frame.
(c) Locate and drill holes for window attaching screws.
1 At middle of contour on forward and
aft edge, drill 0.184 to 0.190 inch diameter holes
through window to match rivnuts in
frame. Install screws as required.
4 Apply thin coat of sealant (C116) to
mating surface of window. Position window in
canopy frame.
2 Continue match-drilling holes and installing screws, working in both directions from
attached points, until all holes matching rivnuts
are drilled.
(g) Align drilled holes in window with
rivnuts in canopy frame and install screws
fingertight. Work from middle of window toward
each end.
(h) Install three hinges halves (1), each with
a shim (2) and three screws (4) (longer than other
,window screws).
Change 54
2-74
TM 55-1520-234-23
Figure 2-43A deleted.
3 Also drill 0.170 to 0.176 inch holes to
match existing holes through explosive window
cutting assembly flange and canopy frame at
lower corner of window adjacent to crew door or
forward lower corner of pilot window, or aft lower
corner of gunner window.
4 Mark trim line on edges of window to fit frame.
5 Remove window from frame. Trim,
smooth and round all edges.
6 Countersink holes 100 degrees on outer
surface of window. Deburr holes on inner surface.
Change 22
7 Prepare mating surfaces of window and
canopy frame:
Provide adequate ventilation when
using methyl-ethyl-ketone. Avoid
breathing solvent vapors and avoid
prolonged skin contact. Do not allow
methyl-ethyl-ketone to contact window
cutting assemblies.
2-74A/(2-74B blank)
TM 566-1520-234-23
prolonged skin contact. Do not allow
methyl-ethyl-ketone (C87) to contact
window cutting assemblies.
a. Removal.
CAUTION
(a) Clean mating surface of canopy
frame with methyl-ethyl-ketone (C87) and wipe dry
with clean cloth before solvent evaporates.
Removal of pilot or gunner door requires two persons to hold the door.
NOTE
(b) Mask inside surface of window
adjacent to mating area with masking tape (C134).
Removal procedure for pilot door is
given. The procedure for gunner door
is similar. See figure 2-45 for view of
gunner door.
(c) Clean mating surfaces with
methyl-ethyl-ketone (C87) and wipe dry with clean
cloth before solvent evaporates.
(d) Apply a bead of sealant (C116) on
mating surface of window.
(d) Position window in canopy frame and
install screws finger-tight.
(e) Tighten screws on forward and aft edges
of window, working from middle of contour in both
directions. Tighten screws on top and bottom edge in
same manner. On screws that pass through frame
and are secured by nut and washer apply tightening
force to screws only.
(f) Wipe off excess sealant around window.
Remove masking material.
2-27. Door Assemblies, Pilot and Gunners .
Maximum permissible standoff
distance between Explosive Cord (LES)
and window is 0.10 inch, without repair.
The pilot and gunner door assemblies consist
of formed acrylic plastic mounted in supporting
framework of aluminum alloy tubes. The doors
are incorporated in the canopy. The forward left
side panel is the gunner door. The aft right side
panel is the pilot door. Both doors are equipped
with linear explosive window cutting assemblies
connected to arming/firing mechanisms at pilot
and gunner station for emergency removal of
canopy windows.
(1) Open pilot door. See figure 2-44.
(2) Disconnect flexible interconnecting line of
window cutting assembly from adapter of the
window cutting assembly mounted in the door.
Install protective steel plug and cap to protect the
window cutting assembly connections from damage.
Refer to Chapter 17 for more information on the
window cutting assembly.
(3) Disconnect cables from door struts (1 and 2,
figure 2-44). See figure 2-46 for detail view of strut
installation. Loosen set screw and remove collar
from cable tip. Loosen screw in cable positioning
clamp and slide cable assembly out of cable
positioning clamp. Retain collar, cable positioning
clamp and horseshoe-shaped clamp for
reinstallation.
(4) Detach door strut (1, figure 2-44) from
lower fitting by removing cotter pin, washer and
headed pin. Detach door strut (2) in the same
manner.
(5) If door struts are to be removed from
helicopter, proceed as follows:
(a) Remove cotter pins, washers, headed
pins, shims, and pivot pin shown in figure 2-46,
detail A. Remove forward door strut (1, figure 2-44).
(b) Remove aft door strut (2, figure 2-44) in
the same manner with the exception that a screw is
used in lieu of the pivot pin arrangement used on the
forward door strut (1).
(6) Support door and remove cotter pins,
washers, and headed pins from two hinges (3, figure
2-44). Remove door.
(7) Remove gunners door in a similar manner.
See figure 2-45.
Change 38
2-75
TM 55-1520-234-23
Figure 2-44. Pilots door (Sheet 1 of 2)
2-76
TM 55-1520-234-23
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
Strut (Forward)
Strut (Aft)
Hinge (2)
Shim (2)
Handle (Outboard)
(Flush-type)
Latch Handle (Inboard)
Roller Assembly (3)
Bellcrank (Aft Roller Rods)
Latch Assembly
Guide Pin
Sealant
Plug Button
Access Cover
Window Cutting Assy.
Spring
Rod
Cable Assembly
Cable Assembly
Handle (Outboard)
(Non-flush-type)
209030-28-2D
Figure 2-44. Pilots door (Sheet 2 of 2)
(3) Install aft door strut (2, figure 2-44) on
pilots door with screw nut and washer.
b. Installation.
NOTE
Installation procedure for pilots door is
given. The procedure for the gunners door
is similar. See figure 2-45 for view of
gunners door.
(1) Identify door struts (1 and 2, figure 2-44)
for installation in the correct position by measuring
distance between centers of rod end bearings with
the strut in minimum length position. Dimensions
are as follows:
Strut (1, figure 2-44) ...................... 15.00 inches
Strut (2, figure 2-44) ...................... 11.84 inches
Strut (4, figure 2-45) ...................... 19.50 inches
(2) Install forward door strut (1, figure 2-44)
on pilots door. Attach strut to door with pivot pin,
shims, headed pins, washers and cotter pins as
shown in figure 2-46, detail A. Shim to 0.005 - 0.015
inch dimension as illustrated.
(4) Position pilots door on canopy. Line up two
hinges (3, figure 2-44) and install headed pins,
washers and cotter pins.
(5) Hold pilots door in open position and align
door strut (1, figure 2-44) in lower fitting. Install
headed pin, washer, and cotter pin. Install door strut
(2) in same manner.
(6) Assemble cable assembly cable positioning
clamp, horseshoe-shaped clamp, sleeves, and collar
on strut (1, figure 2-44). See detail view B on figure
2-46. Set cable positioning clamp to dimension
shown and tighten screw to hold cable positioning
clamp in position on door strut and to hold the cable
assembly in position. Position collar on cable tip and
tighten set screw in collar. Adjust collar position on
cable tip as necessary so that collar will actuate
sleeve to release internal locking mechanism in
strut when door handle is turned to unlatch door.
(7) Assemble and adjust cable assembly on
strut (2, figure 2-44) in the same manner outlined in
the preceding step.
NOTE
The procedure for installation of gunners
door strut (4, figure 2-45) is the same as
for the forward strut on the pilots door
given in the preceding paragraph.
(8) Adjust roller ends (7, figure 2-44) on latch
rods to obtain proper engagement in closed position
tighten locknut after adjustment.
(9) Adjust shims under striker plates and/or
move striker plates on serrated plates to obtain
2-77
TM 55-1520-234-23
Figure 2-45. Gunners door
2-78
TM 55-1520-234-23
Figure 2-46. Strut Installation - pilot and gunners doors
secure locking action with roller ends (7, figure 2-44) and
latch assembly (9).
inch. If necessary, make additional adjustments at hinge
shims, roller ends and strikers to obtain proper closing.
NOTE
The maximum contour mismatch of doors to
the adjacent surfaces shall be 0.12 along forward edges and 0.25 along all other edges.
The maximum gap between door frames and
adjacent structure to be 0.20.
(10) Close door and check for gaps between door
and adjacent surfaces. Maximum allowable gap is 0.20
Change 42
(11) Check door struts to ensure that they both lock
to hold the door open at three positions.
(12) Remove protective cap and plug at window cutting assembly interconnect line and adapter of window
cutting assembly in door. Attach the flexible interconnect line to the adapter and TORQUE TO 30-45 INCHPOUNDS. Secure with lockwire (C151).
2-79
TM 55-1520-234-23
(1) Remove pilots door from helicopter. Refer
to paragraph 2-27.
2-28. Door Assembly Frame, Pilot and Gunners.
(2) Remove hinges (3, figure 2-44) and shims
(4). Identify hinges and shims for reinstallation in
same location.
Ensure that both pilot and gunners
arming/firing mechanism handles are
secured with safety pins prior to entry
into the cockpit area.
(3) Remove screws around outside door handle
(5, figure 2-44).
The door frames are constructed from aluminum
alloy extrusions which are assembled by welding.
Gussets and doublers are used to increase strength.
The frames support the acrylic plastic windows,
window cutting assemblies, handles and latches.
(4) Remove screws around outside side of door
frame that retain window in frame.
CAUTION
Use caution to avoid damage to window
cutting assembly during following step.
a. Inspection. Inspect the door frames for
cracks, corrosion and distortion.
(5) Use a putty knife or similar tool to
separate window from frame and window cutting
assembly. Remove window from frame.
b. Repair.
(1) Repair small cracks in the sheetmetal
portion of the frame by stop drilling.
(6) Remove gunners door window in same
manner outlined above for pilots door window.
(2) Polish out minor corrosion damage and
touch up paint to match surrounding area.
e. Installation. Install door windows as follows:
(3) Replace the door assembly if the frame is
distorted to the degree that it affects opening and
closing the door.
CAUTION
Install canopy windows at 70°F or above.
The windows should be installed with
door frame installed in closed position on
aircraft. When possible, only one canopy
glass should be removed from canopy
structure when replacing glass. The
remainder should remain in place to
shape canopy.
2-29. Door Assembly Windows, Pilot and Gunners.
Ensure that both pilot and gunners
arming/firing mechanism handles are
secured with safety pins prior to entry
into the cockpit area.
(1) Remove all old sealant from door structure
with sharp plastic scrapers and sandpaper (Cl112).
Touch up bare metal with primer (C102).
a. Cleaning. Clean the door windows in the
same manner outlined for the fixed windows. Refer
to paragraph 2-26.
(2) Position new window in door frame. Trim
window edges if necessary.
b. Inspection. Inspect the door windows in the
same manner outlined for the fixed windows. Refer
to paragraph 2-26.
(3) Place door frame and window on
helicopter. Locate and drill holes for attaching screw
around outside edge of windows as follows: use a
hole finder to locate holes. At middle of contour on
forward and aft edges, drill 0.184 to 0.190 inch
diameter holes to match rivnuts in door frame.
Install setup screws. Continue match-drilling and
c. Repair. Repair the door windows in the same
manner outlined for the fixed windows. Refer to
paragraph 2-26.
d. Removal. Remove door windows as follows.
2-80
TM 55-1520-234-23
install screws as needed, work both directions from
attached points but omit hinge locations.
(15) Remove residual sealant and masking tape.
(16) Attach flexible connection to window
cutting assembly. Refer to Chapter 17.
(4) At each of two hinge locations, drill three
0.170 to 0.176 inch diameter holes to match
ivnuts in door frame. Mark cut-out for latch (9,
figure 2-44). Remove window.
(17) Attach door struts to door.
(18) Perform functional check of door handle
latch and door struts.
(5) Trim cut-out for latch.
(6) Countersink all holes, except at two hinge
locations, to 100 degrees on outer surface of window.
(19) Install gunners door window and
functional check door handle latch and door strut in
same manner outlined above for pilots door window.
2-30. Door Handles and Latches, Pilot and Gunners.
Provide adequate ventilation when using
methyl-ethyl-ketone (C87). Avoid
breathing solvent vapors and avoid
prolonged skin contact.
Ensure that both pilot and gunners
arming/firing mechanism handles are
secured with safety pins prior to entry
into cockpit area.
(7) Clean mating surfaces of door with
methyl-ethyl-ketone (C87) and wipe dry with clean
cloth before solvent evaporates.
a. Inspection. Inspect handles and latches (figure
2-47), roller assemblies (7, figure 2-44) and
associated rods and bellcranks for the following
defects: Also inspect the corresponding parts on the
gunners door.
(8) Mask inside surface of window adjacent to
mating areas.
CAUTION
Do not use aliphatic naphtha, Type I, in
or around cockpit. Use of this solvent can
result in damage to acrylic plastic and
window cutting assembly.
(1) Cracks. No cracked parts are acceptable.
(2) Wear severe enough to affect function.
(3) Corrosion severe enough to affect function.
(9) Clean mating surface of window with
naphtha (C88) and wipe dry with clean cloth
before solvent evaporates.
(4) Binding when door handles are moved to
open or close door.
(10) Apply sealant (C116) on mating surface of
window.
b. Removal.
NOTE
(11) Position door handle (5 or 19, figure 2-44)
on door.
Removal procedure for pilots door handle
and latches is given. Procedure for
gunners door is similar.
(12) Position window on door and install
screws fingertight in all countersunk holes.
NOTE
(13) Install two hinges (3, figure 2-44) and
shims (4) with sealant (C116) between shim and
window. Align hinge halves and install hinge pins.
Removal procedure for flush-type door
handle is given. Procedure for
non-flush-type is similar. See figure 2-47.
(14) Tighten window attaching screws and
hinge screws evenly.
(1) Remove screws around edge of outboard
handle plate (2, figure 2-47).
Change 2
2-81
TM 556-1620-234-23
Figure 2-47. Door handles - pilot and gunners doors
(2) Loosen screw (9, figure 2-47) and remove
inboard handle (8). Retain shims (7) for
reinstallation in same location.
(3) Remove access cover (12, figure 2-47).
(4) Remove three roller assemblies (7, figure 2-44).
(5) Remove the pins that attach rods and
cables to latch assembly (9, figure 2-44).
(6) Loosen set screw (11, figure 2-47) and
remove latch assembly (4) from shaft (10).
(7) Remove outboard handle (1, figure 2-47)
and handle plate (2) from door.
2-82
TM 556-1620-234-23
(8) Remove pin attaching rod assembly to
bellcrank (8, figure 2-44).
mechanism operation is not satisfactory, adjust
roller assemblies and strikers. Refer to paragraph 2-27.
(1) Replace any part of the door handles or
latch mechanism that is cracked.
(2) Check operation of door struts to ensure
that they will hold the door open at three positions.
If operation is not satisfactory, adjust as required.
Refer to paragraph 2-27.
(2) Replace door handle or latch mechanism
parts that are worn enough to affect function.
2-31. Door Hinges - Pilots and Gunners Doors.
c. Repair.
(3) Polish out minor corrosion damage and
touch up with primer (C102). Replace door handle
assembly or latch mechanism parts that have
corrosion damage severe enough to affect function.
Ensure that both pilot and gunners
arming/firing mechanism handles are
secured with safety pins prior to entry
into cockpit area.
(4) Replace door handle assembly or latch
mechanism parts if binding cannot be corrected by
adjustment.
d. Installation.
NOTE
(1) Install two roller assemblies (7, figure
2-44) and rod assembly (16) on bellcrank (8) with
pins and cotter pins.
Removal and installation instructions are
given for pilots door hinges. Procedure for
gunners door hinges is similar.
(2) Position outboard handle (1, figure 2-47)
and handle plate (2) on door.
(3) Position latch assembly (4, figure 2-47) on
shaft (10) and tighten set screw (11).
a. Inspection. Inspect door hinges (3, figure 2-44)
and corresponding hinges on gunners door for the
following defects.
(4) Attach rod (16, figure 2-44), cable assembly
(17), cable assembly (18), and forward roller
assembly (7) to latch assembly (9) with pins and
cotter pins.
(1) Cracks. No cracks are acceptable.
(2) Wear severe enough to affect function.
(3) Corrosion severe enough to affect function.
(5) Install access cover (12, figure 2-47).
(4) Binding when door is opened or closed.
(6) Install escutcheon plate (5, figure 2-47),
screws (6), shims (7), inboard handle (8), and
setscrew (9).
NOTE
Ensure that screws (6, figure 2-47) that
secure escutcheon plate are self-locking
setscrews P/N NAS 1189-06P6L.
(7) Install screws around outboard handle (13)
to plate (2, figure 2-47).
b. Removal.
(1) Remove cotter pins, washers, and headed
pins to separate halves of hinge halves (3, figure
2-44).
(2) Remove screws, hinge halves (3, figure
2-44) and shims (4). Identify parts for reinstallation
in the same location.
e. Functional check.
(1) Close the door and check operation of door
handles and latches. Also check for gap in excess of
0.020 inch between door and adjacent surfaces. If
excessive gap exists and/or handle and latch
c. Repair.
(1) Replace hinges if cracked.
(2) Replace hinges if worn enough to affect
functions.
2-83
TM 556-1620-234-23
(3) Polish out minor corrosion damage and
touch up with primer (C102). Replace hinges that
have corrosion damage severe enough to affect
function.
and check the struts to ensure they operate freely
and will lock at all three positions.
b. Removal. Remove the pilot and gunners door
struts. Refer to paragraph 2-27.
(4) Replace hinges that were found to be
binding during inspection.
d. Installation.
c. Disassembly.
(1) Remove snap ring (2, figure 2-48) from
housing (6).
(1) Position shims (4, figure 2-44) and hinges
(3) on door and canopy frame and install screws.
(2) Install headed pins, washers, and cotter
pins to complete installation of hinges.
(2) Extend rod (1) until the six balls (4) seat in
rod recesses and slide sleeve (3) off end of housing
(6).
NOTE
e. Functional check.
(1) Check hinges for binding when door is
opened and closed.
Be careful that none of the six balls are lost.
(3) Remove balls (4) and spring (5) from assembly.
(2) Check door for proper operation of door
handles and latches. Also check for gap in excess of
0.020 inch between door and surrounding surfaces.
If excessive gap exists and/or handle and latch
mechanism operation is not satisfactory, adjust
shims under hinges. Refer to paragraph 2-27.
2-32. Strut Assemblies - Pilot and Gunners
Doors. P/N 209-030-640-1, -3, and -5.
(4) Remove rod assembly (1) from housing (6).
(5) Loosen jam nut and remove rod end
bearing (10).
(6) Remove pin (8) from retainer (7) and
remove rod end bearing (9).
d. Inspection.
NOTE
Pilot and gunners door struts P/N
209-030-687-1/ -3 and -5 are not
reparable. The disassembly, inspection
and repair instructions in this paragraph
apply to struts P/N 209-030-640-1/ -3 and -5.
The door strut assemblies serve to hold the pilot
and gunners doors open at three positions from
partially open to full open. When balls (4, figure
2-48) drop in detents in rod assembly (1) the strut
locks in position. When the door handle is turned,
cable assemblies (17 and 18, figure 2-44) are
actuated. The cables move sleeves, shown on figure
2-46, which allows the strut to unlock and permits
the door to be opened or closed. The pilots door has
two struts (1 and 2, figure 2-44). The gunners door
has one strut (4, figure 2-45).
a. Inspection. Inspect pilot and gunners door
struts for damage visually. Open and close the doors
(1) Inspect rod assembly (1, figure 2-48) for
roughness and/or wear that affects function of strut,
cracks, and corrosion.
(2) Inspect snap ring (2, figure 2-48) for
defects.
(3) Inspect sleeve (3, figure 2-48) for cracks,
wear, corrosion and nicks.
(4) Inspect six balls (4, figure 2-48) for
deformity, wear, and corrosion.
(5) Inspect spring (5, figure 2-48) for cracks,
corrosion and loss of tension.
(6) Inspect housing (6, figure 2-48) for cracks,
corrosion and thread damage. Inspect solid film
lubricant (11) on housing for continuous unbroken
film of serviceable lubricant.
2-84
TM 55-1520-234-23
Figure 2-48. Door strut assembly PIN 209-030-640-1, -3, and -5
2-85
TM 556-1620-234-23
(7) Inspect retainer (7, figure 2-48) for cracks,
corrosion and thread damage.
(b) Polish out minor corrosion damage with
fine crocus cloth (C45).
(6) Repair housing (6, figure 2-48) as follows:
(8) Inspect pin (8, figure 2-48) for cracks,
corrosion and deformity.
(9) Inspect rod end bearings (9 and 10, figure
2-48) for cracks, roughness or binding when bearing
is moved by hand, corrosion and thread damage.
e. Repair.
(a) Replace housing if cracked or if it failed
corrosion or thread damage inspection.
(b) Clean up minor thread damage.
(c) Polish out minor damage with fine crocus
cloth (C45).
(1) Repair rod assembly (1, figure 2-48) as follows:
(a) Replace rod assembly that failed
inspection for roughness. Polish out minor
roughness with fine crocus cloth (C44).
(b) Replace rod assembly if cracked.
(c) Replace rod assembly that failed
inspection for corrosion. Polish out minor corrosion
damage with fine crocus cloth.
(2) Repair snap ring (2, figure 2-48) as follows:
(a) Replace snap ring that failed inspection
for corrosion, resiliancy or loss of tension.
(b) Polish out minor corrosion with fine
crocus cloth and touch up with primer (C102).
(3) Repair sleeve (3, figure 2-48) as follows:
(a) Replace sleeve if cracked.
(b) Replace sleeve that failed inspection for wear.
(c) Replace sleeve that failed inspection for
corrosion. Polish out minor corrosion with fine
crocus cloth and touch up with primer (C102).
(d) Apply new solid film lubricant if old
lubricant failed inspection. (AVIM)
1 Clean the housing with clean cheese
cloth (C36) dampened with naphtha (C88), dry
before naphtha evaporates with clean cheese cloth.
2 Mask off external surface that is not to
be coated with lubricant.
3 Apply lubricant (C85) to the area shown
on figure 2-48 to a depth of approximately 0.002
inch.
4 Remove masking material and cure
housing at 275°F for sixty minutes in a
recirculating-type, automatically controlled oven.
5 Test the solid film lubricant for
adequate adhesion. Apply a strip of adhesive tape
with firm finger pressure, and then remove the tape
with one abrupt motion. If large particles or flakes of
lubricant peel off with the tape, reapply the
lubricant and repeat the test.
(7) Repair retainer (7, figure 2-48) as follows:
(a) Replace retainer if cracked or if it failed
corrosion or thread damage inspection.
(4) Repair balls (4, figure 2-48) as follows:
(a) Replace balls that failed deformity,
wear, or corrosion inspections.
(b) Polish out minor corrosion damage with
fine crocus cloth (C45).
(c) Clean up minor thread damage.
(b) Polish out minor corrosion with fine
crocus cloth (C45).
(5) Repair spring (5, figure 2-48) as follows:
(a) Replace spring if cracked or if it failed
resiliance or corrosion inspection.
(8) Repair pin (8, figure 2-48) as follows:
(a) Replace pin if cracked or if it failed
corrosion or wear inspection.
(b) Polish out minor corrosion damage with
fine crocus cloth (C45).
2-86
TM 556-1620-234-23
(9) Repair rod end bearing (9 and 10, figure
2-48) as follows:
(a) Replace rod end bearing if cracked or if it
failed thread damage, roughness or corrosion
inspection.
a. Inspection. Inspect door latch strikers,
serrated plates and attaching parts that are
contacted by roller assemblies (7, figure 2-44) and (6,
figure 2-45) for the defects listed below. Also inspect
the retainers located forward and aft of the gunners
door latch (9, figure 2-45) for the same defects.
(b) Clean up minor thread damage.
(c) Polish out minor corrosion damage with
fine crocus cloth (C45).
f. Assembly.
(1) Install rod end bearing (9, figure 2-48) in
retainer (7) and secure with pin (8).
(1) Cracks. No cracked parts are acceptable.
(2) Corrosion damage severe enough to affect
function.
(3) Loose and missing screws in latch strikers,
retainers, and plate.
(4) Wear in door latches, strikers, retainers
and plates severe enough to affect function.
(2) Thread nut on rod end bearing (10) and
install bearing in rod assembly (1).
(3) Install retainer (7) on housing (6).
(4) Place spring (5) on housing. Slide sleeve (3)
on housing until ball recess of sleeve extends
slightly outboard from end of housing and install six
balls (4) into sleeve recess, then slide sleeve further
onto housing.
(5) Install rod assembly (1) in housing (6) to
full retract position.
b. Removal.
(1) Remove screws from latch strikers.
Remove strikers, shims and associated parts.
Identify parts for reinstallation in same location.
(2) Remove screws from retainers on canopy
sill of gunners door. Remove the retainers and
identify for reinstallation in same location.
(3) Remove screws from center striker plate.
Remove striker plate, shims and serrated plate.
Identify parts for reinstallation in same location.
(6) Adjust rod end bearing (10) to obtain strut
length dimension illustrated and tighten nut on rod
end bearing (10) to hold it in position.
c. Repair.
g. Installation. Refer to paragraph 2-27.
h. Functional Check. Open door and check to
ensure that struts will lock at all three positions.
2-33. Striker Assembly, Pilot and Gunners Doors.
When closing the gunners door, the forward roller
assembly (6, figure 2-45) and aft roller assembly (6,
figure 2-45) extends and makes contact with forward
and aft door latch strikers. The pilots door has one
forward roller assembly (7, figure 2-44) and two aft
roller assemblies (7, figure 2-44) that extend and
make contact with one forward door latch striker
and two aft door latch strikers. The latch assembly
(9, figure 2-44) and (9, figure 2-45) goes through the
center striker plate on canopy sill and locks the door
securely. The gunners door has two retainers located
forward and aft of center striker plate on canopy sill.
The retainers line up with door and go into slotted
recesses in gunners door.
(1) Replace any parts that are cracked.
(2) Polish out corrosion damage with crocus
cloth (C45) and paint with primer (C102). If
corrosion damage is excessive and severe enough to
affect function of parts, replace affected parts.
(3) Replace loose and missing screws in door
striker latches, retainers, and serrated plates.
(4) Replace worn or defective door strikers,
serrated plates and retainers.
d. Installation.
(1) Position door latch strikers on canopy
frame. Secure door latch strikers with screws.
(2) Position retainer on canopy sill. Secure
retainer with screws and washers.
2-87
TM 556-1620-234-23
(2) Breaks in line. No breaks are acceptable.
(3) Position center striker plate, shims and
serrated plate on canopy sill. Secure with screws,
washers, and nuts.
(3) Pinched line. No pinched lines are
acceptable if area of line is reduced significantly.
e. Functional Check. Open and close the pilots
,and gunners door and check operation of door
latches. Also check for gap in excess of 0.020
between door and adjacent surfaces. If excessive gap
exists and/or if door latch operation is not
satisfactory, adjust roller assemblies and strikers.
Refer to paragraph 2-27 to check the operation of
door struts.
2-33A. Vent and Drain Locations.
Fuel, oil, hydraulic, and battery installations
are vented and drained through a series of
common and individual lines. Most lines exit the
underside of the fuselage, and are identified by
decals adjacent to the exit point. Typical locations
of vents and drains are shown in figure 2-48A.
(4) Chafing. Determine cause of chafing and
extent of damage. Parts with severe chafing damage
are not acceptable. Refer to TM 55-1500-204-25/1.
(5) Loose or missing support clamps.
b. Removal. Remove the B-nut of fuel vent drain
line at nearest connection or fitting. Remove screws
from clamps. Retain screws, nut and washers for
reinstallation.
c. Repair.
(1) Clean clogged vent tube with lockwire (C152).
(2) Replace lines with breaks, cracks,
severely pinched areas or severely chafed areas.
2-34. Vent and Drain Installation - Fuel.
Vent and drain installations for the fuel system
are shown in figure 2-48A.
The engine fuel filter assembly, mounted on a
bracket located on left side of engine compartment,
has an overboard drain line (8) with a manually
operated poppet cock drain valve in corporated into
the line.
The fuel control and overspeed governor (power
turbine governor), located on left side of engine,
each have a vent line that drains into a common
overboard drain line (8).
(3) Correct conditions that cause chafing
damage.
d. Installation. Connect the B-nut of drain line
tube to fittings or connections. Position clamps on
the tubes to structure and secure with screws,
washers, and nuts.
2-35. Vent and Drain Line - Engine Oil Tank.
The engine oil tank scupper, oil tank, oil cooler,
and transmission drain into common lines (9 and
11, figure 2-48A) which carry oil to lower surface of
helicopter. A vent line from the top of the oil tank
extends to the cowling directly above the oil tank.
The flow divider pump valve and the
combustion chamber drain valve, located on the
lower side of the combustion chamber housing,
drain into a common overboard drain line (8).
The ejector (tailpipe cooling assembly), located
at the aft end of the engine, has an overboard drain
line (8).
a. Inspection. Inspect the engine fuel filter
assembly, fuel control and overspeed governor, flow
divider and dump valve assembly, combustion
chamber drain valve, and ejector (tailpipe cooling
assembly) drain lines for the following defects:
a. Inspect for the following defects:
(1) Clogging.
(2) Breaks in line - no breaks are acceptable.
(3) Pinched line - no pinched lines are
acceptable if area of line is reduced significantly.
(4) Chafing. Determine cause of chafing
and extent of damage. Parts with severe chafing
damage are not acceptable.
(5) Loose or missing supporting clamps.
(1) Clogging.
Change 22
2-88
TM 5-1 520-234-23
Figure 2-48A. Vent and drain locations
b. Removal. Remove the B-nut of tube at nearest
fitting or connection. Remove screws from clamps.
Retain screws, washers and nut for reinstallation.
(2) Replace lines with breaks, cracks, severely
pinched areas or severely chafed areas.
(3) Correct conditions that caused chafing
damage.
c. Repair.
(1) Clean clogged vent tube with lockwire (C152).
Change 44
d. Installation. Connect the B-nut of drain line
tube to fittings or connections. Position clamps on
tubes to structure and secure with screws, washers,
and nuts.
2-88A/(2-88B blank)
TM 5-1 520-234-23
tubes to structure and secure with screws, washers,
and nuts. Connect the B-nut of flexible hose
assembly to fitting or connection at each end of hose.
2-36. Vent and Drain Installation, Hydraulic.
Drain installations for the hydraulic system
are shown in figure 2-48A.
2-37. Vent and Drain Installation - Battery.
Each of the two hydraulic reservoirs, located in
the hydraulic compartment, are equipped with
scupper drain lines and reservoir drain lines which
connect into common overboard drain lines (14,
figure 2-48A).
The nickel-cadmium battery is mounted in the
helicopter battery compartment. See figure 1-1.
Two overflow, or vent tube assemblies, extend
from the battery to the left side of the fuselage (10,
figure 2-48A). The vent tube assemblies are
constructed of rubber material.
Each of the two hydraulic pumps, located on
lower side of the transmission, have seal drain
lines which connect into a common overboard
drain line (14).
a. Inspection.
(1) Clogged condition.
a. Inspection. Inspect reservoir and pump seal
drain lines for the following defects.
(2) Breaks in line. No breaks or cracks are
acceptable.
(1) Clogging.
(3) Pinched tube. Pinched tube walls that
significantly reduce area and could possibly result
in clogging are not acceptable.
(2) Breaks in line. No breaks are acceptable.
(3) Pinched line. No pinched lines are
acceptable if area of line is reduced significantly.
(4) Chafing. Determine cause of chafing and
extent of damage. Parts with severe chafing damage
are not acceptable.
(4) Chafing. Determine cause of chafing
and extent of damage. Parts with severe chafing
damage are not acceptable.
(5) Loose or missing supporting clamps.
b. Removal. Remove clamps from both ends of
two vent tube assemblies and remove vent tubes.
Retain clamp for reinstallation.
(5) Loose or missing support clamps.
b. Removal. Remove the B-nut of tube at nearest
fitting or connection. Remove screws. Remove
screws from clamps. Retain screws, washers, and
nut for reinstallation. Disconnect the B-nut at each
end of flexible hose assembly and remove the hose.
c. Repair.
(1) Clear clogged vent tube with probes and
blow out particles with compressed air regulated to
low pressure.
c. Repair.
(1) Clean clogged vent tube with safety wire (C152).
(2) Replace lines with breaks, cracks, severely
pinched areas or severely chafed areas.
(2) Replace lines with breaks, cracks, severely
pinched areas or severely chafed areas.
(3) Correct conditions that caused chafing
damage.
(3) Correct conditions that cause chafing
damage.
d. Installation. Connect the B-nut of drain line
tube to fittings or connections. Position clamps or
d. Installation.
(1) Position vent tube assemblies in
helicopter with lower scarf facing forward and
vent flush with aircraft outer skin to cause air flow
through battery compartment.
Change 2
2-89
TM 55-1520-234-23
(4) Remove side armor panels (1).
(2) Install clamps.
(5) Detach inertia reel control (6) from side of
cockpit by removing screws, washers and spacers.
2-38. Seat - Pilots.
Ensure that both pilot and gunners
arming/firing mechanism handles are
secured by safety pins prior to entry into
the cockpit area.
The pilots seat is one-piece, bucket-type seat
mounted on two vertical tubes which hold it in place
on the airframe structure and serve to make the seat
adjustable vertically. See figure 2-49. The seat is
constructed of armor steel with fittings for armor
side panels.
a. Inspection. Inspect installed pilots seat for the
following defects:
(1) Refer to paragraph 2-43 for inspection
procedure for side armor panels (1, figure 2-49) and
armor steel seat (2, figure 2-50).
(2) Cracks. No cracked parts are acceptable.
(3) Secure mounting of the seat in the
helicopter and secure installation of the inertia reel
and armor panels.
(6) Remove seat attaching parts (7) holding
seat assembly (8) to support brackets and remove
seat assembly from helicopter.
(7) Remove nuts, washers and bolts and
remove seat lap belt (9).
(8) Detach shoulder harness (10) from inertia
reel strap by removing nut, washer and bolt.
(9) Remove four nuts, washers and screws
attaching inertia reel (11) and remove inertia reel.
c. Disassembly. See figure 2-50.
(1) Adjust the seat to "up" position on support
tubes (3). Remove two return springs (5).
(2) Pull up on handle (1) to withdraw pins (10)
from tubes (3). Pull upward on tubes (3) and remove
them from fittings (4 and 7).
(3) Disconnect handle levers (6) from latch
pins (10).
(4) Remove retainer plates (8) from lower
fittings (7) and remove latch pins (10) and springs (9).
(4) Seat cushion band back cushion for wear
and damage, sun fading is not cause for rejection,
but wear and damage that affects comfort must be
corrected by repair or replacement.
(5) Remove two upper fittings (4) and lower
fittings (7) from seat (2).
(5) Check seat vertical adjustment, ease of
operation and secure locking in various height
positions.
(1) Inspect upper guide fittings (4) and lower
guide fittings (7) by fluorescent penetrant method.
d. Inspection. See figure 2-50.
(2) Inspect handle assembly (1), support tubes
b. Removal. See figure 2-49.
(1) Loosen clamp and disconnect air
distribution duct from duct cushion air inlet (2).
(2) Remove seat back cushion (3) and air ducts (4).
(3) Remove seat cushion (5).
CAUTION
Handle side armor panels with care;
ceramic tile is easily broken.
Change 22
(3), latch springs (9), return springs (5) and latch
pins (10) by magnetic particle method.
(3) Inspect seat netting for tears, cuts and
holes. Damage greater than one inch in length or
diameter is not reparable. Temporary repairs can be
made to damage less than one inch in length or
diameter.
(4) Inspect seat netting for deterioration and
discoloration which indicates a decrease in strength.
If integrity of netting is doubtful, the netting must
be replaced.
2-90
TM 55-1520-234-23
Figure 2-49. Pilots seat installation
2-91
TM 55-1520-234-23
Figure 2-50. Pilots seat assembly
e. Repair.
(1) Refer to paragraph 2-43 for repair
procedure for side armor panels (1, figure 2-49) and
armor steel seat (2, figure 2-50).
(2) Replace any parts found to be cracked
during fluorescent penetrant and magnetic particle
inspections in preceding paragraph.
(3) Repair tears, cuts or holes in netting that
are within one inch maximum length or diameter if
new cushion is not available. Use nylon thread, 16
pound test, FSM 8310-227- 1244 to repair by darning
procedure. Pick up at least 1/4 inch of good material
around the repair area. Maintain thread tension to
produce a mend that disturbs the natural lines of the
seat netting as little as possible.
2-92
TM 55-1520-234-23
(4) Replace cushion if deteriorated or damaged
beyond repair limits. Comfort is the determining
factor for replacement of cushions.
(5) Install inertia reel control (6) on side of
cockpit with spacers, washers and screws.
(6) Install seat cushion (5), air ducts (4) and
back cushion (3).
(5) Replace any parts of the seat vertical
adjustment mechanism that are damaged to the
degree that function is affected.
(7) Position air distribution duct on air inlet
(2) and tighten clamp to secure air distribution duct.
(8) Install side armor panels (1).
f. Assembly. See figure 2-50.
(1) Position handle assemblies (1) in place in
each lower fitting (7).
(9) Functional check inertia reel (11) and
control (6) to ensure the reel will lock, unlock and
rewind.
(2) Install lower fittings (7) on seat assembly (2).
(3) Install upper fittings (4) on seat assembly.
(4) Install a latch pin (10) and latch spring (9)
in each lower fitting (7). Install retainer plates (8) to
secure latch pins and springs.
(10) Functional check operation of seat.
vertical adjustment.
2-39. Seat - Gunners.
(5) Connect latch pins (10) to levers (6) of
handle assembly (1) with clevis pins, washers and
cotter keys.
(6) Install support tubes (3) down through
upper fittings (4) and lower fittings (7). Hold handle
(1) in UP position to permit passage of support tubes
through lower fittings (7).
(7) Install return springs (5).
g. Installation. See figure 2-49.
Before any maintenance in or near the
cockpit area, ensure that both pilot and
gunners arming/firing mechanism
handles are secured by safety pins.
The gunners seat is a two-piece, bucket-type seat.
The two major components are the back and the
bottom. See figure 2-51. Construction is ceramic
plate armor.
a. Inspection. Inspect installed gunners seat for
the following defects. See figure 2-51.
(1) Position inertia reel (11) on back of seat
bottom and install attaching screws, washers and
nuts with heads of screws on inside of seat bucket.
(1) Refer to paragraph 2-43 for inspection
procedure for side armor panels (1, figure 2-51), seat
back (7) and seat bottom panel (9).
(2) Thread shoulder harness (10) through
guide at top of seat back and attach to inertia reel
strap with bolt washer and nut. Install the bolt with
head facing seat back.
(2) Seat cushion (3) and back cushion (5) for
wear and damage. Use same procedure described for
pilots seat cushions. Refer to paragraph 2-38.
(3) Install seat lap belt (9) with bolts, washers
and nuts. Install belt half with latch on left side.
Install bolts with heads of bolts on inside of seat
bucket.
(4) Position seat assembly (8) in helicopter and
fit support tubes into support brackets with inertia
reel cable routed inboard of left support tube as
illustrated. Install seat attaching bolts, washers and
nuts (7).
(3) Secure mounting of the seat in the
helicopter and secure installation of the inertia reel
and armor panels.
b. Removal. See figure 2-51.
(1) Loosen clamp and disconnect air
distribution duct from air inlet (6).
(2) Remove seat cushion (3) and air ducts (4).
(3) Remove seat back cushion (5).
2-93
TM 55-1520-234-23
Figure 2-51. Gunners seat installation .
2-94
TM 55-1520-234-23
CAUTION
CAUTION
The pilot's and gunner's shoulder
harness inertia reels are not interchangeable.
Handle side armor panels with care;
ceramic is easily broken.
(4) Remove side armor panels (1).
(5) Remove six screws and washers which
attach seat back (7) to bulkhead and remove seat
back.
(6) Remove screws, washers and spacers which
attach inertia reel control (2) to side of cockpit.
Loosen knurled nut and disconnect control cable
from handle.
(7) Remove four screws and washers which
attach inertia reel (8) to bulkhead. Work shoulder
harness back through bulkhead and remove with
inertia reel.
(8) Remove bolts, washers and nuts which
attach seat lap belt (11) to attaching strap (10) and
remove bolt.
(3) Work shoulder harness through bulkhead
and position inertia reel (8) against bulkhead.
Secure with four screws and eight washers. Place
one washer under each screw head and one washer
under each inertia reel attachment lug.
(4) Route inertia reel control cable along left
hand side of seat and connect to inertia reel control
(2). Tighten knurled nut and position control on
beam. Install attaching spacers, washers and
screws. Lockwire knurled nut.
(5) Attach seat lap belt (11) to attaching strap
assembly (10) with bolts, washers and nuts.
(6) Connect air ducts (4) on seat cushion (3)
and install cushion.
(7) Connect air ducts (4) to seat back cushion
(5) and install cushion.
CAUTION
Handle armor panels with care, ceramic
is easily cracked.
(8) Position air distribution duct on air inlet
(6) and tighten clamp.
(9) Install side armor panels (1).
(9) Remove screws which attach seat bottom
panel (9) to structure and remove seat bottom panel.
c. Repair.
(1) Refer to paragraph 2-43 for repair
procedure for side armor panels (1, figure 2-51), seat
back (7) and seat bottom panel (9).
(10) Functional check inertia reel (8) and
control (2) to ensure the reel will lock, unlock, and
rewind.
2-40. Inertia Reel (Shoulder Harness).
(2) Repair tears, cuts and holes in seat cushion
(3) and back cushion (5). Use same procedure
described for pilots seat cushions. Refer to
paragraph 2-38.
d. Installation. See figure 2-51.
(1) Place seat bottom (9) in helicopter and
secure with screws and washers.
(2) Place seat back (7) in helicopter and secure
with bolts and washers.
Before any maintenance in or near the
cockpit area, ensure that both pilot and
gunner arming/firing mechanism handles
are secured by safety pins.
Inertia reels (11, figure 2-49) and (8, figure 2-51)
serve to control the pilot and gunners shoulder
harness. The pilot can select "Lock" or "Auto" with
control (6, figure 2-49). The gunner has a similar
control.
a. Inspection. See figure 2-49.
NOTE
Instructions to inspect pilots inertia reel
follow. Inspect gunners inertia reel in the
same manner.
Change 38
2-95
TM 55-1520-234-23
upward on strap with at least six pounds of force
and hold. Move control handle to AUTO position
and allow strap to rewind.
(1) Place control (6) in LOCK position and
pull on shoulder harness; the inertial reel should
hold the should harness and not extend. Inertia
reels that will not lock are not acceptable.
(2) Place control (6) in AUTO position and
pull on shoulder harness: the inertia reel should
permit the shoulder harness to be pulled out
against spring tension and should rewind when
pulling pressure on shoulder harness is decreased.
Inertia reels that will not operate as described in
this paragraph are not acceptable.
(3) Replace inertia reel and/or control if cable
is frayed or if the components have incurred damage
that may affect function.
d. Installation. Refer to paragraph 2-38 or 2-39
as applicable for installation and functional check
instructions.
2-41. Shoulder Harness and Seat Belts.
(3) Inspect inertia reel strap for wear fraying
and general condition. (Refer to TM
55-1500-204-25/1.)
(4) Inspect inertia reel (11) and control (6) for
secure mounting and damage.
Ensure that both pilot and gunner
arming/firing mechanism handles are
secured by safety pins prior to entry into
the cockpit area.
(5) Inspect cable between inertia reel (11) and
control (6) visually for fraying and damage.
b. Removal. Refer to paragraphs 2-38 or 2-39 as
applicable.
c. Repair.
The shoulder harness and the seat belt
installations for the pilot and gunner seats is
similar. See figures 2-49 and 2-51.
(1) Replace inertia reel that fails to pass
functional check in preceding inspection paragraph
b, steps (1) and (2).
NOTE
Gunner's shoulder harness webbing adjuster, part number ASE443030, does
not require a Webbing Adjuster Spring.
If spring is installed, it should be removed.
(2) Replace worn inertia reel strap as follows:
(a) Move inertia reel control handle to
AUTO position and pull out slowly on strap
assembly until web retaining insert is visible
through lower slot in reel housing.
a. Inspection.
(1) Inspect pilot and gunner seat lap belts
and shoulder harnesses for fraying, wear, tears,
and general condition. Refer to TM 55-1500-204-25/1.
CAUTION
If reel is inadvertently released while
strap is removed, replace entire reel
assembly.
(2) Inspect to verify belt can be released by
gloved hand.
(b) Move control handle to LOCK position.
(a) Using a steel scale, measure the
distance from the top of the webbing where it
encircles the latch to the inside of the release plate.
This is the space where gloved fingers must be
inserted to release the latch.
(c) Remove web retaining insert and
withdraw strap from reel.
(d) Insert end of new strap through upper
slot in reel housing and through slot in main shaft
until end of strap protrudes through lower slot in
reel housing. Install web retaining insert. Pull
(b) If the measured space is 0.8 inch or
greater, the belt is satisfactory.
Change 48
2-96
TM 55-1520-234-23
(c) If the measured space is less than 0.8
inch, the belt is unsatisfactory. The belt must be
altered in accordance with TM 55-1500-204-25/1 to
provide clearance specified in step (b) above.
b. Removal. Refer to paragraph 2-38 or 2-39
as applicable.
c. Installation. Refer to paragraph 2-38 or 2-39
as applicable.
(1) Remove electronic interface assembly. Refer to TM 9-1270-212-14.
(2) Removal - Release hook and pile
fasteners and remove soundproofing blanket.
c. Repair. Refer to TM 55-1500-204-25/1 for instructions to make minor repairs to soundproofing
blanket.
d. Installation. Position the soundproofing
blanket assembly behind the pilots station and
fasten the hook and pile fasteners. Install electrical interface assembly removed in step b.
2-42. Soundproofing Blanket Assembly.
2-42A. Deleted.
Ensure that both pilot and gunner arming/firing mechanism handles are secured by safety pins prior to entry into
the cockpit area.
A soundproofing blanket assembly is installed behind the pilot station to reduce
the noise level during flight operations.
2-43. Armor - Protective.
Ensure that both pilot and gunners arming/firing mechanism handles are
secured with safety pins prior to entry
into the cockpit area.
a. Inspection. Inspect the soundproofing
blanket for fraying, wear, tears and secure installation.
b. Removal.
Change 56
The crew and engine are protected against
hostile arms fire by ceramic and steel armor
panels. See figure 2-42.
2-96A/(2-96B blank)
TM 55-1520-234-23
Figure 2-52. Armor panels
Change 44
2-97
TM 55-1520-234-23
(g) Damage to armor panel attaching brackets.
Minor damage is reparable.
NOTE
Temporary removal of armor panels: All of the
armor shown of figure 2-52 may be removed in
non-combat areas at the discretion of the unit
commander. If armor is removed, comply with
the following requirements:
Identify armor panels for reinstallation and retain as
flyaway equipment in a safe storage area where it is
readily available.
(h) Refer to TM 55-1500-204-25/1 for additional
inspection criteria if required.
b. Removal.
(1) Refer to paragraphs 2-36 and 2-37 for removal of
armor panels at pilot and gunners seats.
(2) Open engine cowling and remove engine ceramic
armor panels (5, figure 2-52).
Reinstall armor prior to transfer of helicopter.
Make entries in helicopter weighing record and chart C,
basic weight and balance record, when armor is removed
and again when it is reinstalled.
a Inspection.
(1) Inspect pilots armor steel seat (4, figure 2-52) for
damaged brackets and cracks.
(2) Inspect ceramic-type armor panels (1, 2, 3, 5, 6,
7, and 8, figure 2-52) for the following defects:
(a) Damage caused by a ballistic projectile.
Panels with this type damage are not reparable and must
be replaced.
(b) Cracks in ceramic armor and damage to the
glass reinforced plastic backing. Hairline cracks are
reparable if the glass reinforced plastic backing is not
damaged. If cracks are wider than hairline and the glass
reinforced plastic backing is damaged, the panel must be
replaced.
(c) Chipping and spalling damage.
Minor
damage on the edges and up to x inch from the edges is
acceptable without repair. If chipping or spalling damage
is present elsewhere on the panel, the panel must be
replaced.
(d) Damage that results in loose nylon cloth
shield and/or loose neoprene rubber edge moulding. Mark
loose areas for rebonding.
(e). Damage that results in delamination is not
reparable.
(f) Damage to threads in threaded inserts, loose
bonding of threaded inserts to panel and missing threaded
inserts.
C
Repair
(1) Repair hairline cracks in ceramic armor panels
that are within repair limits by laminating fiberglass plies
(C47) with resin (C 107) over the crack and under the
spall shield to a thickness of 0.25 inch and overlapping
the crack by 3 inches on each side. Cover with cellophane
(C33), squeeze and cure at room temperature (77°F) for a
minimum of 8 hours, or for 2 hours at 105°F.
(2) Rebond loose nylon cloth shield or neoprene
rubber edge moulding with adhesive (C20).
(3) Repair threaded fasteners with slightly damaged
threads by cleaning up threads with a tap. Replace the
threaded fastener with a new fastener if thread damage is
severe. Refer to step (5) for bonding procedure.
(4) Repair a loose threaded fastener by rebonding.
Drill two small holes in the backing at an angle down to
bottom of threaded insert and clean out holes. Place
masking tape over the two small holes and the threaded
hole in the insert to keep adhesive out of the threads.
Use a sharp pointed tool to open holes through the
masking tape at the two drilled holes. Inject adhesive
(C12) into one hole with a syringe. Continue injecting
adhesive until it is forced from the second hole. Allow to
cure for 24 hours.
(5) Repair threaded fastener with faulty threads or
one that is very loose in the panel by replacing with a new
threaded fastener. Drill out the old fastener carefully to
avoid damage to the panel. Clean out the hole and bond
a new fastener in the panel in the same manner outlined
in the preceding step.
(6) Repair cracked or distorted armor panel attaching
brackets with standard metalworking
2-98 Change 22
procedures if practical.
reparable.
Replace brackets that are not
(e) Corrosion. Inspect for corrosion damage that
could affect function of engine mounts.
d. Installation.
(1) Refer to paragraphs 2-38 and 2-39 for Installation
of armor panels at pilot and gunners eats.
(f) Cracks. Inspect for cracks visually. If any
areas are suspect, inspect by dye penetrant method. No
cracks are acceptable.
(2) Open engine cowling, install armor panel (5,
figure 2-52) and close engine cowling.
(g) Security. Check for secure installation of all
bolts and rivets in each of the three mounts.
2-44. Engine Mount Installation.
(2) Inspect pillow blocks (4 and 10, figure 2-53) for
the following defects:
The engine is supported on the three engine mounts
shown on figure 2-53. The mounts are made primarily of
steel tubing and rod end bearings. Hinged pillow block
assemblies on the bipod and tripod mounts attach to
fittings on the engine diffuser housing.
The rigid
connecting link attaches to a fitting on the engine inlet
housing. The three engine mounts attach to fittings on
the engine compartment deck. Shims are installed under
the fittings to align the engine with the transmission at
original installation.
(a) Cracks. No cracks are acceptable.
(b) Nicks, scratches and dents severe enough to
affect function.
(c) Wear. If noticeable wear is present, check
bearing with dial indicator. Maximum allowable play is
0.006 inch radial and 0.012 inch axial.
(3) Inspect eye bolts (3 and 9, figure 2-53) for
damaged threads.
a. Inspection.
(4) Inspect fittings (1, 8, 13, 15, 16, and 20, figure 253) for the following defects:
(1) Inspect engine tripod, bipod and rigid connecting
link (5, 11 and 19, figure 2-53) for the following defects.
(a) Cracks. No cracks are acceptable.
(a) Dents. Small, smooth dents which have not
removed material from tubing and occur at least one and
one-half inches from weld clusters or from points where
tubes intersect gussets are considered negligible and do
not require repair. Dents greater than negligible are
cause for replacement of the affected engine mount.
(b) Nicks and dents severe enough to affect
function.
(c) Secure attachment to deck.
(d) Corrosion damage severe enough to affect
function.
(b) Scratches. Transverse scratches longer than
5/16 inch are cause for replacement of the mount. Other
scratches that are less than 0.010 inch deep may be
polished out.
b
(c) Distortion.
Inspect for bends, nicks and
similar damage. Any distortion-type damage that can be
detected visually is cause for replacement of the affected
mount.
Removal.
'
Do not remove fittings (1, figure 2-53) or similar
fittings under engine mounts unless fitting must
be replaced.
(d) Wear. Inspect for worn rod end bearings in
the engine mount tubes. If noticeable wear is present,
remove the affected engine mount and check bearing
wear with a dial indicator. Refer to step b for removal
instructions. Maximum allowable radial play is 0.008
inch. Maximum allowable axial play is 0.016 inch.
(1) Support engine with suitable support to relieve
weight from mounts.
2-99
TM 55-1520-234-23
(2) Remove engine fuel control linkage and bellcrank
from tripod mount on left side. Remove hose clamps
from mount tubes.
(3) Remove bolts, nuts and washers (2 and 7, figure
2-53) and remove engine mount (5). Remove engine
mounts (11 and 19) in the same manner.
(4) If fittings (1, figure 2-53) or similar fittings under
other engine mounts must be removed, remove the bolts
that secure the fitting to the deck and remove fittings.
(5) If pillow blocks (4 and 10, figure 2-53) must be
removed, remove two bolts (6) and remove pillow block
(4). Remove pillow block (10) in the same manner.
c
Repair.
(1) Polish out minor corrosion damage and touch up
with primer (C102).
(2) Replace engine mount parts that are damaged or
worn beyond allowable limits.
Install mount in work aid as shown on sheet 3 of figure 254. Tighten bolts (1), (2), and (6).
(d) Install mount and work aid on a universal drill
press table and adjust so that a 3/16 twist drill mounted in
drill press will pass smoothly through open rivet holes in
mount and rod end. Remove bolt (1, figure 2-54), rivet
and unserviceable rod end. Install new rod end and
tighten bolt (1). Install a 7/32 inch, No. 2 (0.2210 inch)
twist drill in drill press and drill through a new rod end.
Remove rod end, deburr and inspect mount to ensure that
drill did not go off center and enlarge the hole on lower
side of mount.
(e) Reinstall unserviceable rod end in mount and
use same procedure outlined in preceding step to align
and drill out second rivet hole.
(f) Inspect engine mount to ensure that the drill
did not go off center and enlarge rivet holes. Scrap mount
and rod end if damaged during drilling operation.
WARNING
NOTE
209-062-127-1 rod end allows removal and
replacement of the bearing uni-ball.
(3) Rotate uni-ball bearing one-half turn and remove
uni-ball bearing. Install new uni-ball bearing in a reverse
manner.
Cleaning solvent is flammable and toxic.
Provide adequate ventilation. Avoid prolonged
breathing of solvent vapors and contact with
skin or eyes.
(g) Use crocus cloth (C45) to roughen mating
surface of new rod end and mount. Clean area with cloth
dampened with solvent (C87).
(4) Recheck engine mount rod end bearing for radial
and axial play. Refer to paragraph a.(d).
(5) Rod ends that are worn beyond allowable limits
may be replaced as follows:
(a) If not previously accomplished, fabricate a
four-part work aid as shown on figure 2- 54.
(b) Grind off bucked ends of rivets which secure
the unserviceable rod end to engine mount. Use caution
to avoid damaging engine mount. Drive rivets out with a
punch and remove unserviceable rod end.
(c) Temporarily reinstall unserviceable rod end in
mount and align with one new rivet, but do not buck rivet.
Ensure that correct part number rivet is installed
in following step. Refer to TM 55-1520-23423P.
(h) Apply a light coating of adhesive (C12) to
mating surfaces of new rod end and mount. Position rod
end in mount and install rivets. Allow adhesive to cure for
24 hours.
(i) Mark and center punch one location 1.75 inch
from each end of mount tube. Drill a number 30 (0.128
inch) hole at each of the marked locations. Fill mount
tube with lacquer (C78) and allow to drain. Install rivets,
P/N MS20600B4K-1, in the holes.
2-100 Change 7
TM 55--1520-234-23
d. Installation.
NOTE
After proper shim thickness has been selected,
apply proseal (C 119) under the engine mount
fitting.
(1) Install fitting (1, figure 2-53) or similar fittings
under engine mounts if removed. Install the fittings
temporarily. It may be necessary to change shim
thickness under fittings after engine alignment.
(2) Position engine mount (5, figure 2-53) on deck
fittings and install bolts, washers and nuts (2 and 7).
Install mounts (11 and 19, figure 2-53) in the same
manner.
(2) Secure connecting link (19, figure 2-53) with
bolt and washer (18) to forward trunnion fitting. Torque
bolt 50 to 70 inch-pounds.,
(3) Install hose clamps on mount tubes that were
removed in step b.
(4) Install pillow blocks (4 and 10, figure 2-53) if
not previously accomplished. Position pillow block
(4) on bipod mount with eye bolt forward as
Change 65
2-100A/(2-100B blank)
TM 55-1520-234-23
Figure 2-53. Engine mount Installation
All data on pages 2-102 thru 2-104 including figure 2-54 deleted
Change 64 2-101
illustrated. Install two bolts, washers and nuts (6). Install
pillow block (10) in the same manner.
(5) Install fuel control bellcrank and linkage that
was removed in step b.
(6) Remove engine support.
(7) Perform an engine to transmission alignment
check, refer to paragraph 6-8, when any engine mount
component is replaced or when any engine deck mount
fitting is removed and reinstalled.
2-44A. Wire Strike Protection System (After MWO 551520-234-50-2 and MWO 55-1520-234-50-3).
The wire strike protection system consists of three
cutters, and three deflectors. See figure 2-54A for view of
system. A deflector is mounted on the nose of the
helicopter between the telescope sight window and the
laser sight window. Two deflector assemblies, each
consisting of a channel with an insert, are mounted on the
right and left center canopy frames. Two nose deflectors
are mounted on each side of the nose and at the forward
end of each deflector assembly. A wire cutter is mounted
on a honeycomb panel and secured on the aft end of the
pilot overhead window. A second wire cutter is mounted
on the access door below the turret sight unit and forward
of the turret. The third wire cutter is mounted on the
access door below the ammunition compartment.
a.
c
(1) Recoat cutter blades with sealant (CI19E) that
have protective sealant missing. Replace blade (48
and/or 50, detail C) which shows evidence of wire strike
as follows:
(a) Remove screws (12), washers (2 and 11),
and nuts (3) and remove blade (48) or blade (50) and
shims (49) as required. Retain shims (49) for reuse.
(b) Coat replacement blade (48 and 50) with
sealing compound (C1 19E) to seal gaps between blade
and cheekplate (51).
(c)
Position replacement blade in cheekplate
and shim as required to achieve a maximum gap of 0.16
inch. Shim equally either on both sides or same side of
blade to ensure cutting edge alignment.
(d) Install screws (12), washers (2 and 11),
and nuts (3). Do not torque.
(e) Ensure gap, shown in detail F, measures
0.020 inch maximum with back of blades (48 and 50)
forced apart. If overlap exceeds 0.010 inch, dress corner
of lower blade (50) to reduce overlap.
(f)
Torque screws (12) 30 - 35 in-lbs.
(g) Coat cutting edge of replacement blade
with sealing compound (C119E).
(2) Touch up damaged paint to match existing
finish using olive drab paint (C75A).
Removal- Wire Strike Cutter (Upper).
(1) Remove screws (13, figure 2-54A), and
remove cutter (8) and panel (4) as a unit.
(2) Remove attaching screws (1 and 10), and
remove cutter from panel.
(3) Remove bolt (5), washer (6), and nut (7) and
remove struts (9).
b.
Repair-- Wire Strike Cutter (Upper).
(3)
d
Repair panel in accordance with paragraph 2-4.
Installation - Wire Strike Cutter (Upper).
(1) Install struts (9, figure 2-54A) on cutter (8) with
nut (7), washer (6), and bolt (5).
(2) Position panel (4) and cutter (8) on aft end of
pilot upper window.
Inspection - Wire Strike Cutter (Upper).
(1) Inspect cutter for bends, cracks, nicks,
alignment, and presence of protective sealing compound
on blade.
(3) Apply a 2-inch wide bead of sealing compound
(C116) to fill gap between the window and the bottom of
the cutter panel.
(2)
delamination,
(4) Secure panel (4) and cutter (8) with screws (1,
10 and 13).
Inspect panel for cracks, punctures,
and
pulled
or
loose
inserts.
Change 65
2-104A
TM 55-1520-234-23
Figure 2-54A. Wire strike protection system (Sheet 1 of 4)
2-104B
Change 44
TM 55-1520-234-23
Figure 2-54A. Wire strike protection system (Sheet 2 of 4)
Change 44
2-104C
TM 55-1520-234-23
Figure 2-54A. Wire strike protection system (Sheet 3 of 4)
2-104D
Change 44
TM 55-1520-234-23
1. SCREW
2. WASHER
3. NUT
4. PANEL
5. BOLT
6. WASHER
7. NUT
8. CUTTER
9. STRUT
10. SCREW
11. WASHER
12. SCREW
13. SCREW
14. DOOR
15. RIVET
16. STRUT
17. WAS:HER
18. BOLT
19. BOLT
20. WASHER
21. RIVET
22. CUTTER
23. BOLT
24. WASHER
25. NUT
26. STRUT
27. SCREW
28.
29.
30.
31.
32.
33.
34.
35.
36.
37.
38.
39.
40.
41.
42.
43.
44.
45.
46.
47.
48.
49.
50.
51.
52.
53.
54.
CUTTER
WASHER
BOLT
BOLT
BOLT
WASHER
BOLT
DEFLECTOR
INSERT
SCREW
SCREW
CHANNEL
RIGHT NOSE DEFLECTOR
SCREW
WASHER
NUT
LEFT NOSE DEFLECTOR
NUT
WASHER
SCREW
BLADE
SHIM
BLADE
CHEEKPLATE
CHEEKPLATE
BLADE
BLADE
NOTES
SEAL GAPS BETWEEN CHEEKPLATES AND BLADES BY WET INSTALLING BLADES COATED WITH
SEALING COMPOUND (C119E).
SHIM AS REQUIRED BETWEEN CHEEKPLATES AND BLADES TO ACHIEVE A MAXIMUM GAP OF 0.16
INCH TOTAL. SHIM EQUALLY ON BOTH SIDES OR SAME SIDE OF ALL BLADES TO ENSURE
ALIGNMENT OF CUTTING EDGE.
TORQUE SCREWS 30-35 IN-LBS. ,/A GAP 0.020 INCH MAXIMUM WITH BACK OF BLADES FORCED
APART BEFORE TIGHTENING SCREWS.
OVERLAP EXCEEDING 0.010 INCH SHALL BE REDUCED BY DRESSING CORNER OF INDICATED
BLADE.
COAT CUTTING EDGE OF ALL BLADES WITH MINIMAL APPLICATION OFSEALING
(C119E). 209704-40-4
Figure 2-54A. Wire strike protection system (Sheet 4 of 4)
Change 44
2-104E
COMPOUND
e
TM 55-1520-234-23
(1) Replace nose deflector which shows evidence of
wire strike.
Removal- Channel and Insert (Wire Strike).
(1) Remove screws (38, detail A, figure 2-54A) and
remove insert (36).
(2) Touch up damaged paint on nose deflector using
olive drab paint (C75A).
(2) Remove screws (37) and remove channel (39).
f
Inspection-- Channel and Insert (Wire Strike).
l. Installation - Nose Deflector, Right and Left (Wire
Strike).
(1) Inspect inserts (36, figure 2-54A) for bends,
cracks, or loose or missing fasteners or protective sealing
compound.
(1) Position right nose deflector (40) in place and
secure with screws (41 and 47), washers (42 and 46), and
nuts (43 and 45).
(2) Inspect channels (39) for distortion, loose or
missing fasteners, and damage to paint.
(2) Position left nose deflector (44) in place and
secure with screws (41 and 47), washers (42 and 46), and
nuts (43 and 45).
(3) Inspect channel attach holes in center canopy
frames for loose or missing flush filler.
g Repair-- Channel and Insert (Wire Strike).
m Removal- Wire Strike Cutter (Nose).
(1) Remove bolts (30 and 31, figure 2-54A) and
washers (29).
(1) Replace insert which has protective sealing
compound missing or which shows evidence of wire
strike.
(2) Remove cutter (28).
m Inspection- Wire Strike Cutter (Nose).
(2) Flush exposed channel attach countersunk holes
with adhesive, Hysol (C17).
(3) Touch up damaged paint on channel using olive
drab paint (C75A).
(1) Inspect cutter (28, figure 2-54A) for bends,
cracks, nicks, and alignment.
(2) Inspect condition of protective sealing compound
on cutter blade.
h. Installation - Channel and Insert (Wire Strike).
o. Repair-- Wire Strike Cutter (Nose).
(1) Position channels (39, figure 2-54A) on canopy
structure and secure with screws (37).
(2) Position inserts (36) in channels (39) and secure
with screws (38).
(1) Recoat cutter blades with sealant (Cl19E) that
have protective sealant missing. Replace blade (53 or
54, detail E) which shows evidence of wire strike as
follows:
I. Removal-- Nose Deflector, Right and Left (Wire
Strike).
(a) Remove screws (12), washers (2 and 11), and
nuts (3). Remove blades (53 or 54) as required.
(1)
(43).
(b) Coat replacement blades (53 and 54) with
sealing compound (C119E) to seal gaps between each
blade and cheekplates (52).
Remove screws (41), washers (42), and nuts
(2) Remove screws (47), washers (46), and nuts
(45). Remove nose deflectors (40 and 44) as required.
(c) Position blades (53 and 54) in cheekplates
(52).
j. Inspection-- Nose Deflector, Right and Left (Wire
Strike). Inspect right and left nose deflectors for evidence
of wire strike and damage to paint.
k. Repair - Nose Deflector, Right and Left (Wire Strike).
2-104F
(d) Install screws (12), washers (2 and 11), and
nuts (3). Do not torque.
(e) Ensure gap, shown in detail F, measures
0.020 inch maximum with blades forced apart. If overlap
exceeds 0.010 inch, dress corner of upper blade (54) to
reduce overlap.
Change 65
TM 55-1520-23423
Figure 2-54B. Damage limits - wire strike deflector
Change 44
2-104G
TM 55-1520-234-23
(d) Install screws (12), washers (2 and 11), and
nuts (3). Do not torque.
(f) Torque screws (12) 30 - 35 in-lbs.
(g) Coat cutting edge of replacement blade with
sealing compound (C 119E).
(2) Touch up damaged paint using olive drab paint
(C75A).
(e) Ensure gap, shown in detail F, measures
0.020 inch maximum with back of blades (48 and 50)
forced apart. If overlap exceeds 0.010 inch, dress corner
or upper blade (50) to reduce overlap.
(f) Torque screws (12) 30 - 35 in-lbs.
p. Installation - Wire Strike Cutter (Nose).
(g) Coat cutting edge of replacement blade with
sealing compound (C 119E).
(1) Position cutter (28, figure 2-54A) on nose.
(2) Secure cutter (28) to nose with bolts (30 and 31)
and washers (29).
(2) Touch up paint to match existing finish using olive
drab paint (C75A).
(3) Repair door (14) in accordance with paragraph 2-4.
q. Removal-- Wire Strike Cutter (Lower).
(1) Remove screws (27, figure 2-54A), bolts (18 and
23), and washers (17 and 24), securing door (14) to
bottom of ammunition compartment. Remove door (14)
with cutter (22) attached.
t Installation-- Wire Strike Cutter (Lower).
(2) Remove rivets (15 and 21). Remove cutter (22)
from door (14).
(2) Install cutter (22) on door (14) with new rivets (15
and 21).
(3) Remove bolt (19), washer (20), and nut (25).
Remove struts (16 and 26).
(3) Install door (14), with cutter (22) attached, on
helicopter using bolts (18 and 23), washers (17 and 24),
and screws (27).
r.
(1) Install struts (16 and 26, figure 2-54A) on cutter
(22) with bolt (19), washer (20), and nut (25).
Inspection-- Wire Strike Cutter (Lower).
U
(1) Inspect cutter (22, figure 2-54A) for bends,
cracks, and alignment.
(1) Remove bolts (32 and 34, figure 2-54A) and
washers (33).
(2) Inspect condition of protective sealing compound
on cutter blade.
(3) Inspect door (14)
damage, and loose inserts.
s.
for
delamination,
impact
Repair-- Wire Strike Cutter (Lower).
(1) Recoat cutter blades with sealant (Cl19E) that
have protective sealant missing. Replace blade (48
and/or 50, detail D) which shows evidence of wire strike
as follows:
(a) Remove screws (12), washers (2 and 11), and
nuts (3) and remove blade (48) or blade (50) and shims
(49) as required. Retain shims (49) for reuse.
Removal-- Wire Strike Deflector (Nose).
(2) Remove deflector (35) from nose.
v.
Inspection-- Wire Strike Deflector (Nose).
Inspect deflector (35, figure 2-54A) for damage in
accordance with figure 2-54B.
w
Repair-- Wire Strike Deflector (Nose).
(1) Replace deflector (35, figure 2-54A) if damaged
beyond limits shown in figure 2-54B.
WARNING
Cleaning solvent is flammable and toxic.
Provide adequate ventilation. Avoid prolonged
(b) Coat replacement blade (48 and 50) with
breathing of solvent vapors and contact with
sealing compound (C119E) to seal gaps between blade
skin or eyes.
and cheekplate (51).
(2) Polish out corrosion, nicks, scratches, and dents
not severe enough to reject deflector (35, figure 254A).
(c) Position replacement blade in cheekplate and
Use 300 grit sandpaper (C112). Clean sanding residue
shim as required to achieve a maximum gap of 0.16 inch.
with MEK (C87), and touchup bare metal with primer
Shim equally either on both sides or same side of blade to
(C100) and paint to match existing finish. Refer to TB
ensure cutting edge alignment.
746-93-2.
2-104H Change 65
TM 55-1520-234-23
x.
Installation - Wire Strike Deflector (Nose).
(1) Position deflector (35, figure 2-54A) on nose.
NOTE
(2) Install bolts (32 and 34) and washers (33).
When installing a new wire strike deflector, the
unit must be demagnetized prior to installation.
2-45.
Section II. TAILBOOM
(4) Disconnect electrical connectors.
Tailboom Assembly.
The tailboom (11, figure 2-55) is an aluminum alloy semimonocoque structure made up of bulkheads, longerons
and stringers covered by aluminum skins The tail fin is an
integral part of the tailboom and is made up of aluminum
ribs, a spar and honeycomb panels. The tailboom
supports the synchronized elevator, tail rotor, tail rotor
driveshaft, control systems, avionics equipment,
armament system equipment, and cooling equipment for
the avionics equipment.
Premaintenance Requirements for Tailboom.
Condition
Model
Part No. or Serial No.
Special Tools
Test Equipment
Support Equipment
Minimum Personnel
Required
Consumable Materials
Special Environmental
Condition
(3) Position TSU and laser window covers, and check
clearance around deflector (35).
Install covers
Installation
Requirements
AH-IS
All
None
Torque Wrench
1050-1150 inch-pounds
Tailboom,
Support Stand
(5) Disconnect tail rotor control.
(6) Disconnect synchronized elevator controls.
(7) Fabricate a stand or padded support to hold the
tailboom after removal. See figure 2-56.
(8) Place stands, prepared in preceding step, under
tailboom.
(9) Position two men on each side of tailboom
forward of synchronized elevator to support the tailboom
and lower it into stands. Remove two lower bolts (15 and
16, figure 2-57).
Direct men to support tailboom.
Remove two upper bolts (1 and 6). Lift up on vertical fin
and slide tailboom back to clear helicopter. Lower
tailboom into position on stands
and install bolts to secure Talbot to forward stand as
shown on figure 2-56.
b. Installation.
NOTE
Five
(C77)
If a new tailboom is being installed, install
electrical/avionics equipment, synchronized
elevator and controls as outlined in steps (1)
through (6). If the same tailboom is being
installed, proceed to step (7).
None
a. Removal.
(1) Remove tail pipe fairing and open tail rotor
driveshaft covers.
(2) Remove clamps on tail rotor driveshaft section at
forward end of tail boom and remove section of
driveshaft.
(1) Install electrical and avionics equipment in
tailboom.
(2) Install synchronized elevator and control system.
Refer to paragraph 2-0.
(3) Install tail rotor control linkage. Refer to Chapter
11.
just
(3) Open access panel on right side of fuselage and
forward
of
tailboom
attachment
point.
Change 44
(4) Install forty-two degree
gearboxes. Refer to Chapter 6.
2-105
and
ninety
degree
TM 55-1520-234-23
Figure 2-55. Airframe external components
2-106
TM 55-1520-234-23
that the bolts are of varying lengths. Identify bolts so they
can be installed in the proper location.
(5) Install tail rotor drive shaft. Refer to Chapter 6.
(6) Install tail rotor assembly. Refer to Chapter 5.
(7) Open access panel on right side of fuselage and
just forward of tailboom attachment point.
(8) Place countersunk washers (2, 7, 15 and 17,
figure 2-57) on tailboom attaching bolts with countersunk
side of washers toward bolt heads. Note
Figure 2-56.
(9) Position two men on each side of tailboom
forward of synchronized elevator to lift the tailboom into
position for installation. Direct men to sup)port tailboom.
Remove bolts that secure front stand to tailboom. See
figure 2-61.
(10) Ensure that four retainers (5, 9, 11, and 20,
figure 2-57) are in place and that barrel nuts (5, 9, 11, and
20) are aligned for installation of bolts.
Work aid - tailboom support stand
2-107
TM 55-1520-234-23
NOTE
Proper tailboom attachment bolt thread engagement has
been achieved when one thread and not more than three
threads are showing on bolt.
(11) Lift tailboom into position and install bolts with
countersunk washers that were prepared in step (8).
Ensure that the correct length bolt is installed for each
location. Install two upper bolts (1 and 6, figure 2-57)
first, then install two lower bolts. Tighten bolts carefully to
ensure that bolt threads do not bottom in barrel nuts. If
necessary, install flat
Figure 2-57. Tailboom Installation
2-108 Change 7
TM 55-1520-234-23
steel washers (3, figure 2-57), or similar washers on
remaining three bolts, between countersunk washer (2)
and fuselage fittings to obtain proper thread engagement.
Torque the four bolts (1, 6, 15 and 16) 1100 TO 1300
inch-pounds. Retorque bolts after first flight and apply
slippage index marks with lacquer (C77) or other suitable
marking material. Apply a thin bead of proseal (C1]6)
around tail boom to fueslage attachment point to
minimize water intrusion.
(12)
(16)
Install access panels.
(17)
Perform
functional
check
of
electrical/avionics/armament equipment in the tailboom
and perform maintenance test flight. Refer to TM551520-234MTF.
2-46. Tailboom Assembly Skin.
Install tail rotor driveshaft section.
NOTE
(13)
Connect synchronized elevator controls and
check rigging.
Repair is limited to repair of minor cracks, holes,
scratches, corrosion and replacement of loose
or missing hardware. If major damage occurs
which requires use of jigs and fixtures to repair,
forward tailboom to Depot Level Maintenance
for repair.
(14)
Connect tail rotor controls and check rigging.
NOTE
a. Inspection.
If tail rotor controls are found to be out of rig
during preceding step, determine whether tail
rotor has been removed and reinstalled. If tail
rotor has been removed and reinstalled, check
for proper installation of nylatron washer under
bearing at outboard end of crosshead that
supports tail rotor counterweights. Refer to
Chapter 5 for illustration and installation
instructions for nylatron washer.
(1) Wrinkles and buckled areas. All damage of this
type must be repaired.
(2) Popped and cocked rivets.
type must be repaired.
All damage of this
(3) Dents, cracks, holes, tears, nicks, scratches,
corrosion and trapped or stretched skin. Damage limits
are given in table 2-1.
(15)
Connect electrical connectors for electrical
and avionics.
Table 2-1. Tailboom Skin and Structure Classification of Damage
ITEM
1. TAILBOOM
SKINS
See
figure
2-58
DEFECT
a. Dents.
NEGLIGIBLE
DAMAGE LIMITS
REPAIRABLE
DAMAGE LIMITS
a. Smooth contour
free of cracks, nicks,
or wrinkles. Depth
and diameter not to
exceed:
a. Cracks or sharp
nick in dent. Damage
areas after cleanup
(including prior
repairs) shall not
exceed 20
of total area for
a single skin panel.
Damage 6.0 inch
minimum from
similar repair.
Depth
0.0156
0.0468
0.0625
Diameter
1.0 inch
2.0 inch
3.0 inch
Change 65
2-109
DAMAGE
REQUIRING
REPLACEMENT
a. Total damage
(including prior
repairs) exceeds 20
percent of total
area of a single
percent skin panel, or
damage spans entire
distance between
two bulkheads or
or two stringers.
TM 55-1520-234-23
Table 2-1. Tailboom Skin and Structure Classification of Damage (Cont)
ITEM
DEFECT
1. TAILBOOM
SKINS
See
figure
2-58
(Cont)
b. Cracks, holes,
tears, nicks,
scratches, and
corrosion.
c.
Trapped or
stretched skin.
NEGLIGIBLE
DAMAGE LIMITS
REPAIRABLE
DAMAGE LIMITS
DAMAGE
REQUIRING
REPLACEMENT
3.0 inch minimum
undamaged material
between dents and
1.0 inch minimum
from internal
structure. Nicks and
scratches which can
be blended out not
to exceed 10 percent of material
depth.
b. Nicks and
scratches no deeper
than 10 percent of
material thickness
and not exceeding
1.0 inch length by
0.25 inch width after
cleanup. Corrosion
damage less than 10
percent of material
thickness and not
exceeding 4.0 square
inch after cleanup.
Damage no closer
than 1.0 inch to a
supporting structure.
b. Damage exceeds
negligible limits but
does not exceed 25
percent (including
prior repairs) of total
area for a single
skin panel.
b. Same as dents.
c. Inward or outward bulges located
in a sectional area,
that can be corrected
by removing attaching hardware, allowing skin to shift.
Mismatch of rivet
holes shall not exceed that which
can be cleaned up
by drilling and
installing one size
larger rivet and
maintain proper
rivet edge distance.
However, if condition
does not disappear
after unloading panel,
c. Creased dents
not classified as oil
can or stretched
skin, not exceeding
25 percent of a
sectional area and
no closer than 1.0
inch to a supporting
structure. Oil can
condition, free of
sharp dents or
creases and not
extending over or
into supporting
structure may be
repaired by inserting
a backup stiffener
over the damaged
area.
c. Stretched skin,
oil cans, or creased
dents that cannot
be repaired by unloading insertion
repair or back up
stiffeners.
2-110 Change 2
TM 55-1520-234-23
Table 2-1. Tailboom Skin and Structure Classification of Damage (Cont)
ITEM
DEFECT
NEGLIGIBLE
DAMAGE LIMITS
1. TAILI
BOOM
SKINS
See
figure
2-58
(Cont)
area is stretched or
oil canned and must
be replaced or repaired. Oil canning
can be determined by
pressing in on a
sectional area and
that section remains
depressed and a
bulge appears in
that section or
adjacent structure.
2. TAIL- Dents, cracks,
BOOM holes, tears,
STRING- corrosion and
ERS
distortion.
AND
STIFFENERS
NOTE
Scratches or smooth
shallow dents not
extending into formed
radius and less than
10 percent of
material thickness and
0.50 inch length
after cleanup.
Damage in radius
treat as a crack.
One treated area per
length between bulkheads. Edge damage
not to exceed 0.025
inch depth and 0.75
inch length after
cleanup. One repair
per length between
bulkheads.
Dye penetrant
inspect bent
stringers not
requiring sectional removal
(after rework)
Change 2
REPAIRABLE
DAMAGE LIMITS
a. Damage Repairable by Patching:
cracks and smooth
contour dents less
than 1.0 inch depth
that are less than
0.50 stringer width
and do not extend
into radius, stringer
splice or bulkhead.
Longitudinal cracks
maximum 0.10 inch
width and1.0 inch
length.
b. Damage Repairable
by Insertion: Damage
exceeds limits for
patching but does
not exceed 3.0 inch
length after Cleanup.
One repair per length
between bulkheads.
Damage not to extend
into splice or bulkheads. If combined
stringer and skin
damage is present,
above limits and limits
for skin damage shall
not be exceeded.
2-110A
DAMAGE
REQUIRING
REPLACEMENT
Damage requires
more than one
insertion type repair between
bulkheads. Damage exceeds repairable limits or
repair does not
warrant time
expended.
TM 55-1520-234-23
Table 2-1. Tailboom Skin and Structure Classification of Damage (Cont)
ITEM
3. TAILBOOM
DOUBLERS.
DEFECT
DAMAGE
REPAIRABLE
DAMAGE LIMITS
NEGLIGIBLE
DAMAGE LIMITS
a. Cracks and
scratches.
Cracks and scratches
that are no deeper
than 20 percent of
double thickness, and
not exceeding one
inch in length can
be stop drilled at
each end provided
the crack is not closer
than one inch to any
adjacent structure.
b. Nicks, scratches,
and dents.
Nicks, scratches and
dents that are no
deeper than 20 percent of the doubler
thickness and not
exceeding one inch
in length, or 0.025
inch in width may
be polished out and
require no patching.
Holes, tears, and
other damage exceeding the limits in
steps a. and b.
above and no longer
c. Holes, tears,
and other damage
exceeding the
limits in a and
and b. above.
than 4.0 inches or
greater than 6.0
square inches may be
repaired by patching.
4. LONa. Cracks,
GERcorrosion, dents,
ONS
holes, tears,
(EXnicks, scratches,
CLUD- buckle or
ING
wrinkled.
TAILBOOM
ATTACH
FITTINGS)
See figure
a. Nicks and
Scratches: Not to
exceed 10 percent
of material thickness,
0.010 inch width
and 0.75 inch length
after cleanup.
Scratches in web
area that extend into
radius or at angle
greater than 45
degrees into critical
area, treat as a
crack (See figure
2-66A, detail B.)
a. Damage Repairable by Patching:
Smooth contoured
dents, length not
exceeding 1.0 inch
longitudinal, 0.5 inch
lateral and 0.050 inch
depth. If dent limits
are exceeded, treat
as a crack. (See
figure 2-66A, detail
A)
2. Nick and scratch
damage exceeds negligible limits but does
2-110B Change 2
REQUIRING
REPLACEMENT
a Damage exceeds
repairable limits
or two or more repairs required in
a single bay.
b. Damage other
than negligible.
occurs in a bay
containing either
a splice joint or
a previous repair.
c. Damage other2-66A
than negligible in
forward bay.
d. Splice required
TM 55-1520-234-23
Table 2-1. Tailboom Skin and Structure Classification of Damage (Cont)
ITEM
DEFECT
4. LONGERONS
(EXCLUDING
TAIL
BOOM
ATTACH
FITTINGS)
See figure
2-66A
DAMAGE
REQUIRING
REPLACEMENT
NEGLIGIBLE
DAMAGE LIMITS
REPAIRABLE
DAMAGE LIMITS
Nicks or notches
in flange area not
to exceed 0.80 inch
length, 0.04 inch
width and no deeper
than 10 percent of
material thickness
after cleanup. See figure 2-66A, Details B
and C.) No repair
closer than 1.0 inch
to a bulkhead, splice
or doubler. Refer to
attach fitting illustration for damage
limits to fittings.
not exceed 1.0 inch
width by 0.38 inch
height and does not
extend into critical
area after cleanup.
(See figure 2-66A,
detail F, section F-F.)
Damage in critical
area does not
exceed 2.0 inch
length and 0.40 inch
depth after cleanup.
See detail F, section
G-G.)
3. Crack, hole or
tear damage not exceeding limits of
figure 2-66A, Details
D and E, and extending no closer than
1.0 inch to a splice,
doubler or bulkhead
after repair.
in second bay.
e. Damage other
than negligible
comes closer than
1.0 inch to a
doubler, splice or
bulkhead.
f. Any longerons
damaged a sufficient amount to
cause permanent
buckles in tailboom,
sharp wrinkles in
skin or excessive
misalignment.
b. Damage Repairable by Insertion:
1. Repairable by
patching limits exceeded but less than
2.50 inch length after
cleanup. (See figure
2-66A, Details F and
G.)
2. Cracks or sharp
nicks in dent or
damage exceeds repair
by patching, but less
than 2.50 inch after
cleanup.
Requires replacement of both the
longerons and fitting.
Longerons are
replaced at Depot
Maintenance Level.
b. Corrosion: Less
than 10 percent of
material thickness
and not exceeding
an area 0.10 inch
width by 0.75 inch
cleanup. Damage
confined to web area
only and no closer
than 1.0 inch to a
splice, doubler or
bulkhead. One repair
for each longeron
in a bay area. No
damage in forward
bay. See figure 2-66A,
Detail B.)
Change 2
2-110C
NOTE
Damage in forward
bay area (other than
negligible).
TM 55-1520-234-23
Table 2-1. Tailboom Skin and Structure Classification of Damage (Cont)
ITEM
DEFECT
5. TAIL Corrosion, dents,
BOOM cracks, holes,
BULK- nicks and
HEADS wrinkles.
(Does not
include
canted
bulkhead)
See figure 2-66B
NEGLIGIBLE
DAMAGE LIMITS
Corrosion less than
10 percent of web
material thickness
and not exceeding
4.0 square inch after
cleanup. Damage no
closer than 0.260 inch
to a former, stiffener
or radius. Dents
nicks, scratches
in bulkhead web,
refer to skin damage
limits, item 1. Damage
in a radius treat
as a crack
2-110D
REPAIRABLE
DAMAGE LIMITS
a. Damage Repairable by Patching.
1. Corrosion
damage greater than
negligible but does
not exceed 0.70 inch
width or .33 percent
of a cross section after
cleanup. (See figure
2-66B, detail B.)
Damage no closer
than 0.50 inch to a
stiffener or attaching parts after cleanup.
2. Dent, cracks
holes and scratches
greater than negligible
but does not exceed
limits of figure 2-66B
Details A and B.
Maximum three
damages not to exceed limits of detail A
allowed for each bulkhead quadrant. Cracks
or damage in radius
of former on forward
bulkhead except in
area of attach fittings.
b. Damage Repairable by Insertion:
1. Corrosion damage
exceeds repairable
by patching but does
not exceed limits of
figure 2-66B, Detail C.
2. Dent, cracks or
hole damage exceeds
limits of figure
2-66B, Details A
and B, but less than
limits of Details C
and D.
Change 2
DAMAGE
REQUIRING
REPLACEMENT
Replace stiffeners
or any attaching
parts for damage
other than negligible.
Replace bulkhead
if repairable limits
are exceeded or if
more than one repair
to the limits of
figure 2-66B, Detail
D, is required.
NOTE
Bulkheads are re
placed by depot
maintenance.
TM 55-1520-234-23
Table 2-1. Tailboom Skin and Structure Classification of Damage (Cont)
ITEM
DEFECT
6. TAILCracks, holes,
BOOM nicks and
CANTED wrinkles.
BIJLKHEAD
NEGLIGIBLE
DAMAGE LIMITS
REPAIRABLE
DAMAGE LIMITS
Corrosion not to
exceed 1.0 square
inch for single
area, 4.0 square
inch total area
and 10 percent
material thickness
after cleanup. Nicks
and scratches not to
exceed 1.0 inch
length, 0.025 inch
width and 10 percent
material thickness
after cleanup. Treat
damage in radius as
a crack.
Three holes maximum
not exceeding 1.0
inch diameter in web
area and 3.0 inch
minimum distance
between damage.
Cracks in nutplate
hole but not extending
into radius. Cracks
in web area not exceeding 1.0 inch
length after cleanup.
No damage to come
closer than 0.50 inch
to stringer, longeron
or structure attaching point and no
closer than 1.0 inch
to a spar cap.
DAMAGE
REQUIRING
REPLACEMENT
Cracks or holes in
area of longeron,
stringer, or spar
cap attachment
points. Damage
exceeds repairable
damage limits.
b. Repair.
(1) Replace loose, missing or cocked rivets if no
other structural damage is present.
(2) Repair cracks, holes and tears less than three
inches in length as follows:
(a) Remove all the damaged skin and fabricate a
filler plate of the same material as the skin to match the
hole in the skin. Fabricate a backing patch of the same
material. See figure 2- 58.
(b) Rivet filler plate and backing patch in place.
Refer to TM55-1500-204-25/1 for standard repair
instructions.
(a) Stop drill cracks.
(4) Repair corrosion damage as follows:
(b) Smooth out edges of holes and tears.
(a) Polish out minor corrosion damage.
(c) Apply a lay-on patch of like material. See
figure 2-58. Install a minimum of four rivets on each side
of patch. Refer to TM 551500-204-25/1 standard repair
instructions.
(b) Apply chemical film (C37) to bare aluminum
surfaces.
(c) Prime repaired area with primer (C101).
(3) Repair cracks, holes and tears more than three
inches in length as follows:
Change 2
(d) Touch up paint to match surrounding area.
2-110E
TM 55-1520-234-23
Figure 2-58. Tailboom and elevator skins (Sheet 1 of 2).
2-110F Change 29
TM 55-1620-23423
ITEM
1.
2.
3.
4.
5.
6.
7.
S.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
25.
26.
27.
MATERIAL
SPECIFICATION
CONDITION
THICKNESS
5062 Al. Alloy
7075 Al. Alloy
7075 Al. Alloy
6061 Al. Alloy
Fiberglass
Fiberglass
7075 Al. Alloy
Kydex 100
Al. Faced Honeycomb Sandwich
Al. Faced Honeycomb Sandwich
Fiberglass
7075 Al. Alloy
7075 Al. Alloy
7075 Al. Alloy
7075 Al. Alloy
7075 Al. Alloy
7075 Al. Alloy
7075 Al. Alloy
7075 Al. Alloy
7075 Al. Alloy
7075 Al. Alloy
7076 Al. Alloy
7075 Al. Aloy
2024 Al. Alloy
2024 Al. Alloy
6061 Al. Alloy
2024 Al. Alloy
QQA250/8
QCLA250/13
QQA250/13
QQA260/11
TO
T6
T6
T6
0.040
0.032
0.032
0.040
QQA260/13
T6
0.025
Variable
Variable
QQA250/13
QQA250/13
QQA250/13
QQA250/13
QQA250/13
QQA250/13
QOA250/13
QQA250/13
QQA250/13
QQA250/13
QOA250/13
QQA250/13
QQA250/6
QQA250/6
QQA250/11
QQA250/6
T6
T6
T6
T6
T6
T6
T6
T6
T6
T6
T6
T6
T3
T3
TO
TO
0.050
0.050
0.032
0.032
0.032
0.025
0.025
0.025
0.025
0.025
0.040
0.040
0.040
0.040
0.050
209030-300-2A
Figure 2-58. Tailboom and elevator skins (Sheet 2 of 2)
2-47. Tailboom Assembly Doors and Panels .
Premaintenance Requirements for
Tailboom Doors and Panels
Condition
Model
Part No. or Serial No.
Special Tools
Test Equipment
Support Equipment
Minimum Personnel
Required
Consumable Materials
Special Environmental
Condition
Access doors, panels and fairing on the tailboom
make the tail rotor drive shaft, gearboxes, and
internal components of the tailboom accessible. See
figure 2-3, items 44 through 54 for view of doors, panels
and fairings.
Requirements
AH-1S Mod
All
None
None
None
a. Inspection. Inspect doors, panels and fairings
(44, 45, 46, 47, 48, 49, 50, 51, 52, 53 and 54, figure
2-3) for the following defects.
One
(C130)
(1) Cracks.
(2) Corrosion on metallic parts.
None
(3) Secure installation of fasteners and of hinges where
applicable.
(4) Deformity that causes improper fit.
Change 29
2-111
(6) Security and condition of isolation pad on 90
degree gearbox fairing. (See figure 2-58, Detail A.)
TM 55-1520-234-23
or miming hardware. If damage exceeds limits
of table 2-1, or requires use of jigs and fixtures
to repair, forward tailboom to depot level
maintenance for repair.
b. Removal. Remove doors, panels and fairings (44, 45,
46, 47, 48, 49, 50, 51, 52, 53 and 54, figure 2-3) as
follows:
The tailboom structure consists of bulkheads,
longerons, stringers, stiffeners, and doublers. See figures
2-59 through 246.
(1) Remove attaching screws or loosen turnlock
fasteners as applicable.
a. Inspection.
(5) Deteriorated or missing chafing stripe.
(1) Inspect tailboom structure for defects. Refer to
table 2-1 and figures 2-66A and 2-66B.
(2) Remove non-hinged panels.
(3) Disengage fasteners and remove hinge pins.
Remove hinged covers.
c.
Repair.
(1) Repair
204-25/1.
minor
cracks.
Refer
to TM55-1500-
(2) Polish out minor corrosion on aluminum parts.
Apply chemical film (C37) to bare metal surfaces. Touch
up with primer (C102) and paint to match surrounding
area.
(3) Replace missing and unserviceable turnlock
fasteners, hinges and screws.
(4) Replace access doors, panels and fairings that
are deformed to the degree that it does not fit when
installed.
(5) Replace deteriorated or missing chafing strips.
Use anti-chafe tape (C130) for driveshaft cover on
leading edge of fin.
d. Installation. Install doors, panels and fairings (44, 45,
46, 47, 48, 49, 50, 51, 62, 53, and 54, figure 2-3) as
follows:
WARNING
Any cracks to the longeron attach fittings
forward of boom station 70.00 or damage
exceeding the following limitations requires
depot maintenance.
(2) Inspect longeron attach fittings between boom
stations 37.37 and 70.00. Nicks, scratches, and gouges
may be polished with fine India stone (C126) provided
they do not exceed following limitations.
(a) Axial damage (parallel to bolt axis) must not
exceed 0.020 inch in depth or 3.00 inches in length.
(b) Radial damage (normal to bolt axis) must not
exceed 0.010 inch in depth or 3.00 inches in length.
(c) Nicks, scratches or gouges are not allowed
within one diameter of bolt hole, longeron splice rivets or
within 0.250 inch of end of longeron at splice.
(3) Inspect attachment bolt holes in tailboom and
fuselage fittings. Maximum diameter allowed is 0.5616
inch.
(4) Inspect attachment bolts for wear.
(1) Position hinged covers on tailboom and install
hinge pin. Close cover and engage fasteners.
b. Repair - General.
(2) Position non-hinged panels on tailboom and
engage fasteners.
(1) Cracks or damage resulting in distortion of
structural members is not repairable. Send tailboom to
depot level maintenance for repair.
2-48. Tailboom Assembly Structure.
(2) Defects within the limits of table 2-1 may be
repaired by the procedures in this section. If a
NOTE
Repair is limited to repair of minor cracks,
scratches, corrosion and replacement of loose
2-112 Change 29
TM 55-1520-234-23
Figure 2-59. Tailboom and synchronized elevator structure
Change 2
2-113
TM 55-1520-234-23
Figure 2-60. Bulkhead at boom station 59.50 (Sheet 1 of 2)
2-114
TM 55-1620-234-23
Figure 2-60. Bulkhead at boom station 59.50 (Sheet 2 of 2)
2-115
TM 55-1520-234-23
Figure 2-61. Bulkhead at boom station 80.44 (Sheet 1 of 3)
2-116
TM 55-1520-234-23
Figure 2-61. Bulkhead at boom station 80.44 (Sheet 2 of 3)
2-117
TM 55-1520-234-23
Figure 2-61. Bulkhead at boom station 80.44 (Sheet 3 of 3)
2-118
TM 55-1520-234-23
Figure 2-62. Bulkhead at boom station 101.38 (Sheet 1 of 3)
2-119
TM 55-1520-234-23
Figure 2-62. Bulkhead at boom station 101.38 (Sheet 2 of 3)
2-120
TM 55-1520-234-23
Figure 2-62. Bulkhead at boom station 101.38 (Sheet 3 of 3)
2-121
TM 55-1520-234-23
Figure 2-63. Bulkhead at boom station 122.33 (Sheet 1 of 2)
2-122
TM 55-1620-234-23
Figure 2-63. Bulkhead at boom station 122.33 (Sheet 2 of 2)
2-123
TM 55-1520-234-23
Figure 2-64. Bulkhead at boom station 143.28
2-124
TM 55-1520-234-23
Figure 2-65. Bulkhead at boom station 164.23
2-125
TM 55-1520-234-23
Figure 2-66. Bulkhead at boom station 186.18 (part no. 209-961-189-7)
2-126
TM 55-1520-234-23
Figure 2-66A. Longeron damage limits (Sheet 1 of 3)
Change 2
2-126A
TM 55-1520-234-23
Figure 2-66A. Longeron damage limits (Sheet 2 of 3)
Change 2
2-126B
TM 55-1520-234-23
Figure 2-66A. Longeron damage limits (Sheet 3 of 3)
Change 2
2-126C
Figure 2-66B. Typical tailboom bulkhead damage limits (Sheet 1 of 2)
Change 2
2-126D
TM 55-1520-234-23
Figure 2-66B. Typical tailboom bulkhead damage limits (Sheet 2 of 2)
Change 2
2-126E
TM 55-1520-234-23
specific procedure is not given for a repairable defect,
use standard repair procedures in TM 55-1500-204-25/1.
(3) Corrosion.
Polish out minor corrosion
damage to structural members. Apply chemical film
(C37) to bare metal. Prime repair area with primer
(C102).
(i) Secure reinforcing patch in position and
rivet into place.
(j) Apply a coat of primer (C102) over
repaired area.
(4) Replace worn attaching bolts.
c.
(h) Apply a coat of primer (C102) to both
sides of patch and damaged stringer.
Repair - Tailboom Stringers.
(1) Patching. Cracks, tears and punctures in
the stringer may be repaired by patching, provided they
do not extend more than one-half the width of the
stringer. Repair damaged stringer as follows:
(a) Check to see that rivets are not bent or
damaged and that rivet holes are not enlarged or torn.
(2) Insertion. Complete stringer breaks and
cracks extending more than one-half the width of the
stringer which make patching inadequate, necessitates
repair by insertion (splicing). Repair as follows:
(a) Check to see that rivets are not bent or
damaged and that the rivet holes are not elongated or
torn.
(b) Remove damaged or loose rivets.
(b) Remove damaged and loose rivets.
(c) Stop drill end of crack and, if necessary,
cut away damaged part taking care not to cut away
more than necessary.
(d) Re-form damaged stringer and other
displaced areas into correct position.
(e) Form a reinforcing patch of same
material and one gage heavier than damaged stringer.
The patch should extend at least four inches beyond
each end of cutout section. Maximum length of patch Is
12 inches. (Figure 2-66C)
(c) Trim damaged edge of break in stringer.
Do not trim more than necessary. Reform and return
damaged stringer to correct position.
(d) Cut and form an insert of same material
and gage as damaged stringer. Cut and form a
reinforcing patch of same material and one gage
heavier than damaged stringer. The patch should
extend at least four inches beyond each end of the
cutout section.
CAUTION
A filler splice should never exceed 12
inches in length.
(e) Clean dirt from around damaged area
and from both sides of insert and reinforcing patch.
Naphtha is extremely flammable.
Ground container before dispensing.
Use with adequate ventilation. Avoid
repeated skin contact.
(f) Clean dirt from around damaged area
and from both sides of reinforcing patch using a clean
cloth saturated with Naphtha (C88).
(f) Secure the insert and reinforcing patch
firmly in place. Drill rivet holes through reinforcing
patch, insert, and damaged stringer, the same size, and
pitch, as existing rivet holes. Remove burrs from all
holes.
(g) Apply a coat of primer (C102) to
damaged area on both sides of insert and patch.
(g) Secure reinforcing patch firmly in place
and drill rivet holes through patch and damaged stringer
the same size and pitch as existing rivet holes. Deburr
all holes.
Change 65
(h) Secure insert and patch. Rivet in place.
2-126F
TM 55-1520-234-23
d. Repair - Tailboom Longerons. Repair damaged
longeron aft of Boom Station 70.00 as follows:
(9) Secure reinforcing patch in position and
rivet in place while adhesive is still wet.
(10) Apply a coat of primer (C102) over the
repaired area.
e. Repair - Tailboom Bulkheads.
No repairs allowed forward of boom
station 70.00 other than longeron attach
fittings. Refer to paragraph 2-48a. (2).
(1) Check to see if there is any damage to skin
such as bent or damaged rivets or torn rivet holes.
(1) Patching. Cracks, tears, and punctures in
the bulkhead, web and flanges may be repaired by
patching, provided the damage does not extend more
than one-half the width of the bulkhead. Repair damage
as follows:
(2) Cut out damaged area, centering the cut
edges between holes to permit retention of existing rivet
pattern. Do not cut more than necessary. Leave a
generous radii at corners (0.250 inch minimum).
(a) Check that rivets are not bent or
damaged and that rivet holes are not elongated or torn.
(3) Cut and form a reinforcing patch of the
same material and one gage heavier than the longeron
and long enough to extend at least 4.50 inches on each
side or the damage (after cleanup). (Figure 2-66D).
(c) Stop drill end of crack, or if a tear or
puncture exists, cut away damaged part, taking care not
to cut away more thin necessary.
(4) Secure the reinforcing patch in position and
drill out rivet holes of the same size and pitch as in the
existing structure or as specified in TM 55-1500-20425/1.
(5) Mark a line around outer edge of patch
using a soft pencil. Remove patch and deburr holes.
(b) Remove damaged and loose rivets.
(d) Reform damaged member and other
displaced areas into correct position.
(e) Form a reinforcing patch of same
material and one gage heavier than damaged member,
and sufficiently long to give sturdy support. (Figure 266E).
Avoid breathing vapors.
Provide
adequate ventilation. Avoid prolonged
contact with the skin.
Naphtha is extremely flammable.
Ground containers before dispensing.
Use with adequate ventilation. Avoid
repeated skin contact.
(6) Remove paint from between previously marked
lines of damaged area using a clean cloth saturated with
methyl-ethyl-ketone (C87).
(f) Clean dirt from around damaged area
and from both sides of reinforcing patch using a clean
cloth saturated with naphtha (C88).
(7) Buff both sides of patch with Scotchbrite (C113) and
wipe with a clean cloth.
CAUTION
(g) Secure reinforcing patch firmly in place
and drill rivet holes through patch and damaged
member, and same size, and pitch as existing rivet
holes. Deburr all holes.
Do not touch patch with bare hands
after cleaning.
(h) Apply a coat of primer (C102) to both
sides of patch and damaged member.
(8) Apply adhesive (C12) to mating surface of
patch.
Change 65
2-126G
TM 55-1520-234-23
(i) Secure reinforcing patch in position and
rivet into place.
(3) Repair damaged bulkhead web - cracks,
tears and punctures as follows:
(j) Apply a coat of primer (C102) over
repaired area.
(a) Stop drill extreme ends of crack or cut a
round or elongated hole according to the length or shape
of crack, puncture, or tear in order to clean up ragged
edges and stretched metal. Allow generous radii at all
corners.
(2) Insertion. Complete bulkhead breaks, and
cracks, extending more than one-half the width of the
member, which make patching inadequate, must be
repaired by insertion (splicing).
(b) Cut and form a patch of same material
and thickness as damaged web.
(a) Check that rivets are not bent or
damaged and that rivet holes are not enlarged or torn.
(b) Remove damaged or loose rivets.
(c) Trim damaged edge of the break in
bulkhead. Do not trim more than necessary.
(d) Re-form and return damaged bulkhead
to correct position and contour.
(e) Cut and form an insert of same material
and gage as damaged bulkhead.
(f) Cut and form a reinforcing patch of
same material and one gage heavier than damaged
bulkhead, and sufficiently long to give sturdy support.
Use naphtha (C88) in a well ventilated area. Avoid
prolonged breathing of vapors and do not use in an area
with open flame or high temperature.
(c) Remove dirt from around damaged area
using clean cloth saturated with naphtha (C88).
(d) Secure patch in position and drill out a
double row of holes of same size and pitch as
surrounding areas. Remove patch and deburr holes.
(e) Apply a coat of primer (C102) to
damaged area and both sides of patch.
(f) Secure patch and rivet in place.
Use naphtha (C88) in a well ventilated
area. Avoid prolonged breathing of
vapors and do not use in an area with
open flame or high temperature.
(g) Clean both sides of insert and patch
with naphtha (C88). Position insert and patch to drill
rivet holes through patch insert and damaged area. Drill
holes same size and pitch as existing rivet holes.
Deburr all holes.
(g) Apply a coat of primer (C102) over
repaired area.
2-49. Tail Rotor Drive 90 Degree Gearbox Support
Fitting.
Premaintenance Requirements for 90
Degree Gearbox Support Fitting.
(i) Secure insert and reinforcing patch in
position and rivet into place.
Condition
Model
Part No. or Serial No.
Special Tools
Test Equipment
Support Equipment
Minimum Personnel
Required
Consumable Materials
(j) Apply a coat of primer (C102) over
repaired area.
Special Environmental
Condition
(h) Apply a coat of primer (C102) to both
sides of insert, reinforcing patch, and damaged
bulkhead.
Change 2
2-126H
Requirements
AH1S
All
None
None
None
One
(C12), (C87), (C102),
C112), (C113), (C137)
None
TM 55-1520-234-23
MATERIAL
1. Patches shall be of same material or one gauge heavier than damaged stringer.
2. For insertion repairs requiring like material, Bell Standard 110-001 must be used. Dash number is determined by
existing stringer size and thickness.
Figure 2-66C. Stringer Repair (Sheet 1 of 2)
Change 65
2-126J
TM 55-1520-234-23
CONDITION
This repair shall be used for cracks, tears, punctures, breaks in J-Stringer.
RESTRICTIONS
1. Maximum length of doubler is 12 inches.
2. Repair must not extend into bulkhead.
3. One repair per length between bulkheads.
Figure 2-66C. Stringer Repair (Sheet 2 of 2)
Change 65
2-126K
TM 55-1520-234-23
RESTRICTIONS
1. Only one repair may be made on each longeron in any one bay area.
2. No repairs allowed in forward bay.
3. Holes in longerons must not exceed 1.0 inches in diameter after cleanup.
Figure 2-66.D Longeron Material Chart
Change 65
2-126L
TM 55-1520-234-23
Figure 2-66E. Tailboom Structural Material
Change 65
2-126M
TM 55-1520-234-23
Figure 2-67. Damage limits - tail rotor drive support fitting (Sheet 1 of 4)
2-127
TM 55-1520-234-23
Figure 2-67. Damage limits - tail rotor drive support fitting (Sheet 2 of 4)
2-128
TM 56-1620-234-23
Figure 2-67. Damage limits - tall rotor drive support fitting (Sheet 3 of 4)
2-129
TM 55-1520-234-23
Figure 2-67. Damage limits - tail rotor drive support fitting (Sheet 4 of 4)
(2) Repair chafing damage that is within repair
limits as follows:
a. Inspection. See figure 2-67.
(1) Inspect fitting for nicks, scratches, sharp
dents and corrosion.
(2) Inspect fitting for chafing damage where
driveshaft cover and tail rotor gearbox cover contact the
fitting.
(3) Inspect fitting for worn bushings in bellcrank
support lugs.
(4) Inspect fitting for worn bushings in six holes
for gearbox studs.
NOTE
Chafed areas on the fitting, that are not
worn
beyond
minimum
material
thickness, may be built-up with
adhesive to form a new seat for the tail
rotor driveshaft cover and for the tail
rotor gearbox cover.
(a) Ensure that minimum material thickness
limits defined on Sheet 1 and Sheet 2 for chafing
damage have not been exceeded.
b. Repair. See figure 2-67.
(1) Polish out mechanical and corrosion
damage that is within repair limits. Use 400 grit
sandpaper (C112) or Scotchbrite (C113). Remove all
traces of corrosion damage. Ensure that damage limits
were not exceeded during polishing out procedure.
Touch up repair areas with primer (C102).
2-130
Provide adequate ventilation when
using methyl-ethyl-ketone (C87). Avoid
breathing solvent vapors and avoid
prolonged skin contact.
(b) Clean the chafed area with methyl ethyl-ketone
(C87).
(c) Polish chafed area with 400 grit sandpaper
(C112) or Scotchbrite (C113).
TM 55-1520-234-23
vator control linkage from the swashplate is attached to
the horn and controls the rotational movement. The two
elevators are mounted on the horn. Their position is
determined by relational movement of the horn.
a. Inspection.
(d) Apply adhesive (C12) to build-up chafed areas
to provide a new seat for the covers.
1 Build-up area B shown on Sheet 1 to a
thickness of 0.12 TO 0.15 inch in damage sector C shown
on Sheet 2.
2 Build-up area B shown on Sheet 1 to a
thickness of 0.25 TO 0.29 inch in damage sector D shown
on Sheet 2.
(1) Scratches within the following limits
reparable. Replace part if damage exceeds limits:
are
(a) Minor scratches on the elevator skins are
negligible if no crack damage is involved with the
scratches.
(b) Scratches less than 0.005 inch deep on the
horn (11, figure 2-68) are reparable by polishing out.
Refer to paragraph 11-3 for specific damage limits.
3 Build-up area A shown on Sheet 1 to a
thickness of 0.80 TO 0.82 inch.
(e) Allow adhesive applied in step (d) to dry
thoroughly. Apply two coats of primer (C102) to repair
area.
(2) Cracks, tears, holes and nicks within the
following limits are reparable. Replace part if damage
exceeds
limits.
(f) Touch-up repair area with paint to match
surrounding area.
(a) Cracks in elevator skin are reparable by
patching if they do not exceed the following limits:
(g) Apply tape (C137) on forward upper edge of
fitting where corner of drive shaft cover contacts fitting.
2-50. Tailboom Synchronized Elevator.
Minimum Personnel
Required
Consumable Materials
Special Environmental
Condition
On leading edge-6 inches long
2
Between spars-6 inches long
3
On tip-6 inches long.
(b) Holes, tears and cracks in elevator skins
are reparable by cutting out damaged area and patching
with insert plate and backup plate when the repaired area
is not over three inches in diameter. Also, no damage to
the tubular elevator spar is allowed.
Premaintenance Requirements for
Tailboom Synchronized Elevator.
Condition
Model
Part No. or Serial No.
Special Tools
Test Equipment
Support Equipment
1
Requirements
AH-1S
All
None
None
Torque wrench
Force Gauge
(Fish Scale)
(c) No mechanical damage is allowed on
supports (8).
(d) Nicks in elevator trailing edge that are less
than 0.25 inch deep are reparable by rounding and
polishing out.
Two
None
None
The synchronized elevator consists of the right and left
elevators, horn, supports and attaching parts. See figure
2-68. The elevators are built up on a tubular spar with
aluminum ribs and skin. The horn is mounted inside the
tailboom in supports which act as bearings and permit
rotational movement of the horn. The synchronized ele-
(e) Prior to removal of elevator,
inspect
support brackets (4) on both sides of tailboom for loose
attaching rivets. Inspect rivets visually and by hand
contact for signs of movement.
Replace loose,
damaged, or missing rivets.
NOTE
Apply a moderate force when moving the
elevator and use care not to bend the elevator
thus causing false indications.
Change 42 2-131
(f) Check axial play of elevator horn assembly (11)
in support assemblies (8) as follows:
1 Mount dial indicator inside of the tailboom
placing the stylus against the elevator horn at the pivot
point.
2 Move elevator inboard and outboard
(spanwise) and observe the total indicator reading. A
minimum of 0.005 inch and maximum of 0.030 inch play
should be indicated (figure 2-68).
3 If the indicator readings are not within
tolerance, adjust shims (7) as necessary.
(g) Check radial play as follows:
1 Mount dial indicator inside tailboom with
stylus in contact with the upper surface of elevator horn
near the inboard edge of pivot point.
TM 55-1520-234-23
2 Lightly move elevator up and down and
observe total reading on dial indicator. A maximum
reading of 0.010 inch is permissible (figure 2-68).
NOTE
Heavy force in moving the elevator will cause
flexing of elevator spar tube thus producing
false indications of excess radial play.
3 If dial indicator readings are not within
tolerance, adjust shims (16) as necessary.
(3) Corrosion.
Inspect elevators,
horn and
supports for corrosion damage.
Polish out minor
corrosion dam- age that does not exceed 0.003 inch in
depth. Maximum allowable corroded area is four square
inches in a single area or twenty percent of the elevator
skin ara.
(4) No damage to elevator spar (17, figure 2-68) is
allowed.
Change 42 2-132
TM 55-1520-234-23
Figure 2-68. Elevator installation
Change 42 2-132A
TM 55-1520-234-23
(4) Repair holes, tears and jagged cracks that are
within three inch diameter limit as follows:
b. Removal.
(1) Remove bolt (2, figure 2-68) and washer (3).
Slide elevator (1) outboard until elevator tubular spar is
clear of horn (11). Remove opposite elevator in same
manner.
(2) Remove access door on lower side of tailboom
below horn (11).
(3) Disconnect control tube (13) from horn (11).
(4) Remove four bolts (10) and washers (9).
NOTE
Ensure that shims (7), shims (16) and upper
and lower support assemblies are indexed
during accomplishment of the following step.
(5) Remove two nuts (14), washers (15) and screws
(5). Remove support assembly (8), shims (7) and shims
(16). Index these parts for reinstallation in the same
location.
(6) Remove support assembly on opposite end of
horn in same manner described in preceding step and
remove horn (11).
NOTE
Handle support assemblies (8) with care to
avoid damage to bearing surfaces.
c.
(a) Cutout the damaged area. Ensure that
tubular spar inside elevator has not been damaged if the
spar can be inspected in the area of the cutout.
(b) Fabricate a filler plate, to fit the cutout
prepared in the preceding step, from the same material
as the skin. See figure 2-58 for description of elevator
skin. Fabricate backup patch from the same material to
use with filler plate. Install filler plate and backup patch in
accordance with standard instructions in TM 55-1500-20425/1.
(5) Repair minor corrosion damage on elevators,
horn and supports. Polish out corrosion and apply primer
(C102). Touch up paint on elevators to match adjacent
area.
(6) Repair smooth dents by using a suction cup.
Only one pull per dent is allowed. Inspect for cracks after
repair. If only one crack is found, repair per paragraph 250c(3). If two or more cracks are found, send elevator to
depot for
repair.
(7) If dents are found in rivet pattern,
elevator to depot for repair.
send
d. Installation.
Repair.
(1) Polish out minor scratch damage on elevators.
Apply primer (C101) and touch up paint to match adjacent
area.
(2) Polish out scratch damage on horn that is within
limits. Apply primer (C102) to repair area. Refer to
paragraph 11-3.
(3) Repair elevators with crack damage that is
within limits for patch repairs as follows:
(a) Stop drill each end of crack.
Provide adequate ventilation when using
methyl-ethyl-ketone. Avoid breathing solvent
vapors and avoid prolonged skin contact.
NOTE
If solid film lubricant is accidentally removed,
it may be reapplied to the area from which It
was removed.
(b) Ensure that tubular spar inside elevator
has not been damaged if the spar can be inspected in the
area of the crack.
(c) Fabricate a patch from the same material
as the skin. See figure 2-58 for description of elevator
skin.
Install patch in accordance with standard
instructions in TM 55-1500-204-25/1.
(1) Clean the inside part of the horn (11, figure 268) that mates with the tubular spars of the elevators.
Also clean the tubular spars. If there is any zinc chromate
or similar material in these mating surfaces, clean down
to bare metal with methyl-ethyl-ketone (C87) and clean
cheesecloth
Change 65 2-132B
TM 55-1520-234-23
(C36). Do not use excess methyl-ethyl-ketone or allow it
to saturate the tubular spar as it may remove the solid
film lubricant. Allow the methyl- ethyl-ketone to dry.
(4) Position support assembly (8) and shim set (7)
on each end of horn. Install washers (9) and bolts (10)
but do not torque.
(5) Install shims (16) screws (5) aluminum washers
(6 and 15) and nuts (14).
(2) Deleted.
(3) Place the horn (11) inside the tailboom with the
arm that connects to control tube (13) toward the left side
of the tailboom. Position the horn control arm that
attaches to control tube (13) vertical and insert the left
end of the horn through tailboom bracket in left side of
tailboom. This bracket is similar to bracket (4). Raise the
right side of the horn, rotate the horn control arm
downward and insert the right end of the horn through
tailboom bracket (4).
Ensure that correct length bolts (10) are
installed. Refer to TM 55-1520-234- 23P.
NOTE
If no new parts are to be used when installing
the elevator, ensure that shims (7), shims
(16) and upper and lower sup- port assemblies
are installed in the same location from which
removed.
(6) Adjust thickness of laminated shim set (7) to
obtain 0.005 TO 0.030 inch lateral free play of horn.
(7) Adjust thickness of laminated shims (16) as
follows:
(a) Torque two nuts (14) on right end of horn
50 TO 70 inch-pounds, loosen similar nuts on left end of
horn. Attach a spring scale to horn control arm at the
point where control tube (13) attaches and check amount
of force required to rotate the horn in the support
assembly (8). Hold spring scale 90 degrees to control
arm while taking reading. Adjust thickness of shims (16)
until force required to rotate horn is 13 TO 16 pounds with
nuts (14) torqued 50 TO 70 inch-pounds.
(b) After shims (16) are properly adjusted on
right end of horn, repeat the procedure on the left end of
the-horn. The required spring scale reading is 26 TO 32
pounds when shims are properly adjusted on both ends of
horn and all nuts (14) torqued 50 TO 70 inch-pounds.
(8) Torque bolts (10).
Change 65 2-133
(9)
Apply coat of compound (C52) or (C53) to
surfaces of horn (11) contacted by elevator spars.
(10)
TM 55-1520-234-23
(b) Saturate fiberglass repair material with
epoxy resin (C107) and apply over cracked area.
(c) Allow adhesive to cure and touch up
paint to match adjacent area.
Install control tube (13).
(10A) Position elevator (1) tubular spar in horn
and secure with washer (3) and bolt (2). Torque bolts (2)
100 TO 140 inch-pounds.
(10B) Inspect elevator (1) and the opposite
elevator for adequate clearance with the external surface
of the tailboom while the elevators are moved through full
throw. If clearance is not approximately equal for both
elevators,
redistribute shim set (7) under support
assembly (8) and sup- port assembly (12) as necessary to
obtain equal clearance.
(2) Polish out minor corrosion damage. Apply
primer (C102) and touch up paint to match adjacent area.
(3) Replace
loose or missing inserts.
paragraphs 2-5 and 2-6 and figure 2-68A.
(4) Replace
honeycomb panels if damage
exceeds limits in figure 2-68A.
2-52.
(11)
Chapter 11.
Check rigging of elevator.
(12)
tailboom.
Install access door on lower side of
2-51.
Refer to
See
Tailboom Skid.
The tailboom skid is located at the aft end of the tailboom.
The purpose of the tail skid is to warn the pilot of a tail-low
attitude when landing.
a. Inspection.
Tailboom Fin.
The tailboom fin is an integral part of the tailboom.
It is made up of a spar, aluminum ribs and honeycomb
panels. See figure 2-59.
(1) Scratches and nicks. Minor surface scratches
and nicks are negligible and do not require repair. Slight
scratches and nicks require polishing out. Replace tail
skid if very deep scratches or nicks are present
(2) Dents. Smooth dents up to 0.062 inch deep are
negligible and do not require repair. Replace tail skid if
dents deeper than 0.062 inch are present.
a. Inspection.
(1) Cracks: No cracks are allowed in structure.
Minor cracks in ninety degree gearbox cover are
reparable.
(2) Buckled and wrinkled skin.
(3) Cracks.
present.
Replace tail skid if any cracks are
(4) Deformity. Replace tail skid if deformed (bent)
to the degree that it can be detected visually.
(3) Loose, cocked and missing rivets.
(5) Loose attachment to tailboom.
cause for loose attachment
(4) Loose and/or missing fasteners (inserts).
(5) Corrosion.
b. Removal.
(6) Voids, cracks, dents, and other damage in
honeycomb panels. See figure 2-68A for damage limits.
b. Repair. If repair requires use of jigs or fixtures,
forward tailboom to depot level maintenance for repair.
(1) Repair cracked ninety degree gearbox
cover as follows:
(a) Remove cover from helicopter.
Determine
(1) Remove access covers (50, figure 2-3).
(2) Remove nut, washer and bolt that attach
forward end of skid to tailboom and pull tail skid aft out
through support block.
c.
Repair.
(1) Polish out slight scratches and nicks.
If
complete clean up of damage results in removal of
enough material to weaken the tail skid, replace the
Change 48 2-134
tall skid. Apply primer (C102) to bare metal surfaces and
touch up paint to match surrounding surfaces.
(2) Replace tail skids that are cracked, deformed,
or have dents in excess of 0.062 inch limit.
TM 55-1520-234-23
204-030-947-17.
The-17 tailboom skid
installation includes a tube assembly filled with
lead shot. Ensure that the components of
tailboom skid installation P/N 204-030-947-13
are used on MOD S helicopters.
d. Installation.
(1) Position tail skid through support block and
install bolt, washer and nut to secure forward end of tail
skid to tailboom.
It is possible to erroneously install tailboom
skids on MOD S helicopters that are
components of tailboom skid installation P/N
(2) Install access covers (50, figure 2-3).
LIMITS
LIMITS
1. No sharp dents, holes, or damages that penetrate
metal facing.
2. Maximum diameter of damage 2.0 inches, or
maximum length of damage 1.50 inches.
3. Maximum depth of damage 20 percent of panel
thickness.
4. Total damage not to exceed 10 percent of a bay area.
5. Minimum distance of 0.5 inch from adjacent structure,
inserts or beveled edge.
1. Smooth, crack free dent.
2. Maximum diameter of single dent 1.0 inch. Two or
more dents in any 6.0 inch diameter area, consider
as one dent.
3. Maximum depth: 20 percent of panel thickness.
4. Maximum area of all dents combined: 10 percent of a
bay area.
5. Maximum of five dents in a 9.0 square inch area.
6. No voids may be present under tile damage.
7. Minimum distance of 0.5 itch from inserts or beveled
edge.
214020-2-2
Figure 2-68A. Vertical fin honeycomb panels damage limits (Sheet 1 of 3)
Change 7 2-134A
TM 55-1520-234-23
LIMITS
LIMITS
1. Maximum diameter of 3.0 inches after clean-up.
2. Maximum of three patch repairs in a panel. Damage
after clean-up comes no closer than 1.5 inch to a
similar repair or insert and no closer than 1.5 inches
to a beveled edge.
1. Maximum diameter of hole 3.0 inches, after cleanup.
2. Minimum distance from structural members or other
repair: 2.0 inches.
3. Replace panel if water or corrosion found in core.
3. Minimum distance of completed repair from an edge
bevel: 0.50 inches.
4
4. Total damage not to exceed 10 percent of a bay area.
Total damage not to exceed 10 percent of total panel
area or 25 percent of a single bay area after clean-up.
5. Maximum of three patch repairs in a panel.
6. Replace panel if water or corrosion found in core.
214020-2-3
Figure 2-68A. Vertical fin honeycomb panels damage limits (Sheet 2 of 3)
Change 2 2-134B
LIMITS
DAMAGED OR LOOSE INSERTS
1. Maximum total void are not to exceed 5 percent of
panel surface area.
LIMITS
2. Maximum area of a single void: 1. square inch
and a minimum of 2.0 inches between voids.
Maximum length of a void: 3.0 inches in any
direction.
3. Damage not closer than 1.0 inch of a beveled
edge, hole or adjacent structure, or within 3.0
inches of an insert. Void in area of insert limited
to 0.62 square inch with no damage to insert.
1. Remove insert by counter boring
enlarging hole size in panel facing.
without
2. Original hole diameter in panel facing must be
maintained in the replacement process.
3. No damage in area adjacent to insert.
Figure 2-68A. Vertical fin honeycomb panels damage limits (Sheet 3 of 3)
Change 2 2-134C/(2-134D blank)
214020-2-4A
TM 55-1520-234-23
Section III. PYLON
2-53.
a. Inspection. Inspect support for cracks. corrosion
damage, and secure attachment to lift beam.
Pylon Support Installation.
The pylon support installation consists of the
provisions for mounting the transmission in the airframe,
see figure 2-69. Major components of the pylon support
installation are as follows:
b. Removal. See figure 2-70. (1) Remove hydraulic
lines and fittings from support (1). Cap or plug hydraulic
lines to prevent entry of foreign material.
(2) Remove nut, screw and clamp from outboard
side of support (1).
a. Four transmission mount assemblies.
b. Two damper assemblies.
c.
(3) Carefully remove six rivets which secure
support to lift beam and remove support. Do not elongate
holes in lift beam.
Two damper fittings.
d. One fifth mount support fitting assembly.
c. Repair. Fabricate a new support in accordance with
instructions in appendix D. Supports presently installed
may be either 0.040 inch or 0.050 inch thick. Fabricate
the new support from 0.050 inch thick material as shown
in appendix D.
e. One lift beam assembly.
2-54.
Transmission Mounts.
Refer to Chapter 6.
2-55.
d. Installation. See figure 2-70.
Transmission Mount Dampers.
(1) Position support (1) on lift beam in original
position so that hydraulic fittings can be reinstalled and
clamp in place. Drill out holes for six rivets to match
holes in lift beam. Remove support and deburr holes.
Position support on lift beam and install six rivets (item
53, table 2-1).
Refer to Chapter 6.
2-56. Transmission Mount Bushings.
Refer to Chapter 6.
2-57.
(2) Install hydraulic fittings on support and install
hydraulic lines on fittings.
Fittings and Supports.
Refer to paragraphs 2-58 through
instructions to repair or replace supports.
2-58.
2-62
Support (1, figure 2-70). (AVIM)
Premaintenance Requirements for Support.
Condition
Model
Part No. or Serial No.
Special Tools
Test Equipment
Support Equipment
Minimum Personnel
Required
Consumable Materials
Special Environmental
Condition
Requirements
AH-1S
All
None
None
Hydraulic test stand
Riveting equipment
One man
None
for
(3) Install clamp removed in step b above.
(4) Perform functional check of hydraulic system
with hydraulic test stand or by ground run of helicopter.
Check for correct operation of hydraulic system and for
hydraulic fluid leaks.
2-59.
Support Assembly (2, figure 2-70). (AVIM)
Premaintenance Requirements for
Support Assembly.
Condition
Model
Part No. or Serial No.
Special Tools
Test Equipment
NA
Change 22 2-135
Requirements
AH-1S
All
None
None
TM 55-1620-234-23
Figure 2-69. Pylon support installation
2-136
TM 55-1520-234-23
Figure 2-70. Hydraulic fitting supports PIN 209-030-267-11 and 209-030-267-29 installation
2-137
TM 55-1520-234-23
(3) Clean mating surfaces at support assembly (2)
and radius block (4) with 400 grit sandpaper (C112).
Remove all residue with methyl-ethyl-ketone (C87).
Premaintenance Requirements for
Support Assembly (Cont).
Condition
Support Equipment
Minimum Personnel
Required
Consumable Materials
Requirements
Hydraulic test stand,
Riveting equipment
(4) Mix a small quantity of adhesive (C17)
according to directions on container and apply a thin coat
of adhesive to mating surfaces of support assembly and
radius block.
One man
(C17), (C87),
(C112), (C102)
Special Environmental
Condition
NOTE
NA
Pot life of adhesive (C17) is 30 to 50 minutes.
Cure time is 24 hours at 70 to 90 degrees F.
Maximum strength is attained in 6 or 7 days.
a. Inspection. Inspect support assembly for cracks,
corrosion damage, and secure attachment to lift beam.
b. Removal. See figure 2-70.
(1) Remove hydraulic lines and fittings from support
assembly (2). Cap or plug hydraulic lines to prevent entry
of foreign material.
(2) Carefully remove twelve rivets which secure
support assembly (2) and radius block (4) to lift beam. Do
not elongate rivet holes in lift beam.
c.
(5) Place a four mil glass yarn string in bond line at
one inch intervals. The glass yarn will serve as a spacer
to ensure that adhesive thickness will be 3 to 8 mils after
curing.
(6) Clamp support assembly (2) and radius block
(4) in position on lift beam with radius block radius nested
in support assembly radius. Install twelve rivets (item 53,
table 2-2).
(7) Clean all adhesive squeeze out from the parts
before adhesive hardens.
Repair. See figure 2-70.
(1) Do not repair a damaged support assembly.
Procure a new support assembly through normal supply
channels.
(2) Remove radius block (4) from old support
assembly or fabricate a new support in accordance with
instructions in appendix D.
d. Installation. See figure 2-70.
(1) Clamp support assembly (2) and radius block
(4) on lift beam at position illustrated. Ensure that radius
block (4) is nested against mating radius of the support
assembly.
(2) Drill out rivet holes in new support assembly (2)
and radius block to match holes in lift beam. Remove
support and radius block (4). Deburr holes.
(8) Paint support assembly (2) and radius block (4)
with primer (C102).
(9) Install hydraulic fittings on support (2) and
install hydraulic lines on fittings.
(10)
Perform functional check for correct
operation of hydraulic system with hydraulic test stand or
by ground run of helicopter. Check for correct operation
of hydraulic system and for hydraulic fluid leaks.
2-60.
Pylon Damper Fittings.
One pylon damper fitting (11, figure 2-69) is
installed below each of the two dampers (8). The fittings
form the structural attachment point for the lower end of
the dampers.
Premaintenance Requirements for
Pylon Damper Fittings.
Provide adequate ventilation when using
methyl-ethyl-ketone (C87). Avoid breathing
vapors and avoid prolonged skin contact.
Condition
Model
Part No. or Serial No.
Special Tools
2-138
Requirements
AH-1S
All
None
TM 55-1520-234-23
(4) If rivet holes in pylon damper fitting are elongated,
procure a new fitting through normal supply channels.
Premaintenance Requirements for
Pylon Damper Fittings (Cont).
Condition
Test Equipment
Support Equipment
Minimum Personnel
Required
Consumable Materials
Special Environmental
Condition
(5) If rivet holes in helicopter structure for rivets (1,
figure 2-71) are not elongated, proceed to step (10). If
any holes in structure are elongated, install bushings as
outlined in steps (6) through (9).
Requirements
None
Roll staking tool
Two
(C102)
(6) Drill out elongated holes in web (3, figure 2-71) and
extrusion (4) as shown in figure 2-71. Use a letter size N
twist drill. Ream the hole for a class FN2 fit with a 77-331 bushing. Make the hole 0.0004 TO 0.0014 inch
smaller than the bushing.
None
a. Inspection. Inspect damper fittings (11, figure 269) for the following defects:
(2) Corrosion. Severe corrosion damage is cause for
replacement of damper fittings. Minor corrosion damage
may be repaired.
(7) Coat bushing (2, figure 2-71) and the hole with
primer (C102) and press bushing into position while
primer is wet. Install the bushing with the flanged end on
the opposite side of the structure from fitting (6) as
illustrated. The bushing must extend through the web (3)
and extrusion (4). Face off bushing flush with extrusion
(4) as shown on detail view A.
(3) Secure installation of all rivets which attach the two
fittings (11, figure 2-69) to the structure.
(8) Repair all elongated holes as described in steps (6)
and (7).
(1) Cracks. No cracks are acceptable.
(4) Secure installation of bearing in damper fitting and
condition of bearing. Maximum allowable radial play
0.006 inch.
(9) Touch up bare metal with primer (C102).
b. Removal. (AVIM)
(1) Remove transmission. Refer to Chapter 6.
(2) Remove the two aft transmission mount assemblies
(4, figure 2-69). Refer to Chapter 6.
(3) Carefully drill out rivets which secure damper
fittings to structure and remove fitting. See figure 2-71 for
detail view of rivet. Remove fitting.
c.
Do not replace loose or missing rivets with steel
fasteners unless it is a Hi-Lok rivet (HL20PB865-6 (forward) HP20PB86-6-6 (aft)). Rivets are
designed to shear before doing excessive
damage to pylon.
d. Installation (AVIM).
(1) Position fitting in helicopter and install rivets (1,
figure 2-71). Use one of the rivets or the Hi Lock fastener
listed in the preceding "Caution."
Repair. (AVIM).
(1) Replace damper fitting if cracked.
(2) Polish out superficial corrosion damage and apply
primer (C102) to bare metal surfaces.
(3) If bearing in pylon damper fitting failed to meet
inspection requirements, remove old bearing and install
new bearing by roll staking method. Refer to Chapter 5 for
general instructions to remove and install roll staked
bearings. If roll staking tools are not available, procure a
new fitting through normal supply channels.
(a) If the Hi-Lok rivet is installed, remove the two
forward 5/32 inch rivets and the two aft 3/16 inch rivets.
Exercise extreme caution to preclude further enlargement
of original hole.
(b) Ream original 5/32 inch holes 0.1615 to 0.1635
and install HL20PB6-5-6 Hi-Lok pin with HL86PB-5
collars.
(c) Ream original 3/16 inch holes 0.1885 to 0.1895
and install HP20PB6-6-6 Hi-Lok pin with HL86PB-6
collars.
Change 48 2-139
TM 55-1520-234-23
(d) One flat washer (AN960 series or equivalent) may be used under the collar with the nut if the Hi-Lok rivet
shows 1/8 inch or more of unthreaded shank.
(2) Install damper (8, figure 2-69), mount assembly (4) and transmission. Refer to Chapter 6.
(3) Perform ground run for functional check of flight controls and hydraulic system components affected by
transmission removal/installation.
Change 29 2-140
TM 55-1520-234-23
Figure 2-71. Pylon damper fitting supporting structure-repair of elongated holes
2-61.
Flight Control Power Cylinder Support.
The flight control power cylinder supports are the
structural mounting points for the three flight control
hydraulic cylinders.
(3) Replace supports that have damage in excess of
inspection limits.
d. Installation. Refer to Chapter 11.
2-62.
a. Inspection. Inspect each of the three supports for
mechanical damage, corrosion, thread damage and
elongation of holes (bore damage) in excess of limits
shown on figure 2-72.
Fifth Mount Support Fitting Assembly.
a. Inspection.
Inspect fifth mount support fitting
assembly (13, figure 2-69) for damage in excess of limits
shown on figure 2-73.
b. Removal. Refer to Chapter 11.
b. Removal-Refer to Chapter 6.
c.
c.
Repair.
(1) Polish out minor corrosion and mechanical damage
that does not exceed inspection limits. Do not remove
more material than necessary to blend repair smoothly
into surrounding surface. Use fine or medium grades of
sandpaper (C112) or crocus cloth (C45). Do not use
grinding wheels. Polish out mechanical damage only
deep enough to remove traces of damage. Polish out
corrosion damage to twice depth of the deepest pit.
(2) Touch up repair area on aluminum parts with
chemical film (C37) and primer (C102).
Repair.
(1) Polish out mechanical and corrosion damage that is
within limits shown on figure 2-73.
(2) Replace support fitting if damage exceeds limits or
if any cracks are detected.
(3) Touch-up repair areas with chemical film (C37) and
primer (C102).
d. Installation-Refer to Chapter 6.
Change 29 2-140A/(2-140B blank)
TM 55-1520-234-23
Figure 2-72. Damage limits-flight control power cylinder supports (Sheet 1 of 2)
Change 7 2-141
TM 55-1520-234-23
Figure 2-72. Damage limits-flight control power cylinder supports (Sheet 2 of 2)
Change 7 2-142
TM 55-1520-234-23
Figure 2-73. Damage limits-Fifth mount support fitting assembly
Section IV. WING
2-63.
Wing. See figure 2-74.
Stub wings, mounted on the fuselage, supply
additional lift and provide mounting accommodations
for weapons pylons . The structure ; built up with
aluminum alloy spars and ribs covered with sheet
aluminum skin. Each wing is attached to fuselage fittings
with attaching bolts. Four removable panels allow access
to internal provisions.
a. Removal. See figure 2-74.
NOTE
The removal procedure is the same for both
wings.
(1) Remove external stores from weapon pylon, if
Installed. Refer to Chapter 16.
(2) Remove lower access panel (6) on each wing.
(3) Disconnect two pylon hydraulic connections
(5) in each wing. Cap all open lines.
Change 65 2-142A
TM 55-1520-234-23
Support wings to prevent bolts from binding.
Remove bolts in sequence as shown in figure
2-74, detail A.
Provide adequate ventilation when using
methyl-ethyl-ketone (C87). Avoid breathing
solvent vapors and avoid prolonged skin
contact.
(4) Remove five attachment bolts (3) and
separate wing from fuselage.
(5) Disconnect electrical connection (4)
between wing root and fuselage.
b. Inspect wing fitting bushings for damage and for
wear beyond limits of figure 2-75. Bushings worn beyond
limits must be replaced by depot level maintenance.
Inspect barrel nuts for adequate self-locking feature.
Inspect wing bumper. Refer to paragraph 2-65.
Provide adequate ventilation when using
methyl-ethyl-ketone (C87). Avoid breathing
solvent vapors and avoid prolonged skin
contact.
b. Inspect washers (8) for looseness or damage.
Remove loose or damaged washers. Use a heat lamp or
heat gun and apply 200°F (93°C) maximum to soften old
adhesive. Use care not to damage fitting. After removing
washer clean fitting with a plastic scraper and methylethyl- ketone (C87). Wipe dry with clean cloth (C36).
c.
Installation. See figure 2-74.
NOTE
Installation procedure is the same for both
wings.
(1) Install and align barrel nuts in fittings.
(2) Connect and lockwire electrical connections (4).
Change 7 2-142B
TM 55-1520-234-23
(6) Install lower access panel (6).
(2.1)
If washers (8) were removed in step b.1,
install washers as follows:
(a) Clean fitting and washer with methyl-ethylketone (C87). Wipe dry with cloth (C36).
(7) Attach weapons pylon to wing, if required.
Refer to Installation-Outboard Ejection Rack and
Installation Inboard Ejector Rack, Chapter 16.
(b) Apply a small quantity of adhesive (C17) to
chamfered edge of washer and install wet washer into
fitting. Proceed immediately with steps (4), (5), and (6).
(8) Perform functional check of hydraulically
actuated articulated pylon and all electrical circuits in the
wing. Refer to paragraph 9-91.
NOTE
Excess adhesive must be wiped from fitting and washer to
prevent bonding bolt to washer.
2-64. Wing Skins and Panels. See figures 2-76 and
figure 2-77.
(3) Mount wing in fuselage fittings.
a. Inspection. Inspect the wing skins and panels for
the following defects.
(3.1)
Apply anti-seize compound (C26) to bolt
shank and install bolts (3). If MWO 55-1520-244-30-3 has
been incorporated, tiedown fitting (9), washer (10), and
washer (11) must be installed on each of the three lower
bolts (3).
(1) Cracks, holes and tears in the skin and/or
panels. If damage is less than 1.25 inches long and does
not involve damage to the structure, it is reparable by
patching.
(3.2)
follows:
(2) Corrosion.
reparable.
After MWO 55-1520-244-30-3,
torque as
(a) Torque two front bolts 400 inch-pounds.
Torque two center bolts 400 inch-pounds.
(b) Back off to zero torque or until threads are
disengaged.
(c) Gradually retighten bolts until contact
occurs between bolt head and washer or until torque
begins to increase. Note contact torque level.
(d) Apply an additional 100 inch-pounds of
torque above the contact torque to ensure a snug fit, but
do not exceed 450 inch-pounds torque.
NOTE
Some bolts may have a drilled head. Do not lockwire.
(e) Torque aft bolt 80 to 100 inch-pounds.
(4) Before MWO 55-1520-244-30-3, torque two
front bolts 100 TO 150 inch-pounds, two center bolts 100
TO 150 inch-pounds, and aft bolt 80 TO 100 inchpounds.
Minor
corrosion
damage
is
(3) Distortion. Inspect for wrinkles and buckled
skin. If this type damage is detected, inspect the wing
structure for damage.
Damage involving the wing
structure is not reparable at AVIM level.
(4) Rivet condition. Inspect for loose, cocked
and/or missing rivets. Rivet damage may be repaired by
replacing rivets if damage does not involve structural
repair.
b. Repair.
(1) Repair crack, tear and hole damage in skin by
patching as shown on figure 2-78.
(2) Polish out minor corrosion damage, apply
primer (C102) and touch up paint to match surrounding
area.
(3) Replace wing if skin is wrinkled and buckled to
the degree that the wing internal structure is involved.
(4) Replace loose, cocked or missing rivets if no
other structural damage is present.
(5) Remove caps from hydraulic connections (5)
and connect.
2-65.
Change 54 2-143
Wing Bumpers. See figure 2-74.
TM 55-1520-234-23
Figure 2-74. Wing installation
Change 16 2-144
TM 55-1520-234-23
a. Inspect wing bumper (7) for the following defects:
(1) Cuts and tears severe enough to affect
function.
(2) Debonding.
(3) Deterioration.
b. Replace damaged bumper.
(1) Remove wing if not previously accomplished. Refer to paragraph 2-63.
(2) Remove faulty bumper from wing with sharp plastic scraper. Clean residual particles of, bumper and adhesive
with naphtha (C87). Wipe the area dry with a clean cloth.
Change 23 2-144.1/(2-144.2 blank)
TM 55-1520-234-23
Figure 2-75. Limits chart-bushings
2-145
TM 55-1520-234-23
Figure 2-76. Wing skins, doors and doublers (Sheet 1 of 2)
Change 2 2-146
TM 55-1520-234-23
Figure 2-76. Wing skins, doors and doublers (Sheet 2 of 2)
Change 2 2-146A/(2-146B blank
TM 55-1520-234-23
Figure 2-77. Wing structure
(4) Apply a thin coat of adhesive (C14) to
mating surfaces of wing and new bumper. Allow the
adhesive to dry to tacky stage evidenced by adhering
but not transfering to the finger when touched.
(5) Position bumper on the wing. Start at
one edge of wing and roll or press the bumper firmly
against the wing. Remove excessive adhesive with a
cloth dampened with methyl-ethyl-ketone (C87). Allow
adhesive to cure for four hours minimum.
Provide adequate ventilation when
using methyl-ethyl-ketone
(C87).
Avoid breathing solvent vapors and
avoid prolonged skin contact.
(6)
paragraph 2-63.
(3) Sand the area where the new bumper
will be installed with 400 grit sandpaper (C 112).
Remove residue of sanding with methyl-ethyl-ketone
(C87).
2-147
Install wing on helicopter.
Refer to
TM 55-1520-234-23
Figure 2-78. Wing skin repair.
SECTION V.
(Deleted)
Pages 2-149 thru 2-172, including figure 2-79, deleted
2-148
Change 22
TM 55-1520-234-23
Section VI. CORROSION CONTROL
2-67. Corrosion.
same type on non-clad alloy parts is serious.
Corrosion is usually attributed to two factors:
location of helicopter operations and fabrication process
of parts. Corrosion is caused by presence of salt in
moist air, certain chemicals in water, elements in the
metal, treatment of parts, and contact of dissimilar
metals. Corrosion will not normally be as prevalent on
painted, clad, or plated surfaces as untreated surfaces.
However, corrosion can attack painted parts as moisture
can penetrate enamels, lacquers, and primer. Corrosion
on painted parts is usually characterized by a scaly or
blistered appearance, and sometimes by discoloration of
paint. Corrosion on clad or plated parts is recognized by
a dulling and pitting of the surface and is sometimes
accompanied by a whitish or reddish powdery deposit.
The extent and forms of corrosion may be determined
by examination and visual inspection.
A pointed
instrument may be used to make the test. Care should
be taken to avoid further damage. In some cases the
area must be cleaned to remove scales and powdery
deposits before examination can be made. See table 21A for forms and types of corrosion and cleaning and
treating methods.
b.
Electrolytic Corrosion. There are two major
causes for this type corrosion.
Contact between
dissimilar metals and condensation. When dissimilar
metals come in contact with each other with moisture
present, an electrical current flows between the metals
producing chemical by-products that dissolve one of the
metals. Corrosion caused by condensation is a result of
exhaust gases, battery acid, etc., contacting the metal.
a.
Superficial Corrosion. This type is the least
serious on alclad parts. After deposits are removed, an
etching will be noticed which results in the clad surface
having a series of hills and valleys. Provided the
etching has not reached the core, the effect on the
strength of the metal is negligible. Corrosion of this
c.
Intergranular Corrosion.
This form of
corrosion is not easily detected. It is caused by
imperfect heat treatment and occurs mostly in unclad
structural alloy parts. It is the most dangerous form of
corrosion for sheet stock because the strength of the
metal can be lowered without visible surface indications.
d.
Stress Corrosion. This form of corrosion is
caused by the action of sustained tension stresses in the
presence of a corrosive environment.
e.
Hygroscopic Material Corrosion. This form
of corrosion is caused by such materials as sponge
rubber, felt, cork, etc., absorbing water and holding it in
contact with the part.
2-68. Corrosion - Removal and Treatment.
Procedures for repairing corroded surfaces are
given in the following table.
Table 2-1A. Treatment of Corroded Surfaces
METAL
FORM
CLEANING
TREATMENT
Aluminum
Alclad
Surfaces
Mild or heavy
pitting, staining
and superficial
etching
Apply biodegradable
clearing compound
(C39) and rinse
with water. Do not
use abrasives.
Apply paint as
required. On
internal surface
use one coat
of zinc chromate
primer, (C102).
Change 38
2-173
TM 55-1520-234-23
Table 2-1A. Treatment of Corroded Surfaces (Cont)
METAL
FORM
CLEANING
TREATMENT
Paint unfinished
external surfaces
with aluminum
pigmented lacquer
(C78).
Apply paint as
required. On
internal surfaces
use one coat
of zinc chromate
primer, (C102).
Paint unfinished
external surfaces
with aluminum
pigmented lacquer
(C78).
Mild surface
pitting, staining
and superficial
etching
Clean with
biodegradable
cleaning compound
(C39). Remove
products of corrosion
with scotch-brite
(C 113).
Aluminum
Intergranular
corrosion
Remove corroded
area. Burnish part
to remove sharp
edges.
Treat with a
five percent
solution of
Potassium
Dichromate (C98.1)
and allow to
dry. Brush off
excess crystals.
Spray with zinc
chromate primer,
(C102).
Magnesium
Surface pitting
Remove corrosion
with a stiff
bristle brush.
Apply chromepickle solution
(C38) for one
minute. Rinse with
fresh water.
Steel
Lightly rusted
parts. No pitting.
Clean parts with
biodegradable
cleaning compound
(C39) and rinse
with fresh water.
Use steel wool to
remove compound,
if necessary.
Apply a coat of
zinc chromate
primer, (C102)
on previously
cadmium plated
parts.
Badly rusted.
Not applicable.
Replace parts.
2-174
Change 38
TM 55-1520-234-23
Section VII. STRUCTURAL REPAIR MATERIALS
Table 2-2. Structural Repair Materials.
ITEM
NO.
1
DESCRIPTION
REF. NO.
AND FSCM
NSN
Aluminum Alloy Sheet,
0.010 inch thick, 2024-T3
Aluminum Alloy Sheet,
0.012 inch thick, 2024-T3
QQ-A-250/5
(81348)
QQ-A-250/5
(81348)
3
Aluminum Alloy Sheet,
0.016 inch thick, 2024-T3
QQ-A-250/5
(81348)
9535-00-232-0543
4
Aluminum Alloy Sheet,
0.020 inch thick, 2024-T3
QQ-A-250/5
(81348)
9535-00-167-2277
5
Aluminum Alloy Sheet,
0.025 inch thick, 2024-T3
QQ-A-240/5
(81348)
9535-00-167-2278
6
Aluminum Alloy Sheet,
0.032 inch thick, 2024-T3
QQ-A-250/5
(81348)
9535-00-086-9729
7
Aluminum Alloy Sheet,
0.040 inch thick, 2024-T3
QQ-A-250/5
(81348)
9535-00-167-2280
8
Aluminum Alloy Sheet,
0.050 inch thick, 2024-T3
QQ-A-250/5
(81348)
9535-00-232-0569
9
Aluminum Alloy Sheet,
0.060 inch thick, 2024-T3
QQ-A-250/5
(81348)
10
Aluminum Alloy Sheet
0.070 inch thick, 2024-T3
QQ-A-250/5
(81348)
11
Aluminum Alloy Sheet,
0.080 inch thick, 2024-T3
QQ-A-250/5
(81348)
12
Aluminum Alloy Sheet,
0.100 inch thick, 2024-T3
QQ-A-250/5
(81348)
9535-00-288-0675
13
Aluminum Alloy Sheet,
0.020 inch thick, 5052-
QQ-A-250/8
(81348)
9535-00-832-1868
14
Aluminum Alloy Sheet,
0.025 inch thick, 5052-
QQ-A-250/8
(81348)
9535-00-832-1868
15
Aluminum Alloy Sheet,
0.040 inch thick 5052-
QQ-A-250/8
(81348)
9535-00-232-6864
2
Change 38
2-175
9535-00-167-2274
9535-00-232-0398
TM 55-1520-234-23
Table 2-2. Structural repair materials (Cont)
ITEM
NO.
REF. NO.
AND FSCM
DESCRIPTION
NSN
16
Aluminum Alloy Sheet,
0.025 inch thick, 6061-T6
QQ-A-250/11
(81348)
9535-00-250-6502
17
Aluminum Alloy Sheet,
0.032 inch thick, 6061-T6
QQ-A-250/11
(81348)
9535-00-085-4133
18
Aluminum Alloy Sheet
0.010 inch thick, 7075-T6
QQ-A-250/13
(81348)
19
Aluminum Alloy Sheet
0.012 inch thick, 7075-T6
QQ-A-250/13
(81348)
9535-00-236-7091
20
Aluminum Alloy Sheet,
0.016 inch thick, 7075-T6
QQ-A-250/13
(81348)
9535-00-084-4438
21
Aluminum Alloy Sheet,
0.020 inch thick, 7075-T6
QQ-A-250/13
(81348)
9535-00-086-9808
22
Aluminum Alloy Sheet,
0.025 inch thick, 7075-T6
QQ-A-250/13
(81348)
9535-00-086-9864
23
Aluminum Alloy Sheet,
0.032 inch thick, 7075-T6
QQ-A-250/13
(81348)
9535-00-249-5811
24
Aluminum Alloy Sheet,
0.040 inch thick, 7075-T6
QQ-A-250/13
(81348)
9535-00-084-4581
25
Aluminum Alloy Sheet
0.050 inch thick, 7075-T6
QQ-A-250/13
(81348)
9535-00-086-9465
26
Aluminum Alloy Sheet
0.063 inch thick, 7075-T6
QQ-A-250/13
(81348)
9535-00-088-6599
27
Aluminum Alloy Tubing (4" Dia.)
0.083 thickness, 2024-T3
QQ-A-300-3B
28
Magnesium Alloy
AMS4350
29
Rivet, Blind, Flush Head
CR2263-4-1
(11815)
30
Rivet, Blind, Flush Head
CR2248-4
(11815)
31
Rivet, Blind, Flush Head
CR2248-6-3
32
Rivet, Universal
CR2249-6-3
(11815)
2-176
Change 38
5320-00-916-9534
TM 55-1520-234-23
Table 2-2. Structural repair materials (Cont)
ITEM
NO.
.
REF. NO.
AND FSCM
DESCRIPTION
NSN
33
Rivet, Blind, Flush Head,
Monel Sleeve and Inconel Nickle Spindle
NAS1739MW5
(80205)
34
Rivet, Blind, Protruding Head
NAS1738B-4
(80205)
35
Rivet, Blind, Protruding Head
NAS1738B-5
(80205)
36
Rivet, Blind, Protruding Head
NAS1738B-6
(80205)
37
Rivet, Blind, Protruding Head,
Locked Spindle
NAS1398-6
(80205)
38
Rivet, Blind, Structural Pull,
Stem, Protruding Head
MS20600BK-1
(80205)
39
Rivet, Blind, Structural Pull
Stem, Protruding Head
MS20600-B4-W1
(80205)
40
Rivet, Blind, Structural Pull
Stem, Protruding Head
MS20600M6
(80205)
41
Rivet, Blind, Universal Head
CR2249-3
(11815)
42
Rivet, Blind, Universal Head
CR2249-4-1
(11815)
5320-00-866-6114
43
Rivet, Blind, Universal Head
CR2249-4-5
(11815)
5320-00-349-5132
44
Rivet, Blind, Universal Head
CR2249-6-3
(11815)
5320-00-779-0300
45
Rivet, Hi-Loc
HL2086W-5
(73197)
6
Rivet, Hi-Loc
HL2086W-6
(73197)
47
Rivet, Solid, Aluminum Alloy,
Flat Head
MS20426AD3
(80205)
48
Rivet, Solid, Aluminum Alloy,
Flat Head
MS20426AD4
(80205)
Change 38
2-177
5320-00-582-3273
5320-00-117-6948
TM 55-1520-234-23
Table 2-2. Structural repair materials (Cont)
ITEM
NO.
REF. NO.
AND FSCM
DESCRIPTION
49
Rivet, Solid, Aluminum Alloy,
Flat Head
MS20426AD5
(80205)
50
Rivet, Solid, Aluminum Alloy,
Flat
MS20426DD4
80205)
51
Rivet, Solid, Aluminum Alloy,
Universal Head
MS20470AD3
(80205)
52
Rivet, Solid, Aluminum Alloy,
Universal Head
MS20470-AD4
(80205)
53
Rivet, Solid, Aluminum Alloy,
Universal Head
MS20470-AD5
(80205)
54
Rivet, Solid, Aluminum Alloy,
Universal Head
MS20470-DD6
(80205)
55
Rivet, Solid, Universal Head
MS20615-3M3
(80205)
56
Rivet, Solid, Universal Head
MS20615-3M4
(80205)
57
Rubber, Type II, Grade A
Soft, 0.125 x 0.190
MIL-R-6130
58
Steel Sheet, Stainless,
0.016 inch thick
MIL-S-5059A
59
Steel Sheet, 0.032 inch thick,
N-155
9515-00-632-2982
60
Steel Sheet, 0.063 inch thick,
4130 COND-N
MIL-S-18729
61
Titanium
M/L-T-9046
Type 1, Comp. C
62
Rivnut P/N 2R1393
2-178
Change 38
NSN
TM 55-1520-234-23
CHAPTER 3
ALIGHTING GEAR
3-1. Landing Gear.
The landing gear consists of two skid tubes (2,
figure 3-1) and two arched cross tubes (8) of formed
aluminum alloy. The assembly is attached to the lower
fuselage structure with clamps at four points. The part of
the cross tubes extending from the fuselage are
enclosed with thermoplastic fairings (5), which are
streamlined to reduce aerodynamic drag. (It is
permissible to fly this aircraft with both cross tube
fairings removed. No additional flight restrictions will be
required when both fairings are removed.) The lower
fuselage openings, to accept the cross tubes, are
covered with aluminum alloy fairings. Eyebolts (13) are
provided on the skid tubes (2) to accommodate ground
handling wheels. To prevent abrasion and damage from
contact with the ground, replaceable steel skid shoes (3)
cover the bottom side of the skid tubes.
(3) To remove skid tubes from cross
tubes, remove screws from saddles (4) and detach skid
tubes.
b.
Inspection.
(1) Covers (10, figure 3-1) for cracks,
holes and corrosion.
Standard structural repair
procedures are acceptable if repairs are made.
(1)A Inspect support assemblies (11 and
12) as follows:
(a) Presence of deteriorated or
worn rubber bumpers is not cause to replace the support
assembly. Missing rubber bumpers is cause to change
the support assembly:
Pre-Maintenance Requirements for Landing Gear
Conditions
Model
Part No. or Serial No.
Special Tools
Test Equipment
Support Equipment
Minimum Personnel
Required
Consumable Materials
Special Environmental
Conditions
(b) Inspect metal portion of the
support assemblies for cracks, distortion, and elongated
holes. Obvious damage is cause to replace the support
assembly.
Requirements
(2) Landing gear skid tubes (2), skid
tubes in area between cross tube saddles, and cross
tubes (8) for scratches, scuffs, nicks, dents and holes.
AH-1S
All
None
None
Jacks
Three
NOTE
Smooth dents, not exceeding 0.25 inch in
depth and 1.0 TO 1.2 inches in diameter,
between the cross tube saddles may be
disregarded.
(C47) (C87) (C102)
(C107) (C118)
None
(3) Skid shoes (3) for damage, wear, and
loose or missing fasteners.
a.
Removal. Complete landing gear can be
removed as an assembly, or skid tubes (2, figure 3-1)
and cross tubes (8) can be removed separately.
(4) Cross tube retainers (9) for looseness
and for cracks and the cross tubes for cracks per (5)
below.
(1) To remove complete landing gear: (a)
Jack or hoist helicopter off ground. If using jacks, align
legs to allow clearance for removing landing gear.
(Refer to Chapter 1.)
(5)
If it is suspected that the cross tubes
are cracked due to a hard landing or other cause,
prepare the cross tubes for inspection and perform
same as outlined below. (Figure 3-3 specifies the
minimum areas that must be evaluated.)
(b) Remove step (6), fairing (5),
and seal (7). Remove covers (10) under cross tubes.
NOTE
(c)
Remove bolts and four supports
(11) and (12) and lower landing gear to ground. Turn
assembly to clear jacks and slide from under fuselage.
(2) To remove fairing from cross tubes
(8), remove screws and detach from cross tubes.
Change 71
The cross tubes may be inspected by flourescent dye
penetrant inspection (TM 43-0103, Chapter 6),
ultrasonic inspection (TB 55-1520-243-50-2), or
radiographic inspection (TM 43-103, Chapter 5).
Replace cross tube or cross tube saddle if any cracks
are detected
.
3-1
TM 55-1520-234-23
Figure 3-1. Alighting Gear and Support Installation (Sheet 1 of 3)
3-2
Change 29
TM 55-1520-234-23
Figure 3-1. Alighting Gear and Support Installation (Sheet 2 of 3)
Change 2
3-2A
TM 55-1520-234-23
Figure 3-1. Alighting Gear and Support Installation (Sheet 3 of 3)
3-2B
Change 2
TM 55-1520-234-23
(a)
Prepare areas
inspection.
Remove retainers (9, Figure 3-1).
in Figure 3-3 for dye penetrant
NOTE
The success and reliability of penetrant
inspection depends upon the thoroughness
with which inspector prepares the part from the
pre-cleaning process all the way through to the
final interpretation of the indications.
All
inspections should be with the fluorescent
penetrant (Type I, Method C) in strict
accordance with TM43-0103.
MEK (Metyl- Ethyl-Ketone (C74)) or paint remover
(TTR) 248B) and remove the paint.
(c)
Clean the prepared surfaces
with a soft cloth.
(d) Apply
a
fluorescent
dye
penetrant to the prepared surfaces from either a spray
can or with a soft hair brush and in strict conformance to
the procedures specified in TM 43-0103, Chapter 6.
(e) Allow penetrant to dwell for a
minimum of 30 minutes.
(f)
Clean off all excess penetrant in
accordance with TM 43-0103, standard procedures.
(Check for complete excess penetrant removal from
surface by using a blacklight.)
Prolonged or repeated inhalation of vapors or
powders may result in irritation of the mucous
membrane areas of the body.
Provide
adequate ventilation.
Continual exposure to penerrant inspection
materials may cause skin irritation. Avoid
prolonged breathing of solvent vapors and
contact with skin or eyes.
(g) Apply applicable developer
consistent with Type I, Method C penetrant method in
TM 43-0103.
(h) Inspect suspected area with
blacklight source in subdued white light.
NOTE
Normal manufacturing machining marks may
be observed on the tube surfaces. These
will not be cause of part rejection.
(i)
Clean tube with solvent and
wipe dry. U) Recoat cross tube surfaces to be covered
by retainer plates with sealant (C116). Install retainer
plates with rivets coated with primer (C102).
Injury to eyes and skin may occur when
blacklight is not used in accordance with
manufacturer's inspections.
Unfiltered light
sources (if filter is required) may possibly
damage the eyes.
(k)
Repaint cross tubes with primer
(C102).
(6)
Fairings for cracks and security.
(7) After hard landing or overloading,
check landing gear to determine if cross tubes (8, figure
3-1) have taken a permanent set at excessive spread.
Temperatures in excess of 120 degree F may
cause bursting of pressurized cans and injury to
personnel.
(a)
Position the helicopter on a
smooth surface.
(b)
Raise the hehcopter off the
surface with jacks removing all weight from the landing
gear.
Volatile fumes may occur, creating both a fire
and health hazard.
NOTE
Paint will not be removed by any mechanical
means under any circumstances because it
may mask over any potential surface cracks.
(b)
With a soft hair brush, apply
Change 23
(c)
Level the helicopter. (Refer to
Chapter 1.)
(d)
Measure the distance between
the cross tube retainers and divide that distance to
determine the helicopters center line. (See figure 3-2.)
3-2C
TM 55-1520-234-23
NOTE
Distance should be 38 to 40 inches from inside edge of
skid tube to plumb line. If distance exceeds 40 inches
from the inside edge of either skid tube, replace cross
tubes.
(e) Drop a plumb line from helicopter center line to
ground, or floor surface. Measure from plumb line to
the inside of each skid tube at cross tube locations.
(f) Lower helicopter to surface and remove jacks.
Figure 3-2. Checking deflection of cross tube
3-2D
Change 71
TM 55-1520-234-23
c.
Repair or Replacement.
(1)
(b) Dents over 0.25 inch deep and
1.0 to 1.2 inches in diameter between the cross tube
saddles require repair as shown in paragraph 3-2.
Cross tubes:
(a) Minor scratches, scuffs, and
nicks may be polished out to depth of damage, but not
to exceed depth limits as indicated in figure 3-3.
(c)
Holes in skid tubes require
repair as shown in paragraph 3-2.
(b) All other
replacement of cross tubes.
requires
(d) Replace skid tubes which show
excessive wear or damage.
(c)
Replace
cross
tubes
if
deflection dimension exceeds inspection requirements.
(Refer to paragraph 3-1, step b.)
(3) Replace skid shoes if damaged or if
they no longer protect skid tube. Replace missing
fasteners.
damage
(1)A Cross tube support assembly.
will be removed.
(C87) and dry.
(a) All traces of fuel, oil, and grease
Clean support assembly with MEK
(b) Apply
adhesive
EC-2126
(C21A) to support assembly and allow adhesive to dry
tack free. Align rubber bumper on support assembly
and hold firm contact pressure. Place support assembly
in oven and heat to 250°to 260°F.
(c)
Allow support assembly to cool
before installation on crosstubes.
(2)
NOTE
Improper length screws and method of removing and
installing skid tube assemblies may cause rivnuts to
come loose.
(4) Ensure proper alignment of saddle
and skid tube to prevent damage to retaining screw
threads.
(5) Use C clamps on forward and aft ends
of saddle (each side of cross tube), clamping skid tube
on lower side and tighten to prevent movement of parts.
(6) Use new retaining screws on each
side of the saddle when replacing skid tube.
Skid tubes:
(a) Scratches up to 0.03 inch deep
and 1.0 to 1.2 inches long, running directly across top of
tube between cross tube saddles may be polished out.
Scratches beyond these limits require repair as shown in
paragraph 3-2.
(7) Apply sealing compound (C118) to
rivnut and skid tube surfaces for installation.
(8)
Figure 3-3. Damage Limits - Crosstubes
Change 71
3-3
Repair damaged fairings as follows:
TM 55-1520-234-23
location.
WARNING
Provide adequate ventilation when using
methyl-ethyl-ketone. Avoid breathing solvent
vapors and avoid prolonged skin contact.
(a)
Clean area around the crack
perimeter, using a clean cloth dampened with methylethyl-ketone (C87). Lightly sand area using fine grade
of sandpaper (C112). Clean the surface as described
above with methyl-ethyl-ketone after sanding.
(b) Two layers of fiberglass are required. Measure and cut one fiberglass patch from
fiberglass cloth (C47) to cover cracks by a minimum of
one inch around the perimeter. Cut a second patch to
overlap the first patch by one inch on all sides.
(c)
Using a brush, apply a coat of
epoxy (C107) to the cleaned surface of fairing to match
size of patch. Apply fiberglass patch and brush epoxy to
cover patch completely, using brush to work out air
bubbles.
(d) Apply another coat of epoxy
(C107) to cover area for second patch. Apply fiberglass
patch and brush epoxy to completely cover and saturate
fiberglass patch. Work out air bubbles.
(e) Allow epoxy (C107) to cure for
four hours at room temperature prior to flying helicopter.
The complete cure time for epoxy is 24 hours at room
temperature.
Refinish repaired area to match
surrounding surface.
(f)
There is no limit to the number,
size, or length of cracks allowed to be repaired.
d.
Installation.
CAUTION
It is possible to erroneously interchange the fore
and aft crosstubes during assembly to the skid
tubes.
Consult
TM 55-1520-234-23P to
ensure the tubes are installed in the proper
3-4
(1) If separated, assemble skid
tubes (2, figure 3-1) to cross tubes (8). Align holes in
cross tube saddles (4) with holes in skid tubes and install
screws.
(2) Lift and support landing gear in
position in fuselage fittings.
NOTE
Prior to torque application, the gap between
the top of the support assembly and the
bottom of the fuselage at the crosstube
mounting area should be in accordance with
figure 3-3A. The gap shall be measured with
appropriate feeler gage. If the gap is larger
than recommended, fabricate shims per
figure 3-3A If the gap is smaller than
recommended, the rubber pad is worn and
the cap should be replaced. Tighten bolts to
a snug fit, while aircraft is still supported by
jacks or hoists. Depending on thickness of
shim required, longer bolts of the same part
number may be needed to install caps.
(3) Attach each forward support
with one bolt, installed up through support into upper
fitting. Tighten only enough to align four mounting holes
and install four bolts from outboard side, tighten 60 TO
80 inch-pounds torque. No further tightening of vertical
bolt is required.
NOTE
Check for minimum of 0.020 inch gap
between support bracket assembly at
vertical bolt location and support of aircraft.
This is to insure there is no preloaded shear
stress on lateral support bolts.
(4) Install aft support (12) with four
bolts and washers up through support and install four
nuts with washers.
Tighten bolts to a snug fit.
Change 71
TM 55-1520-234-23
Figure 3-3A. Crosstube Shim Fabrication
Change 71
3-4A/(3-4B blank)
TM 55-1520-234-23
(5) Lower helicopter to ground and torque
aft support bolts 70 TO 80 inch-pounds.
(6) Install seal (7), fairing (5), step (6),
and covers (10()).
3-2. Skid Tube Repair.
(2) Cut out damaged portion of skid tube
with hand or powered metal saw.
(3) Fabricate an insert of the required
length from tubing 0.083 inch wall thickness(item 27,
table 2-1) or from scrap skid tube, as shown in figure 34.
(4)
Fabricate splice plates as follows:
a.
Damage Reparable by Patching. Repair
scratches running directly across top of tube more than
0.03 inch deep and 1.2 inches to a maximum of 4.0
inches long, dents more than 0.25 inch deep and 1.2
inches to a maximum of 4.0 inches across and any
holes up to 4.0 inches in diameter through one surface
of tube only as follows:
(a) Cut four plates to the required
dimensions as shown in figure 3-4 from sheet stock
0.100 inch thick (item 12, table 2-1),or use material
salvaged from scrap skid tube.
(1) Lift helicopter and remove skid tube in
accordance with instructions contained in paragraph 3-1.
a.
(b)
Form two plates to fit the
outside diameter of skid tube and the other two plates to
fit the inside diameter as shown in figure 34.
(2) Polish out scratches, trim and smooth
rough edges of holes.
(5) Apply a coat of primor (C102) to
plates and tubes.
(3) Fabricate a patch from 0.100 inch
thick aluminum alloy (item 12, table 2-1) of the required
size as shown in figure 34 or make a patch from
material salvaged from scrap skid tube.
(6) Lay out rivet hole pattern on upper
splice plates and lower side of tubes as shown in figure
3- 4.
(4) Lay out the rivet hole pattern and
form patch to fit contour of skid tube as shown in figure
34.
(5) Locate and securely clamp patch to
skid tube and drill rivet holes with a No. 10 drill.
(6) Rivet patch in place using blind rivets
(item 32, table 2-1).
(7) Apply primer and lacquer in
accordance with painting instructions in TB 746-93-2.
(8) Install skid tube in accordance with
paragraph 3-1. d. and lower helicopter.
b.
Damage Reparable by Insertion. Repair
dents and holes on either top or bottom side of skid tube
which is greater than FOUR inches across in any
direction by inserting a splice of new tubing. Such
repairs are restricted to the areas shown in figure 3-4.
(7) Maintain proper alignment
securely clamp splice plates and tubes together.
(8) Drill rivet holes in plates and tubes
with No. 10 drill. Countersink lower holes with 100
degree countersink. Install blind rivets (item 31, table 21) in upper half of splice and flush rivets (item 31, table
2-1) in lower half of splice as shown in figure 3-4.
(9) If repair involves removal of skid shoe
bolt sleeves, mark sleeve locations using skid shoe as a
template and install new sleeves.
(10) Apply a touch-up coat of primer
followed by lacquer in accordance with TB 746-93- 2.
(11) Install skid tube in accordance with
paragraph 3-1. d., and lower helicopter.
c.
Damage Necessitating Replacement.
Damage to the skid tubes beyond the limits given for
repairs by patching or insertion as shown in figure 3-4
requires
replacement
of
the
skid
tube.
(1) Jack helicopter and remove skid tube
in accordance with instructions in paragraph 3-1.a.
Change 65
and
3-5
TM 55-1520-234-23
Figure 3-4. Skid tube repairs (Sheet 1 of 2)
3-6
Change 65
TM 55-1520-234-23
Figure 3-4. Skid tube repairs (Sheet 2 of 2)
Pre-Maintenance Requirements For
Ground Handling Wheels
3-3. Ground Handling Wheel Actuating
Mechanbm.
Two ground handling wheel assemblies are provided for
quick mounting on landing skids to facilitate moving
aircraft on ground. Each assembly consists of two
wheels on an offset axle, a supporting cradle, and a
hand-operated hydraulic jack with two rams which
actuate axle to extend or retract wheels. (See figure 35.) The cradle is mounted on eyebolts on landing skid by
means of a fixed rear pin and a spring-loaded front l pin.
Change 2
Conditions
Requirements
Model
Part No. or Serial No.
Special Tools
Test Equipment
Support Equipment
Minimum Personnel
Required
Consumable Materials
Special Environmental
Conditions
AH-1S
All
T15) (T16)
None
None
One
3-7
(C45) (C74) (C124)
Dust Free
TM 55-1520-234-23
Figure 3-5. Ground handling wheels
3-8
TM 55-1520-234-23
a. Disassembly.
(1) Remove ball-lock pin (5, figure 3-5)
and remove support rod (6) from axle (4).
(h) Remove overload valve body
(26) from tie rod (28). Remove spring (27) and plunger
(29) from body (26).
(6) Remove cotter pin, washer and
lubrication pin (7, figure 3-5) attaching ram arm (23) to
clevis (8) of hydraulic ram (10).
Deflate tire prior to wheel assembly removal.
(2)
Remove wheel (3) with tire and tube
assembled.
(3) Disconnect and remove flexible
hose(21) from tee on hydraulic pump (15) and hydraulic
ram (10).
(4) Remove nuts and washers and lift Ubolt (22) attaching hydraulic pump (15) to cradle
assembly (12). Remove hydraulic pump (15).
(5)
Disassemble
hydraulic
pump
(8) Remove lubrication fittings (16),
unscrew and remove connecting pin (17) and release
pin (18).
NOTE
When connecting pin (17) is removed, support
pin (19) can be released and spring (20) will
slide from cradle.
as
follows:
(9)
Remove trunnion (13) from cradle
(12).
(a) Remove retaining rings (1,
figure 3-6), fulcrum pins (2) and separate handle
assembly (3) from pump body (4).
(b) Remove tank filler hole screw
(6) and drain oil from tank.
(c)
Pull out piston (7) and remove
clip (8) by spreading clip slightly. Unscrew gland nut (9)
using adjustable spanner wrench and remove packing
support (10), leather packing (11), rubber packing (12)
and spreader (13).
(d)
Pry out filter screen (14) from
hose hole. Remove retaining screw(15),discharge valve
spring (16), 5/16 inch diameter ball (17), suction valve
spring (18) and 3/16 inch diameter ball(19).
(e) Remove screw (32).
Grasp
knob (33) and detach from valve stem (34). Unhook
loop of spring (35) from pin on pump body (4). Slip
knob (33) onto valve stem (34). Remove spring (37),
steel washer (38), packing (39) and 5/16 inch diameter
ball (40).
(f)
Remove nut (20) and packing
(21). Twist tank (22) off pump body (4). Remove seal
(23).
(g)
(25). Discard screen.
(7) Back out set screw (14) and remove
hydraulic ram (10)from trunnion (13). Using clevis (8)
as handle, hold ram housing or cylinder. Separate ram
piston from cylinder.
b.
(1) All foreign particles from magnet
assembly (30, figure 3-6) with clean cheese cloth (C36).
Cleaning solvent is flammable and toxic.
Provide
adequate
ventilation.
Avoid
prolonged breathing of solvent vapors and
contact with skin or eyes.
(2) Recessed hole, into
screen (25) fits, with solvent (C124).
which
filter
(3) Inside of overload valve body (26)
with solvent (C 1 24).
(4) Rod (28) with solvent (C124) to
assure clear passage through hole in end of rod.
c.
Inspection.
(1) Ball-lock pins (5, figure 3-5) for
cracks, corrosion, wear and distortion.
Remove screw (24) and screen
Change 17
Cleaning.
3-9
TM 55-1520-234-23
(2) Wheels (3) for cracks, distortion,
corrosion and damage.
tires, tubes (3, figure 3-).
(3)
(3)
Tires and tubes for cracks, wear and
Replace lubricator pin (7) if worn or
distorted.
abrasion.
(4) Lubricator
distortion and wear.
pin
(7)
for
damage,
(4) Replace
threads are damaged.
(5)
(5) Internal threads of trunnion (13) for
damage and set screw (14) and its internal threads in
trunnion for damage.
(6)
Replace
trunnion
(13)
lubricator
if
internal
fitting (16) if
damaged.
(6) Replace connecting pin (17), support
pin (19), and spring (20) if distorted or damaged.
Lubrication fitting for serviceability.
(7) Connecting pin (19) and spring (20)
for damage or distortion.
(8) Flexible hose (21) for leaks, damage,
and serviceability.
(9) Axle (4), cradle (12) and sleeve for
wear, corrosion damage and cracks.
(7)
or damaged.
Replace flexible hose (21) if leaking
(8) Replace axle sleeve or cradle if
cracked or damaged.
(9) Repair hydraulic pump P/N BU953B if
leaking. (See figure 3-6.)
(10) Hydraulic ram assembly (10) for
leaks, security, corrosion and damage.
(a) If the hydraulic pump is to be
repaired, procure a hydraulic pump parts kit P/N JS953.
Refer to TM 55-1520-234-23P.
(11) Hydraulic pump
security, corrosion and damage.
(b) Replace clip (8, figure 3-6)
leather packings (11), and rubber packing (12).
(15)
for
leaks,
(12) Packings, seals, and gaskets
hydraulic pump for distortion, wear or damage.
of
(13) Washers, screws, retaining rings, clips
and springs of hydraulic pump for damage and
serviceability.
(14) Screens of
damage and serviceability.
hydraulic
pump
(c)
Replace filter screen (14),
discharge valve spring (16), 0.3125 inch ball (17),
suction valve spring (18), and 0.1875 inch ball (19).
(d) Replace release valve spring
(37), washer (38), packing (39), and 0.3125 inch ball
(40).
for
(e)
Replace packing (21) and seal
(f)
Replace screw (24) and screen
(23).
(15) Balls of hydraulic pump for corrosion
and mechanical damage.
(25).
(16) Hole in rod (28, figure 3-) for clear
passage.
(10) Repair hydraulic ram (10, figure
3-5) as follows:
d.
Repair or Replacement.
(1) Replace
ball-lock
unserviceable.
pins
if
Deflate tire prior to wheel assembly removal.
(2)
NOTE
If hydraulic ram does not have piston P/N
330617, which is machined for packing and
backup ring, requisition new piston P/N 330617.
(a) Carefully slip new backup ring
over inboard end of piston P/N 330617 and into packing
groove.
Replace cracked or damaged wheels,
3-10
Change 17
(3)
(b) Carefully slip new packing over
inboard end of piston and into packing groove.
NOTE
Ensure each packing is not spiraled in groove.
(4) Replace outlet check spring (32),
0.3125 inch diameter ball (30) and 0.2187 inch diameter
ball (31).
(5)
(c)
Burnish
scratches
inside
hydraulic ram cylinder that are less than 0.005 inch
deep using crocus cloth (C45).
(11) Replace hydraulic ram if inside of
cylinder has nicks, scratches, or pits deeper than 0.006
inch.
d.1. Repair hydraulic pump, P/N HP9902-4110, if leaking. (See figure 3-6A.)
(1) Procure a hydraulic pump service kit,
P/N KH9000. Refer to TM 55-1520-234-23P.
(2) Replace packing (6), cup retainer (7),
pump cup (8), and spreader (9).
Change 17
TM 55-1520-234-23
Replace release packing (18).
e.
Replace three gaskets (23).
Assembly.
(1) Insert trunnion (13, figure 3-5) in
cradle (12) with threaded openings aft.
NOTE
Steps (2) through (24) are for assembly of
hydraulic pump, P/N BU953B. Step (24A),
substeps (a) through (t) are for assembly of
hydraulic pump P/N HP9902-41-10.
(2) Insert brass spreader (13, figure 3-6)
in body (4), flat side down.
3-10A/(3-10B blank)
TM 55-1520-234-23
Figure 3-6. Ground handling gear pump assembly (Sheet 1 of 2)
3-11
TM 55-1520-234-23
Figure 3-6. Ground handling gear pump assembly (Sheet 2 of 2)
(3)
Slide packing support (10) onto piston
(8) Drive piston and packing down solid,
using medium weight hammer on seating tool.
(7).
NOTE
(9) Remove packing seating tool and
assembly bushing tool.
The "V" must face away from groove on piston.
(4) Dip two leather packings (11), one
rubber packing (12), and third leather packing in
hydraulic fluid (C74) and assemble in that order over
bottom end of piston.
(10) Install and tighten gland nut (9) using
an adjustable spanner wrench.
NOTE
The '"V" on packing must rest on brass spreader.
(12) Insert 3/16 inch diameter ball (19),
suction valve spring (18), 5/16 inch diameter ball (17)
and discharge valve spring (16). Install screw (15).
(11) Replace clip (8).
(5) Place assembly bushing tool (T15) into top
of hole in pump body (4).
(13) Install filter screen (14) in hose hole.
(14) Insert 5/16 inch diameter ball (40,
rubber packing (39), steel washer (38) and spring (37)
into pump body (4).
(6) Insert piston (7) with packings installed into
bellmouth of assembly bushing toll (T15).
(7)
Slip packing seating tool (T16) over piston.
3-12
TM 55-1520-234-23
(15) Install valve stem (34) in base and turn
down against ball by slipping knob (21) into stem and
tightening.
top and in line with pump handle. Tighten nut (20).
(23) Replace tank filler hole screw (6).
(24) Position handle assembly (3) to body
(4) and secure with fulcrum pins (2) and retaining rings
(1).
Do not force down tight.
(16)
(24A)
Assemble
P/N HP9902-41-10, as follows:
Remove knob (33) from stem (34).
hydraulic
pump,
NOTE
(17) Position spring (35) over release knob (33)
and hook one eye of spring onto pin (36). Place knob
and spring over valve stem (34) and hook eye to spring
onto pin (36) in pump body (4).
Procedure outlined in this step is for assembly of
hydraulic pump, P/N HP9902-41-10. For assembly of
hydraulic pump, P/N BU953B, refer to steps (1) through
(24) above.
NOTE
(a) Insert spreader (9, figure 3-6A)
into pump base (10) flat side down.
Do not place knob onto hex of valve stem.
(b)
NOTE
Slide
packing
nut
(5)
onto
plunger (4).
Knob and valve stem should work free and valve should
close firmly when opened and released.
(c) Dip packing (6), cup retainer (7),
and pump cap (8) into hydraulic fluid (C74), and
assemble in that order over bottom end of plunger (4).
CAUTION
NOTE
An improperly operating valve
assembly and an inoperative pump.
means
improper
(18) Hold pump body (4) and valve stem
(34) firmly and twist knob (33) to the left two faces of the
hex. Push knob onto the hex of the valve stem at this
position. Insert flat head socket screw (32) and tighten.
Try knob action to see if closing is positive. If action is
not positive, move knob to the left another face on hex,
and recheck closing.
The "V" on pump cup must rest on spreader.
(d) P lace assembly bushing tool (T15) into
top hole in pump base (10).
(e) Insert plunger (4) with packing nut,
packing, cup retainer, and pump cup installed into
bellmouth of assembly bushing tool (T15).
(f) Slip packing seating tool (T16) over
plunger.
(19) Install filter screen (25) in pump body
(4) and secure with screw (24).
(20) Install plunger (29) and spring (27) in
overload valve body (26). Screw tie rod (28) into valve
body (26) and position in tank (22).
(21) Install gasket (31) and seal (23) and
assemble tank (22) to body (4).
(g) Drive plunger and assembled parts
down solid, using medium weight hammer on packing
seating tool.
(h)
Remove packing seating
tool and assembly bushing tool.
(i) Tighten packing nut (5) using packing nut tool. (Refer
to figure 3-6B.)
(22) Install packing (21) and nut (20) and
tighten nut lightly. Rotate tank so that filler hole is on
Change 7
3-13
TM 55-1520-234-23
(j)
Insert 0.3125 inch diameter ball
(31, figure 3-6A), 0.2187 inch diameter ball (30), and
outlet check spring (32) into base(10). Install valve plug
(33) into base (10).
(k)
Install screen (21) in pump
1
hydraulic test stand.
Connect
hydraulic
pump
to
2
Set relief valve to open at 8500
(plus or minus 300) psi.
base (10).
(I)
Install release packing nut (19)
on release spindle (12). Check threads on release
packing nut (19) and release spindle (12) for free
turning. Remove release packing nut (19) from release
spindle (12).
NOTE
3
Release pressure.
Then
increase pressure to 8000 psi. Observe for 15 seconds.
Loss of pressure in excess of 500 psi is cause for
rejection.
4
Open release valve to drop
pressure. From a 10 degree open position on handle
(13), release quickly, letting return spring (20) close
release valve.
Touch up threads if damaged or binding.
(m)
Dip release packing (18),
release packing nut (19) and release spindle (12) into
hydraulic fluid (C74) and assemble parts in that order.
Install release spindle (12) with release packing (18) and
release packing nut (19) in pump base (10). Tighten
release packing nut (19) until release packing (18)
bottoms out in pump base (10). Loosen release packing
nut (19) and torque to 20 inch-pounds.
(n) Position return spring (20) over
release spindle (12). Use capscrew (17) and washer
(16) to attach return spring (20) to pump base (10).
(o) Hold pump base (10) and
release spindle (12) firmly. Place handle (13) on
release spindle. Handle shall be vertical with release
valve in closed position. Install screw (15) and lock
washer (14). Hook return spring (20) around handle
(13). Try handle action to check for positive closing of
release valve.
NOTE
Do not push on handle.
5
Increase pressure to 8000 psi.
Observe for 15 seconds. Loss of pressure in excess of
500 psi is cause for rejection
(r)
Install shim washer (27) and/or
reservoir shim washer (26) as required to line up filler
plug (25) on reservoir (24) with top of pump within plus
or minimum 10 degrees. Use of shim washers (27) and
reservoir shim washers (26) varies from 0 to a total of 4.
Install reservoir (24) in pump base (10). Ensure that
filler plug on reservoir (24) lines up with top of pump
within plus or minus 10 degrees.
(s)
Position beam (2) to plunger (4)
and pump base (10). Install plunger cross pin (3) and
beam pin (28). Secure beam pin (28) and cross pins (3)
with two cotter pins (29).
(t)
Install reducer bushing (11).
NOTE
If handle action is not positive, move handle to the left
another face on hex shank of release spindle, and
check again for positive closing.
(p)
Install relief valve (22) and
gasket (23) in pump base (10).
(q)
(u) Connect air line to fill port in
reservoir (24). Apply 100 psi air pressure, and check
hydraulic pump under water for signs of leakage. Reject
hydraulic pump if there are signs of leakage. Remove
hydraulic pump from water. Remove air line from
reservoir (24). Install gasket (23) and filler plug (25).
Set relief valve as follows:
3-14
(v)
Change 7
Install handle (1) to beam (2).
TM 55-1620-234-23
Figure 3-6A. Ground handling gear pump, P/N HP-9902-41-10 (Sheet 1 of 2)
Change 7 3-14A
TM 55-1520-234-23
Figure 3-6B. Packing nut tool
NOTE
To prepare a new hydraulic pump (15, figure 35) and hydraulic ram (10) assembly for
installation, remove pipe plug on each and drain
original fluid.
on sleeve, insert axle (4), and secure with bolts. Insert
sleeve through cradle (12) and install hydraulic ram arm
(23) and axle on opposite end. Hydraulic ram must be
forward' of wheel hub center line 1.98 inches. (Refer to
figure 3-7.)
(25) Install hydraulic ram (10, figure 3-5) on
each end of trunnion (13) to bottom out in hole. Back
off until hydraulic outlet is directed down. Secure with
set screw (14).
(27) Position hydraulic pump (15, figure 3-5) on
cradle (12) and secure with U-bolts (22).
(26) Position hydraulic ram arm (23, figure 3-4)
3-14B
(28) Install ram clevis (8, figure 3-5) and with
hydraulic ram fully extended, adjust clevis to hold 1.48
inches diameter. Refer to figure 3-7.
Change 7
TM 55-1520-234-23
(29)
Insert support pin in aft end of cradle, align holes and secure with spring pin.
(30)
Insert release pin (18, figure 3-5) in upper forward orifice of cradle, align holes in both pins and install
connecting pin (17).
(31)
Attach support rod (6) to clevis pin and insert ball-lock pin (5).
(32) Install tire, tube and wheel.
f.
Bleeding - Hydraulic Pump.
(1) Fill hydraulic cylinder (15, figure 3-5) with hydraulic fluid (C73).
(2) Operate pump handle for several strokes to build up pressure.
Figure 3-7. Ground handling wheels adjustment dimensions
Change 7
3-15
TM 55-1520-234-23
(3) Crack (loosen) hose coupling (21) at tee on
pump (15).
is too low, turn the rod clockwise. Test and readjust as
necessary until proper setting is obtained.
(4) Operate pump handle until air bubbles no
longer show at loose hose coupling (21) and fluid runs
smoothly.
(4) When proper setting is obtained, tighten hex
nut (20).
(5) Refill hydraulic pump and repeat previous
steps to be sure all air is expelled from system.
NOTE
Hold tie rod in position using screw driver in slot to
prevent rod turning with the nut.
(6) Tighten hose coupling (21) to pump (15)
and refill cylinder.
h. Testing - Hydraulic Ram.
(1) Screw rain assembly into trunnion (13, figure
3,5) and connect ram to pump (15).
g. Testing - Hydraulic Pump (A VIM).
(1) Fill the oil tank to proper level with hydraulic
fluid (C73).
(2) Connect a 10,000 psi pressure gage to
outlet hole.
(3) Operate pump until pressure builds up and
overload valve unloads. Proper setting is 8300 to 8800
psi. If pressure goes too high, turn the tie rod (28, figure
36) counterclockwise, using a screw driver. If pressure
3-16
(2) Pump until overload in pump goes off with ram
against trunnion stop.
(3) Check for leaks.
(4) Release pressure and pump ram out halfway.
Allow to stand a few minutes.
(5) Check for leaks.
when no leaks are found.
Change 7
Ram is ready for service
TM 55-1520-234-23
CHAPTER 4 - POWER PLANT
Section I. POWER PLANT
4-1.
Power Plant.
The power plant consists of a T53-L-703 series shaft
turbine engine mounted horizontally on the fuselage
behind the main rotor pylon, with adapting parts and
connections to the airframe structure and to fuel, oil,
electrical, instrument, and engine control systems. (See
figures 4-1 and 4- 2.) The engine and transmission are
enclosed by cowling and fairing. Hinged pylon fairing
doors at each side give access to the air induction and
drive shaft area ahead of the engine forward firewall.
These doors also have engine air inlet shield. The
engine compartment between forward and rear firewalls
has hinged side doors equipped with cooling air inlets.
Doors also have armor panels to protect the fuel control
and compressor section. The exhaust area, behind the
rear firewall, is enclosed by removable fairing.
4-2.
Engine Maintenance.
Servicing and lubrication information will be found in
Chapter 1. Special inspections, and schedules for
overhaul and retirement of components are in Chapter
1.
Daily, intermediate, and periodic inspection
requirements are in Preventive Maintenance Checklists,
TM 55-1520- 234PMS. To accomplish maintenance
procedures refer to TM 55-2840-229-23.
4-3.
Engine Mounts.
The engine is suspended on two mount assemblies
at the diffuser housing and one leg assembly at left side
of the inlet housing. For maintenance on the engine
mounts refer to paragraph 2-44.
4-4.
Engine Vibration Tests. (AVIM)
a. Refer to TM 55-2840-229-23 for instructions to
perform engine vibration check.
To prevent a possible hazard to personnel and
damage to equipment, set meter power switch to
OFF, while breaking or making connections to a
source of electrical power.
CAUTION
Leave enough slack in cables to prevent
unnecessary strain on pickups and connectors.
Avoid conditions that would cause cables to
deteriorate from heat or abrasion.
(1) Route all three cables aft to one location and
secure to engine bleed air tube adjacent to engine ignition
exciter box with one MS21919H-15 and one MS21919H-10
cable clamp. See figure 4-3.
(2) Route cables out left engine compartment door
ventilation opening. Make sure cables are clear of engine
throttle control linkages.
Close and secure engine
compartment door.
(3) Route cable forward, over wing, along left side
of aircraft to gunners station. Remove screws in access
panel as shown in diagram. Retain screws for later
installation.
(4) Install one MS21919H-10 clamp at each
location shown. Replace screws with AN3 bolts and secure
clamp and cables. (5) Route cables inboard into gunners
compartment.
NOTE
During test it will be necessary to compress seal on
gunners access door under cables to allow closing
of the door for flight
(6) (Deleted)
b. Route pickup cables as follows:
(7) After test, remove all test equipment, cables
and clamp. Remove all AN3 bolts installed for test and
replace with original screws.
Change 29
4-1
TM 55-1520-234-23
Figure 4-1. Power plant installation
4-2
Change 7
TM 55-1520-234-23
4-5.
Engine Assembly.
(b) Disconnect plug of electrical harness
assembly (6) from aft firewall on left side of engine.
NOTE
When parts are removed, use care to avoid
entry of dirt or foreign material into engine or
components. Seal opening with caps, plugs or
temporary covers.
(c) Remove clamp securing fuel filter cable
assembly (13) and disconnect plug from fuel filter.
(d) Disconnect plug of cable assembly (15)
from engine oil bypass valve.
Premaintenance requirements for engine assembly.
Conditions
Requirements
Model
Part No. or Serial No.
Special Tools
Test Equipment
Support Equipment
Sling, Stand
Minimum Personnel Required
Consumable Materials
Special Environmental
Conditions
AH-1S
All
(T7)
None
Hoist, Lifting
(e) Remove clamp (28) securing harness
assembly (29) to right side aft engine mount tube.
(f) Disconnect plug of harness assembly (29)
from aft right side firewall.
(g) Disconnect exhaust thermocouple harness
plug (30) from aft right side firewall.
Three
(C97)
(h) Disconnect electrical plug (1) at bleed air
valve.
None
a. Removal of Engine Assembly (See figure 4-3A)
(1) Disconnect battery. Refer to paragraph 9-9.
(i) Disconnect oil tank
electrical plug (2) at airframe plug.
Refer to
(4) Remove particle separator.
paragraph 4-8.
Refer to
(5) Remove
paragraph 6-7.
Refer
level
switch
(8) Disconnect fuel, oil, and bleed air hoses. lines
and tubes as follows:
(2) Remove engine cowling and tailpipe fairing.
Refer to paragraph 2-16.
(3) Remove transmission cowling.
paragraph 2-15.
low
NOTE
When removing fuel and oil hoses, lines, fittings,
ensure residual fluids have been drained. Cap all
open ports until reinstallation. All parts removed
should be inspected for serviceability.
to
(a) If not previously accomplished, disconnect
hose assembly (25, figure 4-3A) at oil tank disconnect
fitting.
(6) Remove IR suppressor (if installed), ejector
and tailpipe. Refer to paragraph 4-12 and 4-13.
(b) If not previously accomplished, disconnect
hose assembly (26) at oil tank disconnect.
NOTE
When electrical cables and plugs are
disconnected from airframe plugs; cables, and
plugs should be secured to engine in a manner
that will avoid interference when removing
engine assembly.
(c) Disconnect fuel drain disconnect (27) from
right side engine deck fitting.
main
driveshaft.
(d) Disconnect oil bypass valve hose (l 6) at
engine deck fitting on left forward side of engine.
(7) Disconnect electrical cables and plugs on
airframe as shown in figure 4-3A as follows:
(a) Remove two clamps (5 and 10) securing
harness assembly (6) to left side of engine mount tubes.
(e) Disconnect fuel filter hose (14) at engine
deck disconnect fitting.
(f) Disconnect starter
disconnect (9) at engine deck.
Change 7
4-3
seal
drain
hose
TM 55-1520-234-23
Figure 4-3. Engine vibration test equipment cabling tiedown
4-4
TM 55-1520-234-23
Figure 4-3A. Engine assembly removal/installation (Sheet 1 of 2)
Change 22 4-4A
TM 55-1520-234-23
Figure 4-3A. Engine assembly removal/installation (Sheet 2of 2)
Change 22
4-4B
TM 55-1520-234-23
(g) Disconnect governor and filter bleed
hose disconnect (8) at engine deck.
(1) Remove output driveshaft adapter.
LTCT962 torque adjustment fixture (T7A).
paragraph 6-7.
(h) Disconnect bleed air line coupling (7) at
engine deck coupling.
(i) Disconnect tube assembly (4) from top
of engine. Disconnect opposite end at bleed air valve
and remove tube assembly (4).
(j) Disconnect bleed air tube coupling (3) at
top of engine. Disconnect opposite end and remove
bleed air tube.
(k) Remove cotter pin (20), nut (19)
washers (18), and bolt (17) securing droop compensator
control tube to cambox assembly. Disconnect upper
end of control tube from cambox.
Use
See
(2) Remove air inlet duct.
(3) Remove electrical cables.
(4) Remove starter generator and tachometer.
Remove oil pressure transmitter and pressure switch.
(5) Remove cambox assembly, actuator and all
connecting linkage.
(6) Remove fuel, oil, vent and drainhoses and
connections.
(7) Remove engine mount trunnions and any
support brackets and bellcranks normally attached.
CAUTION
To prevent possible damage to engine mount,
ensure weight is removed from forward engine
mount bolt before removal.
(8) Check engine and remove remaining fittings
and adapters.
Install caps, plugs, and covers over
openings.
c.
(I) Remove cotter pin (24), nut (23),
washer (22), and bolt (21) securing power lever control
tube to fuel control arm. Disconnect control tube from
arm.
CAUTION
If the engine was removed prior to normal overhaul
for internal failure, flush all lines, replace oil cooler
and engine oil filter. Tag removed oil cooler as
contaminated and forward to depot.
(9) Attach engine lifting sling (T7) to engine
assembly. Attach a suitable hoist to sling. Take slack
out of hoist and sling.
(10) Loosen bolts from left and right side pillow
blocks. Disengage pillow blocks. Remove bolt from
forward engine mount trunnion.
Installation of Engine Adapting Parts.
(1) Install engine mount trunnions; also install any
support brackets or bellcranks normally attached to same
bolts.
(2) Install oil system fittings and hoses.
(11) Slowly hoist engine from helicopter
structure ensuring all necessary hoses, tubing, and
electrical cables are disconnected and cleared of
airframe.
(12) Install engine on work stand or trailer as
required.
(3) Install cam
connecting linkage.
box
assembly,
actuator
and
(4) Lubricate splines with plastilube (C97) and
install starter-generator and tachometer generators. Install
oil pressure transmitter and pressure switch.
b. Removal of Engine Adapting Parts. To convert
a quick change engine assembly to a basic engine,
remove all parts not included on or with engines as
supplied in the shipping container. Remove all adapting
parts when preparing an engine for shipment or storage
in container. With engine on stand, remove parts as
follows:
Change 71
WARNING
Disconnect low tension lead prior to connecting
igniters. Tag lead with red tag and note on tag "Do
Not
Hook
Up
Until
Ground
Run."
4-4C
TM 55-1520-234-23
(5) Install necessary connections and hoses.
Install air inlet duct and electrical cables. Connect
engine igniters.
(6) Install output driveshaft adapter. Refer to
paragraph 6-7. Check over engine and remove any
remaining plugs or covers.
oil bypass valve.
(f) Attach clamps (5 and 10) to existing
clamps on engine mount tubes.
(g) Secure cable assembly (13) to fuel filter
using existing clamp on fuel filter.
(h) Connect electrical plug (1) to bleed air
d. Installation of Engine Assembly.
valve.
NOTE
Inspect
all
removed
components
for
serviceability prior to installing on new engine.
(1) Attach engine lifting sling (T7) to engine
assembly.
(i) Connect electrical plug (2) to airframe
mounted plug for oil tank low level switch.
(7) Connect engine fuel, oil and bleed air hoses
and tubing as follows:
(a) Connect fuel disconnect (27) to engine
(2) Attach lifting sling to suitable hoist. Take up
slack in hoist and remove hardware securing engine to
work stand or transportation trailer.
(3) Hoist engine to clear airframe engine
mounts. Ensure airframe pillow blocks are open and
slowly lower engine assembly aligning engine trunnion
bearings over engine mount pillow blocks.
(4) Align forward trunnion with forward engine
mount tube.
Install bolt through tube to forward
trunnion. Tighten bolt 50 TO 70 inch-pounds torque.
Lockwire (C151).
deck fitting.
(b) Secure clamp (28) to two fuel drain hose
clamps and clamp on engine mount tube.
(c) Connect fuel disconnect (14) to fuel filter.
(d) Connect oil hose disconnect (16) to oil
bypass valve.
(e) Connect oil drain disconnect (9) to engine
deck fitting.
(f) Connect fuel drain disconnect (8) to engine
(5) Close pillow blocks over left and right side
aft trunnion bearings. Tighten nuts on pillow blocks 50
TO 70 inch-pounds torque. Lockwire (C151).
(6) Connect electrical cable harness and cable
assemblies as follows:
deck fitting.
(g) Connect engine bleed air tube coupling (7)
to engine deck fitting.
(h) Connect coupling of oil breather hose
assembly (26) to engine oil tank.
(a) Connect plug of electrical harness
assembly (29, figure 4-3A) to aft right side of firewall
plug.
(i) Connect coupling of hose assembly (25) to
engine oil tank.
(b) Connect exhaust thermocouple harness
plug (30) to aft right side firewall plug.
(j) Install tube assembly (4) between bleed air
valve and bleed air fitting on top of engine.
(c) Connect plug of electrical harness
assembly (6) to aft left side firewall plug.
(k) Install bleed air tube and coupling (3)
between bleed air line and fitting.
(d) Connect plug of cable assembly (13) to
fuel filter plug.
(7) Install droop compensator control tube to droop
compensator using bolt (17), two washers (18), and nut
(19).
Tighten nut (19) and install cotter pin (20).
(e) Connect plug of cable assembly (15) to
4-4D
Change 71
TM 55-1520-234-23
(m)
Install power lever control tube to
fuel control lever using bolt (21), washer (22) and nut
(23). Tighten nut and install cotter pin (24).
trunnion (1).
d. Cut lockwire attached to bolts (5 and 6).
Remove lockwire.
(n) Install ejector, IR suppressor (if
installed), and tailpipe assembly. (Refer to paragraph 412 and 4-13).
11).
(o) Install main driveshaft.
paragraph 6-7).
(Refer to
f. Remove cambox assembly bracket (2) and
transducer bracket (3).
(p) Install particle separator.
paragraph 4-8).
(Refer to
(q) Install engine cowling and tailpipe
fairing.
Install transmission cowling.
(Refer to
paragraphs 2-16 and 2-15.)
(r) Connect battery.
4-5.1
e. Remove bolts (5 and 6), washers (4, 7, and
g. Remove trunnion (1).
4-5.4 REMOVAL-AFT
(TRUNNIONS).
MOUNT
4-5.3 REMOVAL-FORWARD
FITTING (TRUNNION).
ENGINE
Engine may be supported with suitable strap
running underneath combustion section and a
spreader bar to clear strap from aft pylon fairing.
b. Remove left aft fitting (trunnion) (13, figure 4-3B )
as follows:
CAUTION
MOUNT
a. Attach engine lifting sling (T7) to engine
assembly. Attach hoist (T49) to sling. Take up slack
between hoist and sling.
FITTINGS
NOTE
FITTINGS
Engine mount fittings (trunnions) are the part of the
engine mounts that are installed on the engine. The
forward fitting (trunnion) is bolted to the forward left side
mount pad. The two aft fittings (trunnions) are bolted to
the rear mounting pads, one left and one right.
MOUN T
a. Attach engine sling (T7) to engine assembly.
Attach hoist (T49) to sling. Take up slack in hoist cable to
relieve weight from engine mount.
ENGINE MOUNT FITTINGS (TRUNNIONS).
4-5.2 DESCRIPTION-ENGINE
(TRUNNIONS).
ENGINE
Ensure that engine is supported by hoist and sling
prior to removal of pillow block.
(1) Remove left pillow block assembly (17) by
procedure outlined in paragraph 2-44.
(2) Remove self-locking nuts (16), washer (15),
and bearing (14) from trunnion on left side of engine.
NOTE
Engine may be supported with suitable strap
running underneath combustion section and
spreader bar to clear strap from aft pylon fairing.
(3) Remove lockwire from four bolts (20). Remove
bolts (20) and washers (21).
(4) Remove trunnion (13).
b. Cut lockwire attached to bolt (8, figure 43B). Remove lockwire.
c.
c. Remove aft right engine mount fitting (trunnion) by
same procedure outlined in step b.
Remove bolt (8) and washer (9) from
Change 71
4-4E
TM 55-1520-234-23
Figure 4-3B. Engine Mount Fittings (Trunnions) Installation (Sheet 1 of 2)
4-4F
Change 71
TM 55-1520-234-23
Figue 4-3B. Engine Mount Fittings (Trunnions) Installation (Sheet 2 of 2)
Change 71 4-4G
TM 55-1520-234-23
Figure 4-3C. Damage Limits - Aft Engine Mount Fittings (Trunnions) and Bearings
4-5.5
INSPECTION - ENGINE MOUNT FITTINGS (TRUN NIONS).
a. Inspect aft engine mount fittings (trunnions) for
damage in excess of limits shown in figure 4-3C.
c. Inspect forward engine mount fitting (trunnion) (1,
figure 4-3B) for nicks, scratches, cracks and thread
damage. No cracks, thread damage or severe mechanical
damage is acceptable.
b. Inspect bearings on aft engine mount fittings
(trunnions) for damage and wear (looseness) in excess
of limits shown in figure 4-3C.
4-4H Change 71
TM 55-1520-234-23
4-5.6 REPAIR
OR
REPLACEMENT-ENGINE
MOUNT FITTINGS (TRUNNIONS).
a. Replace bearings if damaged or worn in excess
of limits (paragraph 4-5.5b).
(1) Remove nut (16, figure 4-38) washer (15)
and bearing (14).
(2) Position serviceable bearing (14) on fitting
(trunnion). Install thin steel washer (15) and self-locking
nut (16). Torque 50 TO 70 inch-pounds.
b. Replace engine mount fittings (trunnions) if
damaged in excess of limits (paragraph 4-5.5).
Torque bolts (5) and bolts (6) 290 to 410 inch-pounds.
e. Position forward engine mount (10) on trunnion (1)
and install special bolt (8) and thin steel washer (9). Torque
bolt (8) 50 to 70 inch-pounds.
f. Lockwire (C151) bolt (8) to upper forward bolt (6),
then to lower forward bolt (6). Lockwire (C151) lower aft
bolts (5), to upper aft bolts (5).
4-5.8 INSTALLATION-AFT ENGINE MOUNT FITTINGS
(TRUNNIONS).
NOTE
c. Polish out mechanical and corrosion damage
that is within limits shown on figure 4-3C with fine India
stone (C128). Touch up repair area with primer (C100
or C102).
If trunnions were removed from an installed
engine, the engine should be supported by a
hoist (paragraph 4-5.4).
4-5.7 INSTALLATION-FORWARD ENGINE MOUNT
FITTING (TRUNNION).
a. Install left aft trunnion (13, figure 4-3B) as follows:
(1) Position trunnion (13) on engine left aft mount
NOTE
pad.
If trunnion was removed from an installed
engine, the engine should be supported by a
hoist (paragraph 4-5.3).
(2) Install four bolts (20) and thin steel washers
(21). Torque bolts (20) 95 to 110 inch-pounds. Lockwire
(C151) bolts in pairs.
a. Position trunnion (1, figure 4-3B) on engine
forward left mount pad.
b. Install right aft trunnion in same manner outlined in
step a.
c.
b. Position cambox assembly bracket (2) on top
mounting holes of trunnion. Position transducer bracket
(3) on lower portion of trunnion (1) with washer (11)
between bracket and trunnion.
Remove hoist and engine sling if applicable.
d. Install main driveshaft (paragraph 6-7).
e. Install particle separator (paragraph 4-8).
c. Install two thin steel washers (4) and bolts (5).
Do not tighten bolts. If required, add a maximum of
three AN960C816 and/or AN960C816L washers
(paragraph 4-120) to obtain a flush fit between
transducer bracket (3) and trunnion (1).
f. Install engine cowling and transmission cowling
(chapter 2).
g. Connect battery.
h. Perform runup and maintenance test flight.
d. Install two bolts (6) and thin steel washers (7).
Change 71
4-5
TM 55-1520-234-23
Section II. COOLING SYSTEM
4-6.
Cooling System.
All power plant cooling is entirely automatic in
operation, having no direct controls.
a. Engine Internal Cooling. Internal parts of the
engine are cooled (and main bearing seals pressurized
against internal oil leakage) by compressed air diverted
from the main stream of air flow through the engine, and
by external air admitted through the hollow struts of the
exhaust diffuser. 'his cooling air passes out of the
engine with exhaust gases.
b. Engine External Cooling. The exterior of the
engine is cooled by air entering scoops on cowling doors
and flowing aft to the tailpipe area, where it is drawn out
through the ejector by the exhaust gas stream. (Refer to
Chapter 2 for further information on cowling, and to Section
IV of this chapter for description and maintenance of the
ejector.)
c. Starter - Generator Cooling. The starter- generator
is equipped with an integral fan and an inlet shroud
connected by a flexible hose to a screened intake duct,
mounted on right 'side of engine inlet housing, near the
cowling air scoop. The duct and hose must be kept free
from obstruction and securely mounted. (Refer to Chapter
9 for further information and maintenance of the startergenerator.)
Section III. AIR INDUCTION SYSTEM
4-7.
Air Induction System.
Engine intake air passes through grass filter screens
and large vertical scoops on both transmission cowling
doors, into a chamber enclosed by induction baffles and
the forward firewall. From this chamber, air is drawn
into the engine inlet through a particle separator which
removes particles of foreign matter.
4-8.
Particle Separator.
The particle separator is an inertial-type separator
consisting of an upper and lower assembly half, a FOD
screen, a deflector, a mounting ring assembly, a flange
assembly and seal, gaskets, and attaching hardware.
(See figures 4-4, 4-5 and 4-6.) Removal of the upper
assembly half permits maintaining aircraft drive shaft
and inspecting the engine inlet. The lower assembly
half mounts the air cleaner which collects particles
removed from the engine inlet air and ejects them
overboard. A flange assembly provides means of
attaching the separator to the engine inlet housing. The
foreign object damage screen consists of two halves which
fit around the sand and dust separator inlet to prevent large
foreign objects from entering the engine. Two latch
assemblies hold the halves together. (See figure 4-5.)
Engine inlet air passes through the FOD screen, where any
large particles are caught immediately, and enters the
separator through a curved, annular, radial inflow bellmouth
provided in the upper and lower assembly halves.
Separation occurs when the contaminated air is drawn
through a turn, causing particles to be forced to the
concave inner flow wall and caught by a protruding lip of
the deflector assembly. Clean air continues into the engine
inlet area while contaminated portion of the air is drawn
through a second turn causing further separation. The clean
air resulting from the second turn is returned to the engine
inlet area while particle- laden air flows into a large annular
chamber and through an air cleaner mounted on the lower
half of the separator. Engine compressor discharge (P3)
air from a fitting mounted on the engine air diffuser flows
through the venturi effect ejector and carries the particles
overboard
through
airframe
plumbing.
4-6
TM 55-1520-234-23
Figure 4-4. Particle seperator
4-7
TM 55-1520-234-23
Figure 4-5. Particle separator - exploded view (Sheet 1 of 2)
4-8
Change 65
TM 55-1520-234-23
Figure 4-5. Particle separator - exploded view (sheet 2 of 2)
Premaintenance Requirements
for Particle Separator
Conditions
Requirements
Model
Part. No. or Serial No.
Special Tools
Test Equipment
Support Equipment
AH-1S
All
None
None
Torque
Wrench
One
(C14) (C15)
(C35) (C87)
(C108) (C112)
(C124) (C138)
(C142)
None
Minimum Personnel Required
Consumable Materials
Special Environmental
Conditions
NOTE
It is not necessary to further disassemble the
separator unless the inspection procedures indicate
that gaskets and seals may be damaged. If further
inspection is required, proceed with the following
steps.
(4) Remove the lower half of the foreign object
damage screen. (Refer to paragraph 4-11.)
(5) Remove main drive shaft from aircraft as a
complete assembly, and remove curvic coupling adapter
from engine output shaft.
(6) Disconnect pressure and overboard plumbing
from air cleaner fittings.
(7) Remove five nuts (20, figure 4-5) and five
washers (21). Remove lower half(14) of separator and
deflector assembly (12).
a. Removal.
(1) Remove
top
half of foreign
(damage screen. (Refer to paragraph 4-11.)
(8) Remove 24 nuts (11), 24 washers (10), 24
sleeve spacers (9) and remove mounting ring assembly (4).
object
NOTE
(2) Release two latches (15,figure 4-5) and
latch assemblies (16 and 22) on front and rear faces of'
upper and lower separator halves (2 and 14) by
simultaneously pressing the safety latch up and lifting
the release catch. Release latch (1) on top of separator
upper half and remove the upper half.
Loosely install spacers, washers, and nuts on
engine inlet housing studs.
(9) Remove washers, screws, and split ring
assembly that secure mounting flange assembly (7) to
aircraft.
(3) Remove gasket assemblies (26).
Change 65
4-9
TM 55-1520-234-23
Figure 4-6. Airflow diagram
4-10
Change 7
TM 55-1520-234-23
(10)
Loosen V-band coupling and remove
mounting flange assembly.
(1) Repair loose seal and gaskets as follows:
Cleaning solvent is flammable and toxic.
Provide adequate ventilation.
Avoid
prolonged breathing of solvent vapors and
contact with skin or eyes.
Cleaning solvent is flammable and toxic.
Provide adequate ventilation. Avoid prolonged
breathing of solvent vapors and contact with
skin or eyes.
b. Cleaning. Clean parts only as required to
facilitate inspection using solvent (C124).
(a) Repair loose gasket (8, figure 4-5) by
recementing gasket to mounting flange assembly with
cement (C35). Clean mating surfaces with methyl-ethylketone (C87).
c.
Inspection.
(1) Inspection seal (5, figure 4-5) on mounting
flange assembly (7) for tearing and/or ripping at the
edges and for lack of adhesion.
(2) Inspect gasket (8) on each side of mounting
flange assembly (7) for lack of adhesion.
(3) Inspect gasket assemblies
permanent set and lack of adhesion.
(26)
for
a
(b) Repair loose gaskets (13 and 26, figure 45) by recementing gaskets to mating surface with adhesive
(C14). Clean mating surfaces with methyl-ethyl-ketone
(C87).
(c) Repair loose seal (5, figure 4-5) by
recementing seal to mating surface, using adhesive (C15).
Clean mating surfaces with trichloroethylene (C142) and
then with methyl-ethyl-ketone (C87).
(2) Replace damaged gaskets (13) as follows:
(4) Inspect all metal surfaces for cracks or other
damage.
(5) Inspect for loose or missing rivets. If rivets
are loose or missing in upper or lower assembly half
perform FOD inspection. Replace the assembly half or
repair in accordance with step d.
(6) Inspect for weld cracks or weld separation
(particularly in the area of the inlet vanes in both the
upper and lower assembly halves).
If cracks or
separation is evident, replace affected assembly half or
repair in accordance with step d.
(7) Inspect for damaged or inoperable safety
latches, damaged positioning pins (23), and angle
brackets (24 and 28). If damage is evident, replace the
affected assembly half or repair in accordance with step
d.
(a) Remove defective gasket and oil adhesive
from deflector assembly.
(b) Wipe all metal surfaces to be bonded with
lint-free gauze moistened (not dripping) with methyl-ethylketone (C87). Continue wiping surface, changing gauze
frequently, until gauze remains clean.
NOTE
All grease, oil, or other surface contaminants must
be removed from the bonding surface.
(c) Using
clean,
stiff
brush,
contaminants from surface of new gasket.
(8) Inspect air cleaner (19) for evidence of
erosion.
(9) Inspect all other parts for evidence of
erosion. Replace damaged parts.
NOTE
Porous surfaces which have been contaminated
with oil or grease cannot be satisfactorily cleaned to
ensure proper bonding and shall be discarded.
d. Repair or Replacement.
Change 7
remove
4-11
TM 55-1520-234-23
(d) Using clean applicator, apply a
continuous uniform film of adhesive (C14) to surfaces to
be bonded. Allow adhesive to thoroughly air dry for
approximately 3 hours.
flange assembly.
Do not stretch seal when
installing it for fit. If the seal is too large, it may
have developed an oversize set during shipment.
If so, cut the seal to the required circumferential
length and adhere as a strip with the "butt joint"
located at either the 6- or 12-o'clock position.
(e) After adhesive coating has dried, apply
a second uniform, continuous film of adhesive to
surfaces to be bonded. Allow adhesive to thoroughly air
dry for approximately 3 hours.
(c) Using sandpaper (Cl12), roughen seal
surfaces to be bonded.
(f) Align surfaces to be bonded to obtain
contact over entire surface.
(d) Using lint-free gauze, thoroughly clean all
surfaces to be bonded with trichloroethylene (C142).
(g) Apply light load to surfaces being
bonded. Allow adhesive to cure under this pressure for
a minimum of 4 hours.
(3) Replace damaged seal (5, figure 4-5) on
mounting flange assembly (7) as follows:
(e) Apply a uniform layer (0.010 TO 0.030 inch
thick) of adhesive (C15) to surfaces to
be bonded.
(a) Remove defective seal. Remove old
adhesive film from metal surfaces with a knife blade.
Complete cleaning with sandpaper (C112).
(f) Fit seal to flange of the mounting flange
assembly and press surfaces together.
NOTE
Use only enough pressure to displace air, but not
so much that the adhesive is forced out of the joint.
(g) Allow adhesive to cure undisturbed.
NOTE
Use trichloroeythelene and methyl- ethyl-ketone
in a well ventilated area. Avoid prolonged
breathing of vapors and do not use in an area
with open flame or high temperature.
Under light pressure, the adhesive will take 24
hours to cure. Under warm, damp conditions the
adhesive may cure sufficiently in 4 hours to permit
reinstallation of mounting flange assembly.
(b) Wipe metal surfaces to be bonded with
lint-free gauze, moistened (not dripping) with
trichloroethylene (C142) followed by methyl- ethylketone (C87).
Continue wiping, changing gauze
frequently, until gauze remains dry.
(4) Replace damaged gasket (8, figure 4-5) on
each side of support of mounting flange assembly (7) as
follows:
(a) Remove gasket material.
CAUTION
To ensure proper bonding of seal, ensure that
all grease, oil, or other surface contaminants are
removed.
Provide adequate ventilation when using methyleythl-ketone. Avoid breathing solvent vapors and
avoid prolonged skin contact.
NOTE
(b) Wipe metal surfaces to be bonded with lintfree gauze moistened (not dripping) with methyl-ethylketone (C87).
Continue wiping and changing gauze
frequently, until gauze remains clean.
Determine whether new seal is proper size by
fitting it to metal mating flange of mounting
4-12
Change 7
TM 55-1520-234-23
CAUTION
For proper bonding of gasket, ensure that all
grease, oil, and other surface contaminants are
removed.
(c) Fit new gasket (0.047 inch thick) to
mounting flange assembly.
(d) Using sandpaper
surface of gasket to be bonded.
(C112)
roughen
(e) Wipe all gasket surfaces to be bonded
with lint-free gauze moistened (not dripping) with
methyl-ethyl-ketone (C87) to remove all powder and
surface contaminants. Allow gasket to dry for 15
minutes.
(f) Apply a continuous uniform film of
cement (C35) to both metal and rubber surfaces to be
bonded. Allow approximately 2 to 3 hours drying time.
(g) Wipe the surface of one adhesive film
with gauze moistened with methyl-ethyl-ketone (C87),
one section at a time. The reactive surface should
immediately become tacky.
(h) Align mating surfaces, one section at a
time, to obtain contact over entire surface, and press
tacky surface to dry surface. Allow adhesive to cure
under light pressure for a minimum of 4 hours.
Change 7 4-12A/(4-12B blank)
TM 55-1520-234-23
(a) Remove two rivets and remove latch.
(5) Replace air cleaner (19, figure 4-5) as
follows:
(a) Remove six nuts (17) and six washers
(18). Remove air cleaner from lower assembly half.
(b) Position new air cleaner on lower
assembly half and secure with six washers (18) and six
nuts (17).
(b) Assemble upper and lower assembly
halves without gaskets installed between halves.
(c) Position latch on lower assembly half in
line with hook or upper assembly half. Mark lines to
position latch.
(d) Separate assembly halves.
(6) (AVIM) Repair nonconverging cracks by
stopdrilling crack ends. Where necessary to prevent air
leakage, seal with tape or silicone rubber, RTV (C108).
Torque nuts 30 to 35 inch-pounds.
(e) Position latch on lower assembly half within
marked lines, and drill 0.128 to 0.133 inch holes for rivets.
NOTE
(7) (AVIM) Use rivets or tack welds to patchrepair converging cracks. Follow standard airframe
sheet metal repair procedures.
If rivet holes are elongated during drilling or if new
holes are drilled, back up sheet metal with doubler
before riveting.
NOTE
(f) Secure latch with two rivets.
Any standard aluminum aircraft rivets, including
Huck, Cherry, etc., of proper size may be used
for all rivet repairs except hardware
replacement.
(13) Replace damaged positioning pins or angle
brackets on lower assembly half as follows:
(a) Remove rivets, bracket and spacer.
(8) (AVIM) Patch-repair punctures too large to
repair with silicone rubber. Follow standard airframe
sheet metal repair procedures. Secure patches with
rivets or tack welds.
(9) (AVIM) Repair torn tack and spot welds with
tack, interrupted, or plug weld repairs; use doublers as
needed. See figure 4-7. Rivet repairs, using doublers
as needed, are also acceptable.
(10) Repair serious erosion damage
replacing damaged parts or using doublers.
latches
on
upper
and
lower
assembly
(c) Position spacer and bracket on lower
assembly half in line with bracket on upper assembly half.
Top of bracket will be slightly higher than the edge of the
assembly half. Mark lines to position
spacer and bracket.
(d) Separate assembly halves.
by
(11) Reshape deformed parts if feasible. If
reasonable conformity cannot be achieved, particularly
at mating edges, sealing surfaces, etc., replace part.
(12) Replace damaged
assembly half as follows:
(b) Assemble
halves without gaskets.
(e) Position spacer and bracket within mark
lines and drill two 0.128 to 0.133 inch holes for rivets.
NOTE
Use vinyl tape (C138) between spacer and bracket.
lower
(f) Secure bracket to spacer with two rivets.
NOTE
Tape (C130) patch between hardware items and
assembly half must be replaced if damaged
during removal of hardware.
(g) Align assembly halves and identify center
on bracket in line with hole of bracket on upper assembly
half.
Change 71
4-13
TM 55-1520-234-23
Figure 4-7. Doubler for inlet vane reinforcement
4-14
Change 7
TM 55-1520-234-23
(h) Drill a 0.250 to 0.252 inch hole through
bracket.
(3) Remove spacers, washers, and nuts from
engine inlet housing studs and discard.
(i) (AVIM) Install new pin in bracket, and
tack weld in two places, 180 degrees apart.
(4) Position mounting ring assembly (4) on engine
inlet housing studs.
(14) (AVIM) Replace new pin in bracket, and
tack assembly half as follows:
NOTE
Ensure that five studs on ring assembly are at the
bottom, with the center stud located at the 6 o'clock
position.
(a) Remove rivets, and remove hook.
(b) Assemble upper and lower assembly
halves without gasket assemblies.
(c) Position hook on upper assembly half in
line with latch on lower assembly half. Mark lines to
position.
(d) Separate assembly halves.
(5) Secure mounting ring assembly with 24 sleeve
spacers (9), 24 washers (10), and 24 nuts (11). TORQUE
NUTS TO 70-80 INCH-POUNDS.
(6) Position deflector assembly (12) over locating
pins and studs on mounting ring assembly (4) and press in
until firmly seated.
(e) Position latch within marked lines and
drill 0.128 to 0.133 inch holes for rivets.
NOTE
NOTE
Secure lines to ejector inlet and discharge ports
before installing lower assembly half (14).
Use vinyl tape (C138) between hook and
mounting surface.
NOTE
(7) Position lower assembly half (14) on locating
pins and studs on mounting ring assembly (4). Secure with
washers (21) and nuts (20). TORQUE NUTS TO 30-35
INCH-POUNDS.
If rivet holes are elongated during drilling or if
new holes are drilled, back up sheet metal with
doubler before riveting.
NOTE
Washer (21) and nut (20) may be omitted from 6
o'clock position to eliminate having to remove main
drive shaft and engine adapter at next separator
removal.
(f) Secure hook with two rivets.
e.
Installation.
(1) Wipe engine inlet housing clean with clean
cloth moistened with solvent (C124).
(2) Position mounting flange assembly (7,
figure 4-5) in front of airframe firewall and on engine
inlet housing. Retain mounting flange loosely with Vband coupling and on firewall with split ring assembly,
P/N 204-060-868-7. Insert screws with washers from
back of firewall to secure ring assembly.
NOTE
Leave mounting flange assembly (7) loose
enough to be rotated
(8) Position upper assembly half (2) on lower
assembly half (14).
NOTE
Do not install two gasket assemblies (26) at this
time.
(9) Rotate mounting flange assembly (7) on inlet
housing to align hook assembly (6) with latch (1) on upper
assembly half (2). Secure mounting flange assembly (7)
with V-band coupling. TORQUE V-BAND COUPLING NUT
TO 40-50 INCH-POUNDS. Tap around coupling from
middle toward each end with a soft-faced mallet to seat
properly.
Change 65
4-15
TM 55-1520-234-23
NOTE
Install V-band clamp assembly (34) with clamp
latch on left side of engine and gap (drain) at
the 6 o'clock position.
(10)
Tighten
assembly to firewall
screws
to
secure
flange
4-9.
Foreign Object Damage Screen.
The foreign object damage screen is mounted on the
forward side of the particle separator and prevents large
foreign objects from entering particle separator.
a. Removal. (Refer to figures 4-8 and 4-9.)
(11)
Remove upper assembly half (2).
(12)
Install curvic coupling in output shaft of
engine and install main drive shaft if removed.
(1) Remove top half of FOD screen from the
particle separator as follows(a) Unlock both latches.
(13)
Connect pressure
plumbing to air cleaner fittings
(14)
and
overboard
(b) Disengage hook portions.
Install baffle panels.
(c) Lift screen free of the sand and dust
separator.
(15)
Position gasket assemblies (26) over
positioning pins (23) on lower assembly half.
(16)
lower half.
If FOD screen is to be installed, install
NOTE
Do not remove lower half of screen during periodic
inspection unless additional inspection is required.
(17)
Position upper assembly half (2) on
lower assembly half (14).
(2) If required, remove upper assembly half of
particle separator. (Refer to paragraph 4-8.)
NOTE
(3) Remove bottom half of FOD screen from the
sand and dust separator as follows:
'Tilt top slightly forward to position assembly or four
positioning pins (23).
(a) Lift forward split portion of the butt molding
free of the vane and hold in that position.
(18)
Engage upper assembly half (2) to
mounting flange assembly (7) with latch (1).
(b) Lift rear (notched) portion free of the curled
inlet of the sand and dust separator.
(19)
Engage latch assemblies (16 and 22) on
front face and latch assemblies (15) on rear curl of
separator
other side.
CAUTION
Ensure that safety catch on latches is
engaged by exerting a slight pull on release
catch. Catch should not open.
(20)
Check for proper seating of seals by
appearance. Approximately 1/8 inch of rubber on
gasket assemblies will be uniformly exposed. Seal (5)
on flange assembly will be approximately half way
compressed.
(c) Repeat preceding steps (a) and (b) for the
(d) Withdraw bottom half of FOD screen from
under the sand and dust separator.
b. Cleaning
(1) Clean all parts as
inspection, using solvent (C124).
c.
required
to
facilitate
Inspection.
(1) FOD screen for damage which would permit
foreign object entry.
(2) Aft
(21)
Install top half of FOD screen. (Refer to
paragraph 4-9.)
4-16 Change 65
molding
for
cuts
or
other
damage.
TM 55-1520-234-23
Figure 4-8. Removal/installation foreign object damage screen
(3) Latch assemblies for damage as follows:
(a) Erosion or damage that may cause tightness
or binding.
(b) Cracks.
(c) Loose or missing rivets.
(4) Replace FOD small mesh screen.
(a) Cut screen (1560-AH-1-080-3)
lengthwise. (See Appendix D.)
in
half
(4) FOD screen for deformation.
Maximum allowable overlay is 2 mesh openings
and .070 inch gap.
d. Repair and Replacement.
(1) Reshape deformed parts, if feasible. If
reasonable conformity cannot be obtained, replace
either half or both as required.
(2) Replace parts having severe damage or
mutilation.
(3) Replace screen halves with missing or loose
rivets.
NOTE
To prevent overlapping of screen on back edges
FOD screen some wedge cuts in screen will be
necessary.
(b) Line screen (1560-AH-1-080-3) along leading
edge of FOD screen. Form screen to existing FOD
screen.
Change 22 4-17
TM 55-1520-234-23
Figure 4-9. Procedural steps installing foreign object damage screen (bottom half)
4-18
TM 55-1520-234-23
(c) Using safety wire (C151), secure one end
and single wire lace every fifth opening along outer
perimeter of screen (1560-AH-1-080-3) to existing FOD
screen.
(d) Repeat steps (3) and (4) for other half of
FOD screen.
(5) Replace damaged hooks on upper screen as
follows:
(a) Remove rivets and remove hook.
(b) Assemble upper and lower screens and
position hook in line with latch on lower screen.
(c) Mark lines to position hook.
(d) Separate assembly halves.
(e)
Position latch within marked lines and
drill 0.128 to 0.133 inch holes for rivets.
NOTE
If rivet holes are elongated during drilling or if new
holes are drilled, back up sheet metal with a
doubler before riveting.
(f) Secure hook with two rivets.
(6) Replace damaged latches on lower screen as
follows:
(a Remove rivets and latch.
(b Assemble upper and lower screens and
position latch on lower screen in line with hook on
upper screen.
(c) Mark lines to position latch.
(d) Separate assembly halves.
(e) Position latch within marked lines and drill
0.128 to 0.133 inch holes for rivets.
NOTE
If rivet holes are elongated during drilling or if new
holes are drilled, back up sheet metal with a
doubler before riveting.
(f) Secure latch with two rivets.
e. Installation. (Refer to figures 4-8 and 4-9).
Ensure integrity of repairs. Materials used to
make repairs may be ingested by engine if not
properly secured.
(1) Position bottom half of the foreign object
damage screen, aft molding side toward engine inlet,
under the sand and dust separator so butt molding is
approximately 2-1/4 inches below horizontal centerline,
and aft molding is seated over the particle separator
split flange.
Improper seating of the aft molding over the
separator split flange can result in cuts or other
damage to the molding as well as placing excessive
stress on all portions of the screens and latches. To
check for proper seating, run hand along the lower
split flange to ensure that the molded channel is
properly fitted over both sides of the split flange.
(See figure 4-8.)
(2) Insert aft molding while holding butt molding
away from the vane in the separator. (Refer to steps I
and II, figure 4-9.)
(3) Line up the slot in forward portion of butt
molding with the vane over which it is to be fitted, and
press into place. (Refer to step III, figure 4-9.)
NOTE
When properly installed, the notched area of the
butt molding should be positioned behind the sand
and dust separator inlet curl, and the forward
portion of the molding should have one part of the
split on the top of the vane and one part underneath
the vane as shown in figure 4-9.
(4) If removed, install top half of the particle
separator. (Refer to paragraph 4-8e.)
(5) Position top half of the FOD screen so as to
engage the aft screen molding slot over the separator
split flange. Position the screen cut out over the latch at
the 12-o'clock position of the separator.
(6) Secure top half to the bottom by engaging and
locking both latches.
NOTE
Both latches must be engaged with the mating
hooks before closing either latch to a locked
position.
4-10. Air Induction Baffle.
The air induction baffle consists of panels
enclosing the engine air inlet area, ahead of the
Change 38 4-19
TM 55-1520-234-23
d.
Repair or Replacement. Make temporary
repair of cracks by drilling a small hole just beyond end
of crack. Replace any panel if damaged so that baffle
cannot be securely fastened or will allow foreign
material to enter induction area.
e.
Installation.
(1) Lift floor and forward baffle panels into place
from right side of transmission compartment. Align on
brackets on engine firewall and pylon fifth-mount
support. Secure fasteners.
(2) Install particle separator lower air filter, main
drive shaft, and upper air filter.
engine forward firewall and within the transmission
cowling. (See figure 4-10.) Panels are formed of
aluminum sheet with riveted doublers, clips, fasteners
and edge seals.
a. Removal.
(1) Open transmission cowling.
(2) Replace fasteners and pull shaft access
panel from left front of forward induction baffle panel.
(3) If so equipped, detach support clamps of
UHF-VHF antenna leads for clips on top and forward
baffle panels. Keep attaching parts with clamps.
(4) Release fasteners and remove top panel.
(3) Place top panel on upper edge of forward
panel and on two firewall brackets,
and secure
fasteners.
(5) Remove upper air filter assembly of particle
separator, main drive shaft, and lower air filter
assembly.
(4) If so equipped, attach support clamps of
UHF-VHF antenna leads to clips on top and forward
baffle panels,
using attaching parts previously
removed.
(6) Release fasteners and remove forward and
floor panels from right side of compartment.
b. Cleaning. Clean all panels with solvent (C124).
c. Inspection. Panels for cracks, damaged seals,
general condition, and security of fasteners.
(5) Position shaft access panel in opening at left
front of baffle. Insert inboard edge into slip joint of
forward panel. Align and secure fasteners.
Section IV. EXHAUST SYSTEM
4-11. Exhaust System.
The exhaust diffuser section, on rear of the
engine,
provides passage for gas flow from the
combustion chamber. The exhaust passage is extended
aft and slightly upward by a tailpipe and an ejector
assembly. A thermocouple assembly, mounted on the
diffuser, has probes in the exhaust stream to provide
continuous indication through the exhaust temperature
gage system. For purpose of engine compartment
cooling, a heat shield is mounted around the end of the
diffuser, the tailpipe is covered by an insulation blanket,
and the ejector surrounds the exhaust gas flow with a
cooling air stream. For helicopters which have IFR
suppression systems installed, refer to paragraph 4-13.
4-12. Tailpipe, Ejector and Heatshield.
The exhaust tailpipe and the heatshield are
mounted on flanges of the engine exhaust diffuser by Vband clamps. (See figure 4-11.)
Premaintenance Requirements
Ejector and Heatshield
Conditions
Model
Part No. or Serial No.
Special Tools
Test Equipment
Support Equipment
Minimum Personnel Required
Consumable Materials
Special Environmental
Conditions
for
Tailpipe
Requirements
AH-1S
All
None
None
None
One
(C116)(C124)
None
a. Removal.
(1) Disconnect drain lines from fittings or tailpipe
and ejector.
4-20
TM 55-1520-234-23
Figure 4-10. Air induction area (cowilng omitted)
(2) Remove ejector and tailpipe fairing as follows:
(a) If ejector is attached to tailpipe, remove
screws and washers at five sets of brackets to detach
and remove ejector. Release fasteners along lower
edge of fairing, and lift off fairing rearward.
(b) If ejector is attached to fairing, release
fasteners and remove fairing with ejector attached.
Detach ejector from fairing by removing screws and nuts
at three mounting clips.
(3) Loosen nuts to open V-band clamp which
secures tailpipe to engine diffuser flange. Pull aft to
disengage locating pins, and remove tailpipe.
(4) When required, remove lockwire and unwrap
insulation blanket from tailpipe.
(5) Remove V-band clamp and heat shield from
flange of diffuser support cone.
Change 2 4-21
(6) Cover exhaust diffuser opening.
TM 55-1520-234-23
Figure 4-11. Exhaust system components
4-22
TM 55-1520-234-23
b. Cleaning. Clean tailpipe and ejector with wire
brush and solvent (C124) when necessary. Clean
clamps and heat shield with solvent. Do not use solvent
on insulation blanket.
c.
Inspection.
(1) Tailpipe and ejector for cracks,
burned out or buckled areas.
d. Repair or Replacement. (AVIM)
(1) Ejector
and
scratches
NOTE
Prior to riveting new angle to ejector to tail pipe,
align flange of new angle with flange of angle to tail
pipe. Mark location for attachment bolt hole in new
angle using mating hole in tail pipe angle as a
guide. Maintain flat plane between the two angles
within 0.020 inch TIR.
(3) Heat shield for cracks and distortion.
dents
1 Drill out six rivets which attach angle to be
repaired. Use rivet holes in ejector as a guide to align
new angle, mark locations and drill six Number 30
(0.128) holes in angle.
dents, and
(2) Insulation blanket for visible damage.
(a) Shallow
disregarded.
(e) Replace brackets as follows:
may
be
(b) (AVIM) Cracks three inches or less in length
may be welded. Welds must be ground down as
smooth as possible to match contour of ejector.
(c) (AVIM) Cracks in excess of 3 inches, tears
or' small bullet holes in any area of the ejector may be
repaired by stopdrilling ends of each crack or tear and
welding a patch of. titanium, MIL-T- 9046, Type 1,
Composition C, over the crack or tear on both inside
and outside of the ejector. Patch edge distance to be a
minimum of 0.5 inch beyond stop-drill.
(d) Cracks in the bellmouth-to-tube weld joint at
angle bracket may be repaired and future cracking
eliminated as follows:
1 Remove angle bracket and spotwelded
doubler. Fabricate a new doubler from titanium (item
61, table 2-1), of same width but 0.5 inch longer than
removed doubler. Position doubler so that it extends
past cracked weld joint up onto bellmouth 0.5 inch.
Form doubler to fit contour.
2 Clean and reweld cracked joint. Grind weld
smooth. Place new doubler in position and weld to
assembly.
3 Attach angle bracket per d(l)(e) below, using
rivet holes in ejector as a guide to drill through doubler.
Change 22
2 Drill one number 1 (0.288) hole in angle
flange at location determined in note above. Attach new
angle to ejector with six rivets.
(f) Burned out parts or dents which cannot be
smoothed out to original contour are cause for
replacement.
(2) Tailpipe.
(a) Scratches and shallow dents may be ignored.
(b)
(AVIM) Cracks or tears in any area of
the tailpipe may be repaired by stop drilling ends of
each crack, and welding a patch of steel (item 59 table
2-2) over the crack or tear, if damage area does not
exceed 4.0 inches, after clean-up. Adjacent repairs on
exterior surface must have at least 2.0 inches of parent
material between patches. Patch edge distance must
be a minimum of 0.5 inch beyond the stop-drill. See
figure 4-12. The patch shall be welded on both inside
and outside of the tailpipe. If the tailpipe ring is welded,
care must be taken that the flange is filed or machined
to a flat surface after welding to provide a flat seat
against the attachment point on the engine.
(c) Burned out parts or damage greater than that
which is repairable by patching is cause for
replacement.
(d) Replace insulation blanket if damaged.
(3) Heat shield.
(a) Replace heat shield for cracks or distortion
that cannot be straightened.
4-23
TM 55-1520-234-23
(4) Position heat shield against flange of diffuser
support cone. Secure with V-band clamp around mating
flanges. TORQUE CLAMP NUTS 40 TO 50 INCHPOUNDS.
(5) Install ejector and tailpipe fairing as follows:
Adjustment of latch mating fitting may be necessary
to ensure proper closing of latch. Do not force latch
closed. Slight tension shall remain on latch in the
closed position.
(a) Place ejector into fairing, align clips and
secure at three locations with screws and nuts. If
necessary, remove plug buttons Detail A, figure 4-11)
and loosen bolts to realign clips on serrated plates,
tighten bolts and reinstall plug buttons. Install fairing
with attached ejector.
(b) Measure gap between tailpipe fairing,
engine cowl doors and pylon fairing. Inspect for a gap
of 0.040 inch at the lower extremities of cowl
components to a maximum of 0.190 inch at tile top. If
adjustment is necessary, proceed as follows:
1 Loosen two attachment screws securing latch
plate to tailpipe fairing bracket (Detail B, figure 4-11 ).
Latch shall remain fastened during adjustment.
2 Adjust fairing to proper clearance allowing
serrated plate to self-adjust.
Figure 4-12. Engine tailpipe repair
e. Installation.
3 After proper clearance is obtained tighten two
attachment screws.
(1) Remove protective cover from engine
exhaust diffuser.
(2) Position tailpipe, with drain fitting down, on
flange of exhaust diffuser. Make sure locating dowels
are engaged, and that inside of pipe aligns with diffuser.
Secure with V-band clamp around mating flanges.
(6) Connect drain lines to fittings on tailpipe and ejector.
(7) Seal all open areas between base of forward and/or
aft firewall and engine deck per figure 4-13 using
sealant (C116).
(a) Seat clamp by tapping with soft mallet,
from middle toward ends, while tightening nuts to 100
TO 130 INCH-POUNDS TORQUE.
(b) Check torque again after first ground run-up.
(3) If removed, install insulation blanket around
tailpipe with joint at top. Secure with lockwire (C151)
installed in zig-zag pattern. (See figure 4-11).
4-13. Infra-red Exhaust Suppression System.
The infra-red exhaust suppression system
includes an upturned insulated exhaust duct
assembly, an exhaust extension and forward duct
assembly. For helicopters which do not have IRS
systems installed, refer to paragraph 4-12.
4-24
TM 55-1520-234-23
Figure 4-13. Engine deck and firewall sealing
4-25
TM 55-1520-234-23
Premaintenance requirements for infrared suppression
system.
Conditions
Model
Part No. or Serial No.
Special Tools
Test Equipment
Support Equipment
Minimum Personnel Required
Consumable Materials
Special Environmental
Conditions
Requirements
AH-1S
All
None
None
None
Two
(C17) (C56) (C87)
C106A) (C124)
None
a.
Removal.
(1) Disconnect drain line (6, figure 4-14) from
fitting on exhaust tailpipe (1).
(2) Release fasteners on tailpipe fairing (4) and
remove tailpipe fairing, duct assemblies (2 and 5) and
exhaust extension (3) as an assembly.
(3) Detach forward duct assembly (2) from
exhaust extension (3) by removing first V-band clamp
(8).
(4) Detach exhaust extension (3) from aft duct
assembly (5) by removing second V-band clamp (8).
(5) Remove 10 screws and washers and remove
fairing assembly (7) from aft duct assembly (5) and
fairing (4).
(6) Detach aft duct assembly (5) from tailpipe
fairing (4) by removing 12 bolts and washers.
(7) Loosen nuts on third V-band clamp (8) which
secures exhaust tailpipe (1) to engine diffuser (9). Pull
aft to disengage locating pins and remove exhaust
tailpipe.
(8) Remove lockwire (11) and unwrap insulation
blanket (10) from exhaust tailpipe (1).
engine diffuser opening to prevent entry of foreign
material.
Cleaning solvent is flammable and toxic. Provide
adequate ventilation. Avoid prolonged breathing of
solvent vapors and contact with skin or eyes.
b. Cleaning.
Clean exhaust tailpipe (1),
duct
assemblies (2 and 5) and exhaust extension (3) with a
wire brush and solvent (C124). Clean clamps, heat
shield, and tailpipe fairing (4) with solvent (C124). Do
not use solvent on insulation blanket (10).
c. Inspection.
(1) Forward exhaust duct and exhaust extension.
(a) Inspect ducts for cracks,
dents,
and
overheating.
1 Cracks and holes in the surface of the duct
should not exceed 3.0 inches in diameter after cleanup.
Adjacent repair areas must allow a minimum of 2.0
inches of parent metal between patches.
2 Heating
as
evidenced
merely
by
discoloration of the metal is permissible. However, if
the condition becomes progressive,
indicating a
possible burn-through, the part should be replaced.
3 Dents in the surface of the duct are
permissible, providing the surface is not broken, and
there are no sharp creases or projections into the
exhaust stream.
4 Damage to circumferential mounting frames
shall be evaluated locally as to feasibility to repair or
need for replacement. It is deemed not feasible to
replace the frame on the duct.
(b) Ensure that all attaching bolts are secure.
(c) Ensure that drain hose on forward exhaust
duct is intact and securely attached.
NOTE
See figure 4-11 for view of heat shield.
(9) Remove third V-band clamp (8) and heat
shield from engine diffuser (9). Cover
4-26 Change 7
(2) Aft exhaust duct.
TM 55-1520-234-23
Figure 4-14. Exhaust infrared suppression system
Change 2
4-27
TM 55-1520-234-23
(a) Inspect interior and exterior surface of duct
for cracks, nicks, dents, and overheating.
1 Cracks and holes in the interior surface of the
duct shall not exceed 3.0 inches in diameter after
cleanup. Adjacent repair areas must allow a minimum
of 2.0 inches of parent metal between patches.
2 Cracks and holes in the exterior surface of
the duct shall not exceed 4.0 inches in diameter after
cleanup. Adjacent repair areas must allow a minimum
of 2.0 inches of parent metal between patches.
3 Damage which penetrates the temp-mat
insulation between the interior and exterior surfaces of
the duct shall not exceed 4.0 inches in diameter after
cleanup.
4 Dents in the interior or exterior surfaces are
permissible, providing the surfaces are not broken, and
there are no sharp projections into the exhaust stream.
5 Damage to the mounting flange and outlet
rim of the duct shall be evaluated locally as to feasibility
to repair, or need for replacement.
(b) Check duct and supporting fairing for security
of mounting and for loose or missing rivets.
d. Repair or Replacement. A repair kit (C106A) is
available for use in repairing damaged components.
(1) Dents.
(a) Minor dents in exterior surfaces require no
rework if the surface if not broken, or if no sharp crease
or projection exists in the interior surface.
(b) Work out minor dents having sharp
projections into the interior of the duct, by restoring to
original contour and smoothing off any sharp projections
with fine abrasive paper.
(c) Large dents (surface impressions) shall be
worked out by restoring the surface to original contour.
(2) Cracks.
(a) All cracks shall be stopdrilled to prevent
continuation.
(b) Cracks which penetrate interior surfaces of
the forward exhaust duct or exhaust extension, shall be
repaired using method I. See table 4-1 and figure 4-15,
detail A.
(c) Cracks in interior surface of aft exhaust duct
shall be repaired using method II. See table 4-1, and
figure 4-15, detail B.
(d) Cracks in exterior of aft duct shall be
repaired using method III. See table 4-1.
(e) Two or more adjacent cracks, or two or more
converging cracks, shall be treated as a single repair
area, within limitations specified in Section III.
(3) Holes.
(a) Holes in the forward exhaust duct, or in the
exhaust extension, shall be repaired using method I.
See Table 4-1.
(b) Holes in the interior surface only of the aft
exhaust duct, shall be repaired using method II. See
Table 4-1.
(c) Holes which penetrate only the outer surface
of the exhaust duct, shall be repaired using method III.
(d) Holes which penetrate completely through
exterior and interior surfaces of aft exhaust duct, shall
be repaired using method II. See Table 4-1.
Table 4-1. Repair Methods
METHOD I. (See figure 4-15, detail A.)
a. Stop drill crack at both ends.
b. If two or more cracks converge, cutout area encompassed by cracks and smooth out edges to form a hole,
not to exceed 3.0 inches.
4-28 Change 7
TM 55-1520-234-23
Table 4-1. Repair Methods (Cont)
METHOD I. (See figure 4-15, detail A.) (Cont)
c. Trim patch, 205-706-083-3, as necessary, to provide a minimum of 0.25 inch edge distance between rivets
and hole in parent metal, and between rivets and edge of patch.
d. Position patch and layout rivet hole locations.
e. Drill rivet holes and install rivets, (item 56, table 2-2), with heads on interior of duct.
METHOD II. (See figure 4-15, detail B.)
a. Cut hole through outer surface of duct to encompass damaged area, not to exceed 4.0 inches in diameter.
b. Cut out and remove insulation material (temp-mat) not to exceed 4.0 inches.
c.
Cut out damaged area of interior surface of duct, not to exceed 3.0 inches in diameter.
d. Trim patch, 205-706-083-3, as necessary to provide a minimum of 0.25 inch edge distance between rivets
and hole, and between rivets and edge of patch.
e. Position patch on outside of interior surface, and layout rivet hole locations.
f.
Drill rivet holes, and install rivets, (item 56, table 2-2) with heads on interior of duct.
g. Cut patch, 205-706-0835, to fit insulating core area.
h. Apply adhesive paste, No. 19, and install patch.
i.
Trim patch, 205-706-083-7, to provide a minimum of 0.50 inch overlap on outer surface of duct.
Provide adequate ventilation when using methylethylketone. Avoid breathing solvent vapors and avoid
skin contact.
j.
Lightly abrade area 0.60 inch wide around hole in outer surface. Clean area with methyl-ethyl-ketone (C87).
k.
Apply adhesive (C17) to area around hole.
l.
Remove backing from patch, and apply over hole. Press smoothly into place.
m. Allow adhesive to cure.
NOTE
Preferred cure time is 24 hours at room temperature. Cure time may be reduced by
applying heat not to exceed 175 degrees F (79 degrees C) for one hour.
Change 7 4-29
TM 55-1520-234-23
Table 4-1. Repair Methods (Cont)
METHOD III. (Figure 4-15, Detail B, View C-C.)
a. Cut out damaged area of outer skin, not to exceed 4.0 inches in diameter.
b. Trim patch, 205-706483-7, to provide a minimum of 0.50 inch overlap on outer surface of duct.
Provide adequate ventilation when using methyl-ethyl-ketone. Avoid breathing solvent vapors
and avoid prolonged skin contact.
c.
d.
e.
f.
Lightly abrade area 0.60 inch wide around hole. Clean area with methyl-ethyl-ketone (C87).
Apply adhesive (C17) to area around hole.
Remove backing from patch, and apply patch over hole. Press smoothly into place.
Allow adhesive to cure. (See note in Method II.)
e. Installation.
Use extreme care in positioning aft duct
assembly (5, figure 4-14). Failure to do so
will damage gasket.
(1) Loosely install third V-band clamp (8) on
exhaust tailpipe (1).
(2) Install the two insulation blankets (10) with
lockwire (11) on exhaust tailpipe (1).
NOTE
See figure 4-11 for view of heat shield.
(3) Engage locating pin and install heat shield
and exhaust tailpipe (1) on engine diffuser (9) with third
V-band clamp (8). Seat V-band clamp by tapping with a
soft mallet, from middle toward ends, while tightening
nuts 100 TO 130 inch-pounds.
(4) Position aft duct assembly (5) on rear flange
of tailpipe fairing (4) with lower external supports
extending over the external supports mounted on the
fairing transition section. Support aft duct assembly (5)
with 12 bolts and washers. Position fairing assembly (7)
in place and secure to aft duct assembly (5) and tailpipe
fairing (4) with 10 screws and washers.
(5) Position exhaust extension (3) inside fairing
(4) and mount to forward section of aft duct assembly
(5) with second V-band clamp (8). Seat V-band clamp
(8) by tapping with a soft mallet, from middle toward
ends, while tightening nuts 100 TO 130 inch-pounds.
(6) Install forward duct assembly (2) on forward
section of exhaust extension (3) with first V-band clamp
(8). Seat V-band clamp (8) by tapping with a soft
mallet, from middle toward ends, while tightening nuts
100 TO 130 inch-pounds.
(7) Place tailpipe fairing (4),
with exhaust
extension (3), and duct assemblies (2 and 5) installed,
in place and over tailpipe (1). Engage fasteners around
tailpipe fairing (4).
(8) Connect drain line (6) to fitting on tailpipe (1).
Page 4-30A/4-30B blank deleted
4-30 Change 7
TM 55-1520-234-23
Figure 4-15. Repair procedures - infrared suppression
Change 7 4-31
TM 55-1520-234-23
Section V. - OIL SYSTEM
4-14. Oil System.
is controlled by a float switch in the oil tank and a switch
on the pilots console. If oil level in the tank becomes
low enough to operate the float switch, (3.8 quarts low
from spill over) the ENGINE OIL BYPASS caution panel
segment will light, and the valve will automatically shut
off flow to the cooler and return engine oil directly to the
tank. The pilot can use his switch to reopen the valve
and use oil cooling when conditions warrant such action.
Oil is supplied to the engine from a self-sealing
tank mounted in the aft fairing above the engine. (See
figure 4-16). After passing through the engine, oil is
delivered from the scavenge side of the engine oil pump
to an oil cooler bypass valve mounted on the engine
compartment deck at left side. In normal operation, oil
passes through the oil cooler and then returns to the
tank. The cooler is mounted below an engine deck
opening, under a metal plenum. Cooling air is drawn in
through a screened duct on left side of the fuselage up
through the cooler, then aft to pass out through
screened openings of the tailpipe fairing. A turbine fan
driven by engine bleed air is mounted under the cooler.
The oil cooler emergency bypass valve
4-14A. Troubleshooting - Oil System.
Table 4-2 is provided as an aid in troubleshooting
malfunctions in the oil supply system.
Table 4-2. Troubleshooting - Engine Oil System
NOTE
Before using this table, be sure all normal operational checks have been performed.
CONDITION
TEST OR INSPECTION
CORRECTIVE ACTION
Do not operate engine until it is determined that oil pump failure or oil starvation has not occurred.
1. No engine oil pressure.
STEP 1. Ensure that tank is filled to proper level.
Fill tank to proper level if required. (Refer to paragraph 1-3a.)
STEP 2. Check for loose connection and/or clogged hose.
Inspect entire lubrication system for leaks and obstruction. Pay particular attention to quick
disconnect fittings.
4-32 Change 2
TM 55-1520-234-23
Table 4-2. Troubleshooting - Engine Oil System (Cont)
CONDITION
TEST OR INSPECTION
CORRECTIVE ACTION
STEP 3.
Check for proper operation of oil pressure transmitting system.
Check system using pressure source at pressure tap. Replace oil pressure transmitter if faulty.
(Refer to paragraph 8-12f.) Replace oil pressure indicator if faulty. Check continuity of wiring
circuit between transmitter and indicator. (Refer to paragraph 8-11 c and 8-12 c.)
STEP 4.
Check that oil pump coupling is not sheared and female spline on oil pump driveshaft gear is not worn.
Replace oil pump for sheared coupling. Inspect for worn spline on shaft gear. (Refer to TM 55-2840229-23.)
2.
Fluctuating oil pressure.
STEP 1.
Check oil quantity in tank.
Fill tank to proper level. (Refer to paragraph 1-3a.)
STEP 2.
Check for dirty piston in oil pump pressure regulating valve.
Remove, clean, and reinstall piston. (Refer to TM 55-2840-229-23).
STEP 3.
Check oil pump and/or oil pump driveshaft gear for failure.
Remove and replace oil pump or oil pump driveshaft gear. (Refer to TM 55-2840-229-23).
STEP 4.
Check for faulty transmitter or circuit to indicator.
Perform continuity check and replace components as necessary. (Refer to paragraph 8-12c and 812e.)
3.
Low engine oil pressure.
STEP 1.
Check for low quantity in tank.
Fill tank to proper level. (Refer to paragraph 1-3a.)
STEP 2.
Check oil pump pressure regulating valve for proper adjustment.
Adjust regulating valve. If no response, remove valve, clean and reinstall. (Refer to TM 55-2840-22923.)
STEP 3.
Check for clogged oil filter.
Clean oil filter. (Refer to TM 55-2840-229-23.)
Change 29
4-32A
TM 55-1520-234-23
Table 4-2. Troubleshooting - Engine Oil System (Cont)
CONDITION
TEST OR INSPECTION
CORRECTIVE ACTION
STEP 4.
Check oil pressure transmitter for faulty operation. (Refer to paragraph 8-14c.)
Remove and replace faulty transmitter. (Refer to paragraph 8-14f.)
STEP 5. Check oil pump for faulty operation.
Remove and replace faulty oil pump. (Refer to TM 55-2840-229-23.)
4.
High engine oil pressure.
STEP 1.
Check for restrictions in oil flow lines.
Check quick disconnect couplings for proper connections. Clear oil lines of restrictions.
STEP 2.
Check oil pump pressure regulating valve for proper adjustment.
Adjust regulating valve. If no response, remove valve, clean and reinstall. (Refer to TM 55-2840-22923.)
STEP 3.
Check oil pressure transmitter for faulty operation. (Refer to paragraph 8-14c.)
Remove and replace faulty oil pressure transmitter. (Refer to paragraph 8-14f.)
NOTE
High engine oil pressure may be due to cold oil on start.
Allow engine to reach operating temperature by
operating engine at idle.
5.
High engine oil temperature.
STEP 1.
Check that oil cooler bypass valve is not stuck in bypass position.
If valve is stuck in bypass position, remove and replace valve. (Refer to paragraph 4-19a.)
STEP 2.
Check that emergency bypass valve is not stuck in bypass position.
If valve is stuck in bypass position, remove and replace valve. (Refer to paragraph 4-19a.)
STEP 3.
Ensure that oil cooler blower is operating correctly.
Check for proper blower operation. Repair and replace as necessary. (Refer to paragraph 4-18a.)
4-32B
Change 29
TM 55-1520-234-23
Table 4-2. Troubleshooting - Engine Oil System (Cont)
CONDITION
TEST OR INSPECTION
CORRECTIVE ACTION
STEP 4.
Ensure that cooling air inlet is not blocked.
Inspect screened inlet and remove all grass, leaves, and other foreign material. Also, check cooler
core air passage and remove any grass, dirt or other foreign material. (Refer to paragraph 4-7 and
4-8.)
STEP 5.
Check engine oil level.
Fill tank to proper level. (Refer to paragraph 1-3a.)
STEP 6.
Ensure that temperature indicating system is operating correctly.
Check operation of oil temperature indicator, resistance bulb and related circuitry. Replace faulty
components. (Refer to paragraph 8-49, 8-50, and 8-49f. and 8-50e.)
STEP 7.
Check for bleed air leak between engine and oil cooler turbine fan.
Isolate and correct any leaks.
STEP 8.
Check operation of low level emergency system float switch in oil tank. (Refer to paragraph 4-14.)
Replace faulty float switch in oil tank. (Refer to paragraph 4-14e.)
6.
Low oil temperature.
STEP 1.
Check that oil cooler bypass valve is not stuck in failure to bypass position.
Repair or replace valve as necessary. (Refer to paragraph 4-19a.)
STEP 2.
Check operation of temperature indicating system.
Check operation of oil temperature indicator, resistance bulb, and related circuitry. Replace faulty
components. (Refer to paragraph 8-49, 8-50, and 8-49f. and 8-50e.)
7.
No oil temperature.
STEP 1.
Check operation of temperature indicating system.
Check operation of oil temperature indicator, resistance bulb, and related circuitry. Replace faulty
components. (Refer to paragraph 8-49, 8-50, and 8-49f, and 8-50e.)
Change 2 4-32C
TM 55-1520-234-23
Table 4-2. Troubleshooting - Engine Oil System (Cont)
CONDITION
TEST OR INSPECTION
CORRECTIVE ACTION
8.
Excessive engine oil consumption.
STEP 1.
Check for leakage at fittings and hose connection.
Tighten or replace fittings or hose assemblies.
NOTE
Refer to TM
55-2840-229-23 for additional excessive engine oil consumption
troubleshooting procedures.
4-15. Pressure Transmitters and Switch. (Refer
to Chapter 8.)
4-16. Engine Oil Tank.
The engine oil system tank is a self sealing cell
equipped with a filler cap, an oil level sight glass, and a
scupper with drain line. (See figure 4-16.) The tank is
located in the aft fairing, secured by bolts on a
horizontal firewall above the engine, and accessible for
service through right side of transmission cowling.
Bosses on bottom of the tank provide mounting for an
outlet coupling, a drain valve, and a float switch. A
plate on top of the tank provides connections for a vent
tube, and for oil return and engine breather line tubes.
NOTE
The oil level sight gage is provided for the
purpose of determining a low oil condition.
When oil level is at sight gage level, oil supply
is 2.75 ± .25 quarts low. When servicing oil
tank, fill completely to a spill over condition.
a. Removal.
(1) Open engine compartment cowling. Remove
center fairing to allow access through front of aft fairing.
Prolonged contact with lubricating oil may cause
a skin rash. Those areas of skin and clothing
that come in contact with lubricating oil should
be thoroughly washed immediately. Areas in
which lubricating oil is used should be
adequately ventilated to keep mist and fumes to
a minimum.
(2) Place a suitable vessel under drain line.
Drain oil tank by opening valve.
(3) Disconnect hoses from outlet coupling and
coupling of breather line and return line. Remove
breather and return lines, and vent tubes from top plate
of tank.
(4) Disconnect drain tubes from tank scupper and
from drain valve. Remove valve from tank boss.
(5) Disconnect electrical leads of float switch from
relay on underside of firewall. Remove support and float
switch assembly and outlet coupling, with gaskets,
from tank bosses.
(6) Remove three bolts and washers to detach
tank from horizontal firewall.
4-32D Change 29
TM 55-1520-234-23
(7) Remove oil tank. Cover open ports and ends
of tubes.
(4) Install drain valve
Connect scupper drain line.
and
connect
tube.
b. Cleaning. Clean oil tank with solvent (C124).
Drain off solvent and dry with filtered compressed air.
(5) Assemble float switch, with gasket, on
support. Install support, with gasket into tank boss.
Connect electrical leads to relay located on underside of
firewall near tank. (Refer to Appendix F for circuit wiring
diagram.)
(6) Install outlet coupling, with gasket, in tank
boss. Connect hose from engine oil pump inlet.
Do not pressurize tank with compressed air.
c. Inspection. Tank for evidence of damage or
leaks; bosses and fittings for damaged threads. Sight
glass for clear condition.
(7) Connect hose from engine breather (on
accessory drive gear box) to coupling on tube. Connect
oil return hose to coupling on tube.
d. Repair or Replacement.
(8) Service tank.
cowling.
(1) Replace packings or gaskets under tank
fittings when leaks occur.
Check for leaks.
Reinstall
4-17. Engine Oil Cooler.
(2) Replace any damaged fittings.
The engine oil cooler is mounted on underside of the
engine compartment deck. Its inboard side is attached
to mating flanges of the transmission oil cooler, but
there is no oil connection between the two coolers.
(3) Replace tank if punctured, cut or otherwise
damaged.
(4) For repair of self-sealing tank refer to TM 551500-204-25/1.
NOTE
If oil cooler is known to have been contaminated
with metal particles. Replace cooler and tag
removed cooler as being contaminated.
Forward to depot.
e. Installation.
NOTE
Check protrusion of three indicator pins to
ensure security of quick-disconnect couplings.
(1) Check that filler cap, scupper, sight glass,
and top plate are securely installed on tank.
Use back-up wrenches when removing and
installing oil cooler drain fittings, valves and
lines.
(2) Place tank in position on horizontal firewall of
aft fairing, with filler cap forward. Align holes and install
three bolts, with washers, through firewall into threaded
inserts of tank.
Premaintenance Requirements for
Engine Oil Cooler.
Conditions
Requirements
(3) Install and connect vent tube.
Connect
breather tube and return tube to top plate of tank.
Model
Part No. or Serial No.
4-33
AH-1S
All
TM 55-1520-234-23
b. Cleaning. (A VIM)
Premaintenance Requirements for
Engine Oil Cooler (Cont)
Conditions
Requirements
Special Tools
Test Equipment
Support Equipment
Minimum Personnel Required
Consumable Materials
None
None
None
One
(C14), (C43)
(C90), (C110)
(C124)
Dust Free
Special Environmental
Conditions
When using steam and compressed air, be
careful not to damage air fins by high
pressures.
(1) Steam clean the exterior surfaces and
corrugated air fins of each core. Remove obstructions
from air fins with a pick and compressed air.
(2) Prepare oil cooler for internal cleaning as
follows:
a. Removal.
(a) Remove lockwire and unscrew by-pass
control valve body from valve housing in cooler.
(b) Press a rubber plug into the by-pass opening
in the valve housing.
(c) Reinstall by-pass valve into valve housing so
valve body bears up against the rubber plug.
(3) Connect oil cooler in line with cleaning
equipment in reverse of normal flow for first flush. See
figure 4-17.
(1) Remove oil cooling duct from left side and
access door from right side of fuselage. Remove turbine
fan and duct. (Refer to paragraph 4-19 a.)
(2) Drain trapped oil from cooler.
(3) Disconnect oil lines from cooler fittings. Cap
open lines.
NOTE
Centrifugal pump in cleaning equipment must
be capable of supplying fluid at approximately
40 gpm while maintaining pressure of 75 psi.
(4) To remove oil and loose sludge and to reduce
contamination of cleaning solutions during following
operations, pre-clean cooler interior as follows:
Prolonged contact with lubricator oil may
cause a skin rash. Those areas of skin and
clothing that come in contact with lubricating
oil should be thoroughly washed immediately.
Areas in which lubricating oil is used should
be adequately ventilated to keep mist and
fumes to a minimum.
(4) Remove bolts, nuts and washers that attach
engine oil cooler to transmission oil cooler.
Use solvent in a well ventilated area. Avoid
prolonged breathing of vapors and do not use
in an area with open flame or high
temperature.
(5) Remove bolts and washers around
mounting flange of engine oil cooler. Remove cooler
from fuselage.
(a)
Flush core, in. reverse direction, with
solvent (C124) for 30 minutes or until solvent appears
clean.
(6) If cooler is being replaced, remove inlet
and outlet fittings for use on replacement assembly.
4-34
TM 55-1520-234-23
Figure 4-16. Engine oil supply and cooling system with bleed air-driven fan
Change 2 4-35
TM 55-1520-234-23
Figure 4-17. Oil cooler cleaning equipment setup - typical
(b) Reverse lines to cooler and flush core in
direction of normal flow for approximately 15 minutes.
(c)
Remove oil cooler from cleaning equipment
and drain all fluid from cooler.
(5) Remove dirt, carbon deposits, oil gum, lead
deposits, and other contaminants by connecting oil
cooler to cleaning equipment. See figure 4-17. Use
cleaning compound (C43).
(d) Flush oil cooler in normal direction for 15
minutes to clean by-pass passage.
(e) Remove plug from cooling section opening in
valve housing and install into by-pass opening.
Reinstall by-pass valve.
(a) Flush core 30 to 60 minutes in direction
opposite to normal flow.
(b) Reverse lines and flush core in normal
direction for 15 minutes.
(c)
Remove plug installed in by-pass
opening of valve housing and insert plug in cooling
section opening. Reinstall by-pass valve.
Use solvent (C124) in a well ventilated area.
Avoid prolonged breathing of vapors and do
not use in an area with open flame or high
temperature.
(6) Connect oil cooler to cleaning equipment
containing cleaning solvent (C124). Install 100-mesh
screen at inlet and outlet ports of oil cooler.
(a) Flush core for 10 minutes in each direction.
4-36 Change 7
TM 55-1520-234-23
(b) Check 100-mesh screens between each
flush.
2 Connect other end of air line to an adjustable
air pressure source and adjust pressure to 12 psig.
(c) If screens are not clear, reflush core for 5
minutes in each direction, repeat until screens are
clear.
3 Submerge cooler in water at approximately
140°F (60°C). Gradually heat water to 180°F (85°C)
and check for leaks.
(7) Remove rubber plug from by-pass valve
housing in oil cooler.
c. Inspection.
(1) Inspect air fins and air passages for distortion
and foreign particles that may obstruct air flow.
Increase air pressure gradually to avoid burns
from leakage and hot water.
(2) Inspect cooler for damaged or bulged plates,
cracked castings and flanges,
and broken welds.
Inspect studs for stripped threads-and cracked or
ineffective lock rings.
4 After 5 minutes of submersion, gradually
increase air pressure to 100 psig. Inspect cooler for
leaks as evidenced by presence of air bubbles.
(3) Inspect all openings in oil cooler for evidence
of foreign matter inside of the cooler.
(4) Inspect rubber gaskets on top of cooler for
security, rips, tears, or scores, and missing sections
that may prevent gaskets from sealing.
5 Remove cooler from water and relieve air
pressure.
(b) Make final check of oil cooler as follows:
1
Dry oil cooler externally using compressed
air.
(5) Inspect by-pass control valve and valve
housing for stripped threads and distortion, scoring, or
wear of the seat surfaces. Check functioning of by-pass
control valve as follows:
2 Plug outlet port and apply room temperature
water at 400 psig to the other (inlet) port.
(a) Submerge valve in water heated to 150° to
155°F for 5 minutes. Valve should open.
4 Inspect cooler for any visible leaks and blown
or bulged plates.
3
(b) Remove valve and measure length.
5
(c) Submerge valve in water heated to 176° to
1800F for 5 minutes. Valve should open.
(d) Remove valve and measure length.
Minimum increase in valve length is 0.090 inches.
(6) With by-pass
pressure-check oil cooler.
control
valve
installed,
(a) Make preliminary check of oil cooler for air
leaks as follows:
1Plug outlet port and connect an air line to inlet
port.
Lock liquid in oil cooler for 10 minutes.
Release pressure and drain cooler.
d. Repair or Replacement. (AVIM)
(1) Replace oi lcooler if damage, other than minor
distortion to air fins, is detected or if cooler fails to meet
inspection requirements. Tag oil cooler as applicable
and forward to depot.
(2) Repair minor bends or distortion of accessible
air fins with flat duckbill pliers.
(3) Replace missing, torn, ripped, and scored
gaskets with rubber (C110). Attach gaskets to oil cooler
surface with rubber adhesive (C14). Replace by-pass
control valve seal.
Change 2 4-37
TM 55-1520-234-23
(6) Service oil tank. Check for leaks and proper
operation at next ground run.
(4) Replace damage or faulty by-pass control
valve.
(5) (AVIM) If oil cooler is serviceable,
thoroughly with preventive oil (C90) as follows:
flush
f.
Preparation for Storage and Shipment.
(1) Flush oil cooler thoroughly with preventive oil
(C90).
NOTE
(2) Place cooler in container which will prevent
damage during shipment or storage.
The interior of the cooler should be completely
dry before final flush with preventive oil to
prevent fouling of the mixture.
4-18. Oil Cooling Turbine Fan.
(a) Connect cooler to cleaning equipment
containing preventive oil and 100-mesh screens.
A turbine fan driven by engine bleed air is used to
blow air through the engine and transmission oil coolers.
(See figure 4-16.) The fan is suspended on an adapting
duct under the coolers, in the fuselage compartment
below the engine deck.
(b) Flush oil through cooler in each direction for
10 minutes.
(c) Check 100-mesh screens between each flush
to ensure that no metal particles have appeared.
Premaintenance Requirements for
Oil Cooling Turbine Fan.
(d) Drain cooler and install plugs in both inlet
and outlet ports. Secure by-pass control valve with
lockwire.
Conditions
Model
Part No. or Serial No.
Special Tools
Test Equipment
Support Equipment
Minimum Personnel Required
Consumable Materials
e. Installation.
(1) If replacing cooler, install inlet and outlet
fittings with new gaskets.
(2) Position cooler on underside of support,
below engine deck, and secure with bolts and washers.
Install bolts through mating flanges of engine and
transmission oil coolers, and secure with nuts and
washers.
Special Environmental
Conditions
Requirements
AH-1S
All
None
None
None
One
(C37) (C45)
(C90) (C102)
(C124)
None
a. Removal.
(3)
Install turbine fan. (Refer to paragraph 4-18
(1) Remove oil cooling duct from left side of
fuselage. (See figure 4-16.) Remove access door from
right side.
1.)
(2) Disconnect bleed air hose from inlet fitting of
turbine fan. Cover open end of hose.
Check proper alignment of flared ends of
tubing to valves and fittings. Do not allow
preloading or stresses due to misalignment or
improper fit.
NOTE
Keep attaching parts in sets,
changing adjustment.
(4) Align and connect oil lines to cooler fittings.
and avoid
(3) Detach two brace tubes and brackets from
flange of fan.
(5) Install cooling duct and access door.
4-38
TM 55-1520-234-23
(4) Remove eight screws and washers that
secure fan to bottom flange of duct. Remove fan
assembly.
(5) Match-mark four hanger brackets to mating
brackets on sides of duct. Detach each hanger bracket
from duct bracket by removing two bolts and washers.
(8) Detach screen if installed from fan assembly
by removing two remaining bolts with nuts, washers and
vibration-isolating grommets. When replacing turbine
fan, also remove inlet fitting and cap open port.
b (A VIM) Disassembly (Janitrol). Disassemble
turbine fan, P/N 158001-1 as follows: See figure 4-18.
(6) At forward and rear sides, detach duct
brackets from lower bolts that secure flanges of two
coolers together. Remove duct. Reinstall bolts through
cooler flanges, adding washers as needed.
(2) and remove cover and bellmouth assembly
(1) from housing (16).
(7) Detach four hanger brackets from bolts that
secure coolers to deck structure.
(2) Remove nut (14) and washer (15) from end
of shaft (10).
(1) Remove nuts (3), washers (4), and bolts
Figure 4-18. Oil cooling turbine fan (Janitrol)
Change 22
4-39
TM 55-1520-234-23
(3) Remove nut (6) and washer (7) from shaft
(10); then remove fan and turbine assembly (5) and key
(11) from shaft.
(4) Cut lockwire and remove four screws (9) and
retainer (8) from housing.
(5) Carefully pull shaft (10) with bearings (12 and
13) from housing, as a unit.
may be used to dislodge stubborn deposits. Wipe clean
and dry with filtered compressed air.
(2) Remove corrosion deposits on shaft (10,
figure 4-18) and housing (16) bearing liners using fine
crocus cloth
(C45).
Clean parts after removing
corrosion with corrosion preventive oil (C90).
e.
(Deleted.)
f.
(A VIM) Inspection (Janitrol).
NOTE
Do not remove identification plate (17) or
rotation directional arrows from housing
unless damaged.
(6) Using a suitable bearing puller,
bearings (12 and 13) from shaft (10).
c.
d.
remove
(Deleted.)
(A VIM) Cleaning (Janitrol).
(1) Clean all parts with lint-free cloths saturated
with solvent (C 124). A soft bristle brush
(1) Visually inspect all parts for nicks, burrs,
scratches, dents and weldment cracks and for evidence
of excessive wear.
(2) Inspect ball bearings (12 and 13, figure 418) for wear or damage.
(3) Inspect fan and turbine assembly (5) for
cracks, nicks, and scratches and for bent or cracked
fan blades.
(4) Inspect parts for dimensional tolerances.
See table 4-3.
(5) General pitting to a depth of 0.060 inch is
acceptable on the fan housing.
g. (Deleted.)
All data on page 4-41 including Figure 4-19 deleted
4-40
Change 22
TM 55-1520-234-23
Table 4-3. Dimension Tolerance - Turbine Fan
FIG.
NO.
INDEX
NO.
4-18
1
Cover & Bellmouth Assy
Nozzle
4-18
10
Shaft
Replace if front end bearing journal is not within
0.6695 TO 0.6691 inch diameter or if rear end
bearing journal is not within 0.4726 TO 0.4722
inch diameter.
4-18
16
Housing
Replace if front bearing liner I.D. is greater than 1.3791
inches or if rear bearing liner I.D. is greater than 1.1034
inches.
4-18
Deleted
4-19
Deleted
NOMENCLATURE
REMARKS
JANITROL
Replace if throat diameter is over 0.324 inch.
h. Repair or Replacement. (Janitrol) (AVIM)
(3) Replace bearings (12 and 13) if they do not
meet inspection criteria.
(1) Remove burrs and blend minor nicks and
scratches from fan with a fine india (C128) or
carborundum stone (C128).
(4) Replace components which do not meet the
dimensional tolerances set forth in table 4-3.
(5) Replace nuts (3,
condition.
Do not attempt to remove nicks or scratches
from turbine blades. If the turbine blades are
damaged, replace fan and turbine assembly
(5, figure 4-18).
(2) Refinish all exposed aluminum surfaces,
after repair, with chemical film (C37) and repaint with
one coat of primer (C102) as required.
6 and 14) regardless of
(6) Match-drill any replacement hanger brackets
at installation, with 0.280 TO 0.297 inch diameter holes
through ends to match existing bolt holes in cooler
flanges and deck structure, and 0.266 TO 0.263 inch
diameter holes in lower legs of hanger brackets to
match existing holes in brackets on duct.
4-42 Change 65
TM 55-1520-234-23
i.
(Deleted.)
j. (AVIM) Assembly. (Janitrol) Assemble turbine
fan, P/N 158001-1 as follows: See figure 4-18.
(5)
Install washer (7) and nut (6) on shaft and,
holding fan and turbine assembly to prevent rotation,
torque nut 115 TO 140 inch-pounds.
CAUTION
(6) Install washer (15) and nut (14) on shaft
and torque nut 48 TO 55 inch-pounds.
Do not force bearings into housing. If bearings do not slip
in place with slight hand pressure, check bearing liners
for burrs or corrosion.
(7) Position cover and bellmouth assembly (1)
on housing (16) and secure with bolts (2), washers (4),
and nuts (3).
(1) Press bearings (12) on shaft (10) to seat
firmly against shoulder on shaft. Insert shaft and
bearing into housing (16).
(2) Press bearing (13) on shaft (10) and into
housing (16).
(3) Position retainer (8)in housing with four
screws (9). Tighten screws and secure with lockwire.
(4) Install key (11) in shaft (10) and install fan
and turbine assembly (5) on shaft, align keyway in fan
with key in shaft.
Change 71
k.
(Deleted)
l.
Installation.
(1) At each side of cooler, attach two hanger
brackets on second bolt from each end in row of five bolts
that attach cooler flange to deck structure.
(2) At forward and rear sides of coolers,
remove lower bolts that attach coolers to each other. Lift
duct to position under coolers. Align bolt holes of two
duct brackets to holes in cooler
4-43
TM 55-1520-234-23
flanges.
Install bolts from left side, using washers,
between brackets and flanges and thin washers under bolt
head and nuts.
(11) At next ground run-up,
for proper operation.
4-19.
(3) Attach each hanger bracket to mating
bracket on duct with two bolts and washers.
check installation
Engine Oil Bypass Valve.
A two-position motorized valve, located on left side of the
engine compartment deck, is connected in the engine oil
return line. (See figure 4-16.)
a. Removal.
Use of incorrect reducer fitting (204-060-4941) in bleed air line may cause blower
overspeed.
(1)
Open engine compartment cowling at left
side.
(4) Check installation of restrictor fitting with
preformed packing in air inlet port of turbine fan.
Use of a screen on turbine fan is optional. A
defective screen may be removed and
discarded. If screen is not installed, ensure
decal (WARNING- HIGH SPEED FAN) is
installed and is visible from side duct opening.
(5) Position screen over intake end of fan
assembly, with screen cut-out next to air inlet. Secure
two screen legs located nearest cutout on bolts through
fan flange. On each bolt install a grommet and a thin
washer between flange and screen, and use a washer
under bolt head and under nut.
(6) At two bolt holes of fan flange that align
with two remaining legs of screen, install bracket instead
of grommets and install bolts in same manner as in step
(6).
(7) Position fan assembly under duct, and
align with air inlet at left side pointing aft. Install eight
screws and washers to attach fan to duct.
Prolonged contact with lubricating oil may
cause a skin rash. Those areas of skin and
clothing that come in contact with lubricating oil
should be thoroughly washed immediately.
Saturated clothing should be removed
immediately. Areas in which lubricating oil is
used should be adequately ventilated to keep
mist and fumes to a minimum.
(2) Remove oil cooling duct from left side of
fuselage. Drain oil cooler and lines.
(3)
Disconnect electrical cable connector from
valve.
(4) Disconnect engine scavenge oil hose from
valve coupling. Disconnect oil cooler lines and tank
return line from valve fittings. Cap open ends of lines.
(5) Detach valve from brackets by removing
two screws at each side. Lift out valve assembly.
(6) Remove fittings from valve by removing
attachment screws. Remove check valve and gasket
from right-hand fitting.
(8) Attach two brace tubes between brackets
on right side of fan and on fuselage structure. If
necessary, adjust inboard ends of braces.
(9)
Connect engine bleed air hose to fan inlet
fitting.
(10) Reinstall access door and cooling duct on
fuselage openings.
Change 7
Cleaning solvent is flammable and toxic.
Provide
adequate
ventilation.
Avoid
prolonged breathing of solvent vapors and
contact with skin or eyes.
4-44
TM 55-1520-234-23
Do not immerse valve motor in solvent.
b Cleaning.
solvent (C124).
Clean valve and attaching fittings with
Change 7
4-44A/(4-44B blank)
c. Inspection. Mating surfaces of valve and fittings
for nicks and burrs and threads for damage.
d. Repair or Replacement. Replace any damaged
fittings or attaching parts. Replace valve assembly if
malfunction occurs.
(4)
TM 55-1520-234-23
Connect electrical cable to connector on
valve.
(5) Service oil tank. Reinstall oil cooling duct
with screws. Close cowling.
(6) Check for leaks and proper operation at
next ground run.
e. Installation.
4-20.
NOTE
Be sure flow arrow is toward fitting.
(1) Assemble fittings on valve with attaching
screws.
Use new gasket when installing check valve
in return line fitting.
Engine Chip Detector.
A chip detector is installed in the lower right side of the
accessory drive gearbox. This unit is wired into the
master caution panel. (Refer to TM 55-2840-229-23.)
(2) Position valve assembly between brackets
and install two screws at each side.
(3) Connect oil cooler lines and tank return line
valve fittings. Connect engine scavenge oil hose to
coupling.
No more than 15 INCH-POUNDS OF TORQUE
shall be applied to the chip detector centerpost
nut when installing the chip detector wire.
Section VI. IGNITION SYSTEM
Refer to TM 55-2840-229-23 for maintenance instructions.
Section VII. POWER CONTROLS
quadrant marked with power settings in its range of travel
between stops which are pre-adjusted by the engine
manufacturer or overhaul facility. The linkage is a series
A mechanical linkage system, actuated by twist-grips
of control rods, bellcranks, idlers and levers. Control
on collective pitch control sticks, provides manual control
rods (6 and 26), at each end of the series, are
of the power lever on the engine fuel control unit. The
adjustable. Bellcrank (15) has an adjustable connection
power lever modulates the engine from zero to full power
for control rod (14), which determines the travel of
by controlling the gas producer (N1) turbine rpm.
linkage above the bellcrank. An adjustable cam (16) will
make contact with the spring-loaded plunger of a solenoid
4-22. Power Lever Control Linkage.
(19) to arrest linkage motion at the flight idle position
when power is being reduced from higher settings. The
The power lever shaft (28, figure 4-20) is serrated
solenoid plunger can be
and grooved to accept a control arm (27), and has a
4-21.
Power Controls.
Change 29
4-45
TM 55-1520-234-23
Figure 4-20. Power lever control system
Change 2
4-46
TM 55-1520-234-23
retracted by use of the ENGINE IDLE STOP REL,
pushbutton switch, located on the pilot's collective stick to
allow control movements in the OFF position.
a. Removal.
(1) To remove control rod (6, figure 4- 20):
Remove Screw-mounted access panel from left side of
fuselage in line with lower end of pilot's collective control
stick. Detach rod from bellcrank on control stick and
bellcrank (7) by removing bolts with nuts, washers and
cotter pins.
(2) Leave lower linkage (7 through 13) in place
for normal maintenance and inspection. If necessary to
replace damaged parts, obtain access by removing
screw-mounted panels from lower skin and detach parts
by removing bolts, washers and nuts.
(3) To remove bellcrank (15) and cam (16):
Obtain access to compartment, below engine and behind
aft fuel cell, by removing oil cooler air intake duct from
left side of fuselage. Disconnect control rods (14 and 24)
from bellcrank, keeping attaching parts with rod-ends.
Remove attaching bolt and lift out bellcrank with cam
attached. To remove cam, use an allen wrench to
remove two special bolts and serrated washers.
(4) To remove solenoid (19): Obtain access as
in step (3). Disconnect electrical connector from solenoid.
Remove four bolts and washers to detach solenoid
assembly and base (18) from bulkhead (17).
(5) To remove control rod (24) and boot (21):
Loosen upper clamp on boot. Disconnect control rod from
bellcrank (15) and lever (25). Remove rod with retainer
(22) attached. Remove snap-ring and split bushing (23)
and separate retainer from rod. Remove boot and clamps
from housing (20).
(6) To remove lever (25): Disconnect control
rods. Remove bolt with nut, washer, spacer and cotter
pin to detach bellcrank from engine mount pillow block.
(7) To remove control arm (27): Disconnect
control rod (26). Remove lockwire and screw from arm.
Pull arm from fuel control power lever shaft (28). Keep
screw with arm.
Change 7
Cleaning solvent is flammable and toxic.
Provide
adequate
ventilation.
Avoid
prolonged breathing of solvent vapors and
contact with skin or eyes.
b. Cleaning. Clean external surfaces of parts by
wiping with a cloth moistened with solvent (C124). Do not
permit solvent to enter bearings or solenoid.
c.
Inspection.
(1) Control rods for cracks and general
condition, end fittings for security, bearings for binding or
rough operation.
(2) Bellcranks, levers and idlers for security,
cracks or damage, and binding or rough bearings.
(3)
Solenoid for security and proper operation.
(4)
Boot assembly for cracks and wear.
d. Repair or Replacement.
(1) Replace control rods for cracks, distortion,
or loose or binding end fittings.
(2) Replace bellcranks, levers and idlers for
cracks, or for loose or binding bearings.
(3) Replace solenoid if malfunction occurs.
Secure replacement solenoid in bracket with four
countersunk screws, using shims as required so that
solenoid plunger operates freely in bracket bushing.
(4)
Replace boot if cracked or damaged.
e. Installation.
(1) Place arm (27, figure 4-20) on fuel control
power lever shaft (28), aligned with stop arm. Install
screw through arm and groove of shaft. After rigging
controls, lockwire screw.
4-47
TM 55-1520-234-23
(7) Connect control rod (14) to bellcrank
(13) with bolt, two thin washers, nut and cotter
pin. At upper end, insert bolt through large
safety washer,
rod-end,
aluminum-alloy
washer,
slot of bellcrank and a serrated
washer. Set rod-end at middle of bellcrank slot
and secure with nut and washer. Install cotter
pin after rigging is complete.
(2) Position lever (25) on outboard side of engine
mount pillow block, with spacer between lever and block
and with marked arm of lever pointing up. Insert bolt
through large safety washer, lever bearing, spacer and
pillow block. Secure bolt with thin washer, nut and cotter
pin at inboard end.
(3) Adjust control rod (26) to nominal length
of 11.17 inches between centers of rod end bearings.
Attach adjustable end of lever (25) with bolt,
washers, nut, and cotter pin. Forward end of rod will
be corrected to arm (27) in rigging procedure.
(4) Place smooth side of cam (16) on bellcrank
(15). Place a serrated washer on each of two special
bolts and start through slotted holes of cam and into
nutplates of bellcrank. (See figure 4- 21.) Use allen
wrench at shank ends to tighten bolts. Place bellcrank in
support and install bolt, thin washers, nut and cotter pin.
Cam position will be adjusted during rigging.
(5) Attach solenoid assembly (19, figure 4- 20)
and base (18) to mounting holes in bulkhead (17) with
four bolts and washers. Set solenoid position so that
plunger will not engage stop cam, until ready for
adjustment during rigging.
NOTE
(8) Adjust control rod (6) to nominal length of
20.0 inches between center of rod-end bearing and clevis.
Position clevis, with rod offset inboard, on throttle control
bellcrank of pilot's collective stick (3). Install bolt from
inboard side, with thin washers under bolt head and nut.
Connect lever rod-end to bellcrank (7) in the same
manner. Install cotter pins when rigging is complete.
f.
(1)
Check the power lever control linkage is
completely installed except as follows:
(a)
Idle stop solenoid (19, figure 420) should not make contact with stop cam (16).
(b)
Arm (27) should be installed in
fuel control power lever shaft (28) as nearly parallel to
shaft stop arm as serration alignment permits. Control rod
(26) should not be connected to arm (27).
If bracket (19) P/N 204-060-797-1 is installed,
use shims P/N 120-031-12-7 to adjust solenoid.
If bracket P/N 204-060-797- 5 is installed,
shims are not required.
NOTE
Before starting rigging procedure, make sure
rod (6) and rod (26) are at nominal length.
(6) Place retainer (22) on control rod (24).
Insert split bushing (23) between rod and retainer and
secure with snapring. Secure boot (21) with clamp on
housing (20). Insert rod down through boot and housing.
Connect rod to lever (25) and bellcrank (15). At lower
rod-end, insert bolt through large safety washer, rod-end,
aluminum alloy washer, and bellcrank. Secure bolt with
aluminum-alloy washer, nut and cotter pin. Attach boot
to retainer with clamp.
NOTE
Check for 0.06 inch minimum clearance
between rod and engine mount leg. If needed,
install not more than three thin steel washers
under spacer on pivot bolt of lever (25).
Rigging.
(2) Center control rod (14) in bellcrank (15)
before trying to obtain center of travel in step (3).
(3) Support free end of control rod (26) level
with fuel control power lever shaft (28). Operate pilot
control grip (5) to full on to full off, and check that end of
control rod (26) moves equal distances from centerline of
shaft (28) at both positions. Adjust control rods (26 and 6)
as close to nominal lengths as possible. Refer to
paragraph 4-22e, step (3) and step (8) for nominal
lengths.
(4)
Position control rod (26) in arm (27). Install
bolt.
(5) Turn pilot control grip (5) in one direction
until fuel control shaft bottoms on stop. Disconnect control
rod (26) from arm and check that control grip (5) will turn
approximately 5 degrees further before bottoming.
Repeat procedure with grip rotated in opposite direction.
Change 38
4-48
TM 55-1520-234-23
Make corrections by adjusting position of control rod (14)
on slotted bellcrank (15). When satisfactory, leave
control rod (26) connected to arm (27). Install cotter pin.
(6) Operate control grip (5) to set power lever
shaft stop arm (27) to 47 degree mark on fuel
control. Adjust positions of idle stop cam (16)
and solenoid (19) so that cam rests against
extended plunger of solenoid. Check that
solenoid bracket clears cam by 0.06 inch in
all conditions (figure 4-21).
Serrations of cam and square washers
must be matched.
(7) In next ground run, make final adjust- ment
of idle stop cam to obtain 68 TO 72 percent gas producer
rpm.
Figure 4-21. Flight idle stop installation
Change 38
4-48A/(4-48B blank)
TM 55-1520-234-23
a. Removal-Actuator and Control Lever.
(1)
Open engine compartment cowling at left
side.
(2) Remove terminal cover with attaching
screws from top of linear actuator (20, figure 4-22).
Disconnect and stow electrical leads. Reinstall cover.
(3) Detach actuator jackshaft end-fittings from
lever (21) on governor control shaft, and from slider of
cambox (19), by removing bolts with nuts, washers and
cotter pins. Use care to avoid losing spring washer,
which is installed between actuator clevis and slider, also
washers installed between rod-end and lever (21).
(4) Remove lockwire and clamping bolt, and
pull lever from serrated shaft at top of overspeed
governor.
b. Removal-Cambox and Linkage.
(1) Disconnect control rod (18, figure 4-22)
from bellcrank of cambox (19) by removing bolt with nut,
washers and cotter pin.
4-23.
Power Turbine Governor RPM Controls.
Engine power turbine speed (N2 rpm) is controlled
through the overspeed governor by means of an actuator
and a droop compensator cam and linkage.
(2) Remove nuts and washers from inboard
ends of two bolts that attach cambox to support bracket.
Remove cambox with bolts in place. Reinstall nuts and
washers on bolts, with care that shims remain in place on
bellcrank pivot bolt between bearing and sides of housing.
NOTE
4-24. Governor Actuator and Droop Compensator
Linkage.
An electrically operated linear actuator (20, figure 422) controlled by the GOV RPM INCR/ DECR switch on
the pilot's collective pitch control stick, moves a lever
(21) on the fuel control overspeed governor to change
sttings of power turbine rpm. Droop compensation, to
stabilize rpm as engine load fluctuates with changes of
main rotor pitch, is provided by mounting the actuator to
a cambox (19) which is mechanically linked to a bellcrank
(2) in the collective pitch control system. The droop
compensator linkage consists of control rods, levers,
arms and bellcranks. Bellcrank (4) is attached on shaft
(7) by means of a shear pin (9), which is designed to
shear to allow unhindered operation of the collective pitch
controls if the compensator linkage should become
fouled.
As an alternate method, remove cambox and
bracket as an assembly by removing two bolts
that secure bracket to forward engine mount
trunnion.
Reinstall bolts to secure mount
trunnion.
(3) To remove bellcrank (4) and shaft (7) and
associated parts: Disconnect control rods (3 and 13).
Remove nut and washer from inboard end of shaft.
Remove three attaching bolts and bracket (11) from
outboard end of shaft. Pull shaft free of inboard bracket
(8). Remove shaft assembly with bellcrank, shear pin
(9), shims (5) and special washer (6). When replacement
of shear pin is required, remove parts from inboard end
of shaft.
(4)
required.
Change 2
4-49
Remove other components of linkage as
TM 55-1520-234-23
Figure 4-22. Power turbine governor rpm controls
Change 48
4-50
TM 55-1520-234-23
f.
(1) Position cambox (19,
figure 4-22) on
outboard side of bracket. Insert two bolts and secure with
nuts and washers. Be sure shims are in place on
bellcrank pivot bolt.
Cleaning solvent is flammable and toxic.
Provide
adequate
ventilation.
Avoid
prolonged breathing of solvent vapors and
contact with skin or eyes.
NOTE
c. Cleaning-Governor
Actuator
and
Droop
Compensator Linkage. Clean external surfaces of parts
by wiping with a cloth moistened with solvent (C124). Do
not permit solvent to enter bearings or actuator.
d. Inspection-Governor
Compensator Linkage.
Actuator
and
Droop
(1) Linear actuator for evidence of damage or
malfunction.
(2) Cambox for security of parts and smooth
operation with no evidence of binding. Check for proper
clearance of 0.001 to 0.006 inch between bellcrank (4)
and flange of shaft (7). Refer to figure 4-22. view A-A.
(3) Check for broken shear pin (9, figure 4- 22)
by manually holding lever (10) and applying slight force to
bellcrank (4).
(4) Other parts of droop compensator linkage
for freedom of operation, looseness and damage.
(5)
Boot (16) for cracks, wear, and security.
(6) Inspect adjustment shaft of single screw for
not more than 4 threads showing beyond the locknut.
e. Repair or Replacement-Governor Actuator and
Droop Compensator.
(1)
Installation-Cambox and Linkage.
Replace actuator if damage or malfunction
If bracket was also removed reinstall on two
upper bolts of engine forward mount
trunnion. Lockwire both heads.
(2) Adjust control rod (18) to nominal length
of 19.0 inches between centers of rod end hearings.
Connect nonadjustable end to forward arm of lever
(17) with bolt, thin aluminum alloy washers, nut,
and cotter pin. Adjustable end will be connected to
bellcrank of cambox (19) during rigging.
(3) If bellcrank (4), shaft (7), and associated
parts are removed, reinstall as follows: Attach bracket (8)
on pylon support with two bolts and washers. Place shims
(5), special washer (6) and thin steel washer on shaft.
Insert shaft through bearing of inboard bracket and secure
with thin steel washer and nut, fingertight. Install bracket
(11) over outboard end of shaft and attach to pylon
support with two bolts and washers. Tighten nut on
inboard end of shaft. Check for 0.001 to 0.006 inch
clearance between bellcrank and shaft as shown. (See
Section A-A.) If necessary, disassemble to change shim
thickness and reassemble.
(4) Adjust control rod (3) to nomi nal length
of 32.46 inches between center of bolt holes in rod
end and clevis. Connect clevis to collective system
bellcrank (2) with bolt, thin aluminum alloy washers,
nut, and cotter pin. Align upper rod end in fork of
bellcrank (4), with a thin steel washer between each
side of bearing and inside of fork. Install bolt,
secured by washer, nut and cotter pin.
occurs.
(2) Replace cambox and bracket assembly if
damaged or failing to operate smoothly.
(3) Replace shear pin (9, figure 4-22) in event
of failure. Investigate cause of failure and correct any
fouling of linkage or other faulty condition.
(5) Assemble retainer, split bushing, snap ring
and boot with clamps on rod (15). Insert rod- end aft
through firewall retainer and connect to upper arm of
lever (17) with bolt, thin aluminum alloy washers, nut
and cotter pin. Connect forward end of rod to inboard
end of arm (14) in the same manner. Secure boot with
clamps on retainers.
(4) Replace other parts in droop compensator
linkage where found unserviceable.
Change 48
4-51
TM 55-1520-234-23
Arm (14) should be installed with angled clevis
outboard and sloping down forward to meet
with control rod (13).
NOTE
If rod (15) is not fixed-length type, adjust to nominal
length of 16.82 inches between center of rod ends
and install adjustable end aft.
(2) Align actuator (20) with front end-fitting
clevis on end of cambox slider. Insert spring washer
between clevis and underside of slider and install bolt
from top, secured with washer and nut. Torque nut 5 TO
15 inch-pounds and insert cotter pin.
(3) Attach actuator shaft rod end with a thin
steel washer on each side of rod end bearing into clevis of
governor control lever with bolt (washer under head).
Install washer and nut, omit cotter pin until rigging is
complete. If necessary, loosen bolts attaching cambox
bracket on engine to align actuator to lever. After
installing actuator, tighten and lockwire bracket bolts.
g. Installation-Actuator and Control Lever.
(1) Place control lever (21, figure 4-22) on
control shaft of fuel control overspeed governor,
approximately 90 degrees to centerline of stop arm
on shaft. Install retaining bolt into lever and through
shaft groove. Torque bolt 12 TO 15 inch- pounds.
Secure bolt to lever to with lockwire (C151).
(4) Remove actuator terminal cover. Connect
electrical leads on terminals. (See wiring diagrams,
Appendix F.) Reinstall terminal cover.
h. Rigging-Power
Turbine
Governor
Controls. (See figure 4-22 and 4-23.)
Figure 4-23. Rigging diagram-governor rpm controls
Change 7
4-52
RPM
TM 55-1520-234-23
cambox bellcrank at this setting.
NOTE
NOTE
Collective pitch control system rigging must be
complete before using this procedure.
Keep control rods as near as possible to
nominal lengths, for safe thread engagement
of rod ends (Refer to paragraph 4-24f.)
The engine should be at flight idle before the
rod end bearing is disconnected from the
governor control lever.
(1) Check that installation of governor control
linkage is complete except:
(5) At overspeed governor control shaft, adjust
upper stop screw to extend 0.21 inch from its mounting
boss. Adjust lower stop screw to extend not less than
0.06 inch from its boss. (See figure 4-23.) Check
installation of lever on governor shaft, to be as nearly 90
degrees to shaft stop arm as serrations permit.
NOTE
(a)
Linear actuator (20, figure 4-22)
disconnected from governor control shaft lever. Support
actuator near normal position, so that its jackshaft rod
end can be moved freely.
(b)
Vertical
control
rod
disconnected from bellcrank of cambox (19).
(18)
(1.1) Set control rod (3) to nominal length of
32.46 inches.
(1.2)
19.0 inches.
Never shorten either-stop screw on
governor to less than 0.06 inch length from
inner side of boss.
Set control rod (18) to nominal length of
(2) Set cam adjustment bolt in middle of slot,
within 0.06 inch. Match serrations of square washer and
cam while tightening nut on bolt. Refer to figure 4-23.
(3) Measure stroke of actuator jackshaft rod
end while operating GOV RPM switch, on collective
control stick, to INCR and DECR. Set actuator adjusting
screw (or screws) to limit stroke to 1.20 inches. After
adjustment,
leave actuator at full INCR (retracted)
position.
(6)
Move collective stick to full up position.
(7) Manually position lever so that shaft stop
arm is 0.010 inch from upper stop screw (figure 4-23).
Adjust actuator jack shaft rod end and connect to lever at
this setting. Install washers on both sides of rod end
bearing between bearing and level (one washer on each
side). Install rod end bolt, washers and nut (one washer
under head of bolt and one washer under nut). Torque 12
TO 15 inch-pounds and install cotter pin. Check that rod
end is centered and torque jam nut 60 TO 85 inch-pounds
and lockwire.
(8) Place collective control stick full down.
Hold GOV RPM switch to DECR until actuator is fully
extended. Adjust lower stop screw on governor to be
0.010 inch away from governor shaft stop arm. Tighten
jam- nuts and lockwire both stop screws s.
NOTE
If actuator has two adjustment screws:
Electrically
position
actuator
shaft
to
approximately midpoint of stroke. Turn both
adjusting screws to obtain maximum stroke.
Reduce stroke by turning each screw equal
number of turns away from maximum
adjustment until rod end travel length is 1.20
inches.
Any adjustments after preliminary rigging will
require recheck and adjustment of governor
stop screws for 0.010 inch clearance of upper
and lower stop bolts according to steps (7)
and (8).
(4) Place collective stick full down. Manually
position cambox bellcrank so that 0.09 ( + 0.03) inch of
cam slot is visible below cambox housing. Adjust two
vertical control rods (3 and 18, figure , 4-22) equally and
in opposite directions, to align and connect upper rod to
(9) After preliminary rigging by preceding
steps, final rigging adjustments will be made as required
in next ground-run or flight:
Change 56
4-53
TM 55-1520-234-23
(a) With collective control stick full down, full throttle,
and governor rpm increase/decrease switch at full
decrease, the rpm should be 6000. At full increase, the
rpm should be 6700. Readjust linear actuator screws to
obtain 6000 + 50 to 6700 ± 50 rpm range.
NOTE
If linear actuator has only one adjustment screw,
adjust screw to obtain a 700 rpm range. Then
adjust linear actuator rod end to obtain 6000 +50
to 6700 +50 rpm range.
(b) Set droop compensator cam to maintain 6600 (±
40) rpm engine output shaft speed from flat pitch to full
power. If rpm droop occurs, move cam adjustment bolt
toward forward (max compensation) end of slot.
If
maximum cam compensation does not correct' rpm droop,
lengthen control rod attached to cambox bellcrank,
increasing amount of cam slot visible below cambox
housing. Cam slot must not bottom out against follower in
either extreme collective stick position.
4-25.
Installation-Engine Electrical Cables.
NOTE
Refer to wiring diagram in Appendix F for
wiring or connector identification.
a. Install nipples over ends of wires of startergenerator cable (6, figure 4-24).
(1)
One nipple on wires K5C4 and K5A1.
(2)
One nipple on wires P26A1 and P26C4.
(3)
One nipple on wires K4B4 and K4D4.
(4)
One nipple on wire P25A16.
b. Remove nuts and washers from terminals C, B, and
E of starter-generator.
NOTE
If terminals of starter-generator are too short,
thin washers may be used.
e. Position main engine cable (10) on engine and
connect electrical connector (3) to airframe main
connector.
f. Remove cover from linear actuator wire terminals
and connect wiring (figure 4-24) as follows:
(1)
Connect wire Q23C18 to terminal R.
(2)
Connect wire Q22C18 to terminal E.
(3)
Connect wire Q26A18N to terminal GND.
g. Secure wiring with bracket and clamp. See Detail A.
h. Connect electrical connector (8) to engine oil
pressure transducer and connect electrical connector (9) to
engine oil pressure switch.
i. Secure cable with clamps as shown in details D and
E.
j. Install one nipple on wire W71D18 and one nipple on
wire W71E18.
k. Remove nuts and washers from terminals of fuel
pressure switches (4). Position one wire on each pressure
switch and reinstall washers and nuts. Place nipples over
terminals.
l. Connect electrical connector (5) to fuel filter bypass
switch and connect electrical connector (7) to engine oil
bypass valve.
m. Secure cable with bracket and clamp. See detail C.
n. Connect electrical connector (2) to bleed air valve
and connect electrical connector (1) to engine oil low level
switch receptacle.
o. Secure harness to bleed air duct with clamps. See
figure 4-25, detail C.
p. Connect electrical connectors (1 and 2) to torque
pressure transducers.
q. Assemble terminal board assembly details (6 thru 11)
to engine support.
Connect engine ground wires to
terminal board 1TB1. Secure wires to terminal board with
washers (5), nuts (4), and cover (3).
c. Position wires K5C4 and K5A1 on terminal E, wires
P26A1 and P26C4 on terminal B, wires K4B4 and K4D4
on terminal C and wire P25A16 on terminal A. Reinstall
washers and nuts. Place nipples over terminals.
s. Install nipple on wire W10C18. Remove nut and
washer from terminal of chip detector (12).
d. Loosely assemble clamps on starter-generator cable.
Clamps are used to secure cable during engine installation.
See Detail B.
t. Install wire W10C18 on chip detector terminal.
Reinstall washer and nut and place nipple over terminal.
Secure wire with clamps. See detail E.
r. Secure cable with clamps. See detail D.
u. Secure remainder of harness to engine using clamps
and brackets. See details A, B, F, G, and H.
v. Secure exhaust thermocouple cable (13) to engine
with clamp and bracket. See detail I.
Change 29
4-54
TM 55-1520-234-23
Figure 4-24. Electrical Cable Installation-Engine Left Side (Sheet 1 of 2).
Change 29
4-55
TM 55-1520-234-23
Figure 424. Electrical Cable Installation-Engine Left Side (Sheet 2 of 2).
Change 29
4-56
TM 55-1520-234-23
Figure 4-25. Electrical Cable Installation-Engine Right Side (Sheet 1 of 2).
Change 29
4-57
TM 55-1520-234-23
Figure 4-25. Electrical Cable Installation - Engine Right Side (Sheet 2 of 2).
Change 29
4-58
TM 55-1520-234-23
CHAPTER 5
ROTOR SYSTEM
5-1.
Rotor System.
The rotor system is comprised of the main rotor system and the tail rotor system.
Section I. MAIN ROTOR SYSTEM
5-2
Main Rotor System.
The main rotor system consists of the main rotor hub
and blade assembly, swashplate assembly, scissors and
sleeve assembly, and connecting tubes (pitch links.) See
figure 5-1.
5-3.
Troubleshooting and Operational Check-Main
Rotor System.
Condition
Minimum Personnel
Required
Consumable Materials
Special Environmental
Condition
Premaintenance Requirements for
Troubleshooting and Operational Check of Main
Rotor System.
Condition
Model
Part No. or Serial No.
Special Tools
Test Equipment
Support Equipment
Requirements
Three Men
(C151)
NA
a. Troubleshooting-Main Rotor System.
Troubleshoot the main rotor system in accordance
with table 5-1 and the specific procedures set forth in
"Operational Checks", in step b.
Requirements
AH-1S
A11
(T44) (T51)
Two fifteen-inch
crescent wrenches
Tracking Flag
Grease pencils for
tracking blades
Table 5-1. Troubleshooting-Main Rotor System
NOTE
Before you use this table, be sure you have performed
all normal operational checks.
CONDITION
TEST OR INSPECTION
CORRECTIVE ACTION
1. Lateral vibration.
STEP 1. Rotor spanwise unbalance.
Balance dynamically with weight in blade bolt. (Refer to paragraph 5-3).
Change 71
5-1
TM 55-1520-234-23
Table 5-1. Troubleshooting-Main Rotor System (Cont)
CONDITION
TEST OR INSPECTION
CORRECTIVE ACTION
STEP 2. Rotor chordwise unbalance.
Balance dynamically by adjusting drag brace (sweeping blade). (Refer to paragraph 5-3.b)
2. Vertical 1/rev vibration.
STEP 1. Rotor blades out of tracks.
Track rotor blades. (Refer to paragraph 5-3.b)
3. Steady or intermittent 1/rev vertical vibration.
STEP 1. Loose collective mast friction assembly.
Check friction sleeve force adjustment. (Refer to paragraph 5-7.g.)
STEP 2. Worn bearings in lever assembly and link.
Replace worn bearings. (Refer to paragraph 11-5.d.)
STEP 3. Worn pitch link rod end bearing.
Replace if wear is excessive. (Refer to paragraph 11-6.d.)
STEP 4. Swashplate pivot ball adjustment incorrect.
Adjust swashplate pivot ball with AVIM assistance. (Refer to paragraph 6-8g(13).
STEP 5. Excessive wear in scissors assembly.
Replace scissors and sleeve assembly. (Refer to paragraph 5-7.e.)
STEP 6. Internal wear or damage in main rotor hub assembly.
Replace hub. (Refer to paragraph 5-6.e.)
NOTE
Wear at one bearing or combined wear at these locations significantly contributes to
vibration.
4. Pylon rock.
STEP 1. Defective fifth mount.
Replace mount. (Refer to paragraph 6-14.)
Change 22
5-2
TM 55-1520-234-23
Table 5-1. Troubleshooting-Main Rotor System (Cont)
CONDITION
TEST OR INSPECTION
CORRECTIVE ACTION
STEP 2. Defective transmission mount dampers.
Repair or replace transmission dampers. (Refer to paragraph 6-16.)
STEP 3. Mount bolts bottomed or stripped.
Replace bolts. (Refer to paragraph 6-16.)
STEP 4. Worn or dirty trunnion bearings.
Inspect, clean, or replace trunnion bearings. (Refer to paragraph 5-4.)
STEP 5. Loose trunnion.
Adjust trunnion, with AVIM assistance.
5. 2/rev vibration, approximately ten per second.
STEP 1. Transmission mounts deteriorated.
Replace transmission mounts. (Refer to paragraph 6-12.d.)
6. High frequency vibration.
STEP 1. Loose elevator linkage at swashplate horn.
Replace worn parts. (Refer to paragraph 2-50.)
STEP 2. Loose elevator.
Reshim bearing. (Refer to paragraph 2-50.)
7. Rotor RPM high or low in autorotation.
STEP 1. Low pitch blade angle incorrect.
Adjust both pitch links equally. (Refer to paragraph 5-36b(7) for B540
blades or paragraph 5-3c(7) for K747 blades.)
8. Slow control response.
STEP 1. Internal leakage in servo cylinder.
Replace cylinder or seals as necessary. (Refer to paragraph 7-11 .h.)
Change 4
5-2A
TM 55-1520-234-23
Table 5-1. Troubleshooting-Main Rotor System (Cont)
CONDITION
TEST OR INSPECTION
CORRECTIVE ACTION
9. 2/1 rev vertical vibration.
STEP 1. Worn or dirty grip/feather bearing.
Inspect. clean. or replace bearing. (Refer to paragraph 5-5.)
10. Unable to get full stroke on the collective during bleed down.
STEP 1. Worn or dirty grip/feather bearing.
Inspect, clean, or replace bearing. (Refer to paragraph 5-5.)
b. Operational Checks-Main Rotor System
(B540 Blades).
NOTE
The strobe-type tracking device may be used if available.
Instructions for use are provided with the device.
Change 29
5-2B
TM 55-1520-234-23
Figure 5-1. Main rotor installation (Sheet 1 of 2)
Change 71
5-2C
TM 55-1520-234-23
Figure 5-1. Main rotor installation (Sheet 2 of 2)
Change 71
5-2D
TM 55-1520-234-23
Figure 5-1A. Main rotor installation torque values (Sheet 1 of 2)
Change 76
5-3
TM 55-1520-234-23
Figure 5-1A. Main rotor installation torque values (Sheet 2 of 2)
Change 71
5-4
TM 55-1520-234-23
Perform following procedures as required for acceptable
smooth operation of main rotor.
Recommended
sequence of procedures is also provided in charts. See
figures 5-2, 5-3 and 5-4.
(1) Coat tracking tips of rotor blades with suitable
grease pencils, different colors on each blade, as
preparation for use of tracking flag. Set both trim tabs at
trail (zero degrees) using gage (T44) and tab bender
(T51).
CAUTION
Run-up of helicopter shall be performed only
by personnel authorized by AR95-1.
NOTE
Designate blades as A and B. Keep a running
record of check results and any adjustments.
Change 74
(2) Perform a low speed blade track at 4700 rpm.
See figure 5-6. If track is satisfactory, omit step (3) and
proceed to step (4).
(3) Prior to accomplishment of MWO 65- 1520-24450-9 correct a low speed out-of-track condition by
shortening pitch link attached to the low blade to roll the
blade up. See figure 5-6. Loosen jamnuts and turn barrel
to shorten tube. Turn barrel ten flats will change blade
track approximately 3/8 inch. Torque 700 inch-pounds
and lockwire (C152) nuts. Repeat checks and adjustments
until satisfactory.
5-4A/(5-4B blank)
TM 55-1520-234-23
(3A) After accomplishment of MWO 551520-244- 50-9 correct a low speed out of track
condition by shortening pitch link attached to the low
blade to roll the blade up. See figure 5-6A. Loosen
jamnuts and turn entire tube assembly to shorten tube.
Turning tube assembly five flats will change blade track
approximately 3/8 inch. Torque jamnut closest to rod
end bearing 1400 TO 1600 inch-pounds and torque
jamnut closest to universal bearing 1100 TO 1300 inchpounds. Lockwire (C-152) jamnuts. Repeat checks and
adjustments until satisfactory.
(4)
Perform a high speed track at 6600
rpm. If out-of-track, record which blade is low but make
no adjustments.
(5)
Test fly helicopter.
If vertical
vibration is not evident, proceed to step (6). If vertical
vibration requires correction, begin sequence of
adjustments indicated on figure 5-3 according to
airspeed where vibration occurs.
If maximum blade sweep adjustment fails to correct
rotor vibrations, remove main rotor hub and blade
assembly and align blades. Refer to paragraph 5-5.
(d) After each adjustment, test fly
helicopter to observe effect. Continue until vertical
vibration is reduced to acceptable level.
(6)
Test fly helicopter through full
airspeed range to check for lateral vibrations. The
lateral vibrations are usually more pronounced in hover.
If lateral vibration is severe enough to require
correction, follow the sequence shown in figure 5-4.
(a) Use two-inch width masking
tape (C134) and apply to blade near the tip. When
satisfactory operation is obtained, remove the tape and
count number of wraps as the tape is removed.
Remove hex-head plug from top of blade retaining bolt
assembly (22, figure 5-1) on taped blade. For each full
wrap of tape, add 2.4 ounces of lead in bolt. Reinstall
plug.
NOTE
After alignment, ensure that all inner threads of the
clevis ends are in contact with those of the drag brace
and the drag brace is no less than flush with the clevis
ends.
(a) When bending blade trim tabs,
do not exceed eight degrees up and/or eight degrees
down (16 degrees maximum both blades).
(b)
To roll a blade: Adjust pitch link
as in steps (3) and (3A) above.
(c) To sweep a blade: Loosen
jamnuts on blade drag brace (11, figure 5-1) enough to
turn barrel one full turn AFT as shown by decal arrows.
Torque jamnuts 150 TO 200 foot-pounds without moving
barrel. Record all such adjustments, and do not exceed
two full turns total adjustment.
NOTE
The blade sweep adjustment is also used to correct
lateral vibration. Refer to step (6).
(b) Refer to paragraph 5-3, b, (5),
(c) for instructions to sweep a blade.
(c)
Refer to paragraph 5-6 for grip
spacing and paragraph 5-5 for alignment instructions.
(7)
Prior to accomplishment of MWO 551520-244-50-9 check rotor rpm in autorotation. If rotor
overspeeds, shorten both pitch links equally. If rotor
underspeeds, lengthen both pitch links equally. One turn
of barrels will change rotor rpm approximately 3 percent.
Torque 700 inch-pounds and lockwire (C152) jamnuts.
Repeat flight check and adjustment as necessary.
(7A)
After incorporation of MWO 551520- 244-50-9 check rotor rpm in autorotation. If rotor
overspeeds, shorten both pitch links equally. If rotor
underspeeds, lengthen both pitch links equally. Onehalf turn of tube assemblies will change rotor rpm
approximately 3 percent. Torque jamnut closest to rod
end bearing 1400 TO 1600 inch-pounds, and jamnut
closest to universal bearing 1100 TO 1300 inch- pounds.
Lockwire (C152)jamnuts. Repeat flight check and
adjustment as necessary.
NOTE
Change 74
5-5
TM 55-1520-234-23
Figure 5-2. Main rotor tracking chart
5-6
TM 55-1520-234-23
Figure 5-3. Vertical vibration correction chart (Sheet 1 of 2)
5-7
TM 55-1520-234-23
Figure 5-3. Vertical vibration correction chart (Sheet 2 of 2)
5-8
TM 55-1520-234-23
Figure 5-4. Lateral vibration correction chart
5-9
TM 55-1520-234-23
Figure 5-5. Tracking main rotor
Figure 5-6. Pitch link adjustment (Prior to accomplishment of MWO 55-1520-244-50-9)
5-10
Change 71
TM 55-1520-234-23
209010-241
Figure 5-6A. Pitch link adjustment (After accomplishment of MWO 55-1520-244-50-9)
Change 71
5-10A
TM 55-1520-234-23
c.
Operational
ChecksMain
Rotor
System(K747 Blades). Perform following procedures as
required for acceptable smooth operation of main rotor.
Recommended sequence of procedures is also provided
in charts (figures 5-2, 5-3, and 5-6B).
WARNING
Runup of helicopter shall be performed only
by personnel authorized by AR 95-1.
K747 main rotor blades have a tendency to
attain a higher percent RPM during
autorotation, than B540 main rotor blades.
DO NOT RIG (adjust length of pitch change
links) beyond the limits established in
paragraph 5-4b(13) to obtain a lower main
rotor percent RPM.
(4)
Perform a high speed track at 100
percent rpm. If out-of-track, record which blade is low
but make no adjustments.
(5)
Test fly helicopter.
If vertical
vibration is not evident, proceed to step (6). If vertical
vibration requires correction, begin sequence of
adjustments indicated in figure 5-6B as applicable
according to airspeed where vibration occurs.
(a)
as in step (3) above.
(b) To sweep a blade, loosen
jamnuts on blade drag brace (11, figure 5-1) enough to
turn barrel one full turn as shown by decal arrows.
Torque jam- nuts 150 TO 200 foot pounds without
moving barrel. Record all such adjustments, and do not
exceed TWO full turns total adjustments.
(1)
Adjust pitch links (figure 5-10) to
accomplish main rotor tracking for K747 blades.
(2)
Perform a low speed blade track at 91
percent rpm. track is satisfactory, omit step (3) and
proceed to step (4).
(3)
Prior to accomplishment of MWO 551520-244-50-9 correct a low speed out-of-track condition by shortening pitch link attachment to the low
blade to roll the blade up (figure 5-10). Loosen jamnuts and turn barrel to shorten tube. Turning barrel
three turns will change blade track approximately 0.375
inch. Tighten and lockwire (C152) nuts. Repeat checks
and adjustments until satisfactory.
(3A) After accomplishment of MWO 551520-244-50-9 correct a low speed out of track condition
by shortening pitch link attached to the low blade to roll
the blade up (figure 5-10). Loosen jamnuts and turn
entire tube assembly to shorten tube. Turning barrel one
and one-half turns will change blade track approximately
0.375 inch. Torque jamnut closest to rod end bearing
1400 TO 1600 inch pounds and torque jamnut closest to
universal bearing 1100 TO 1300 inch-pounds. Lockwire
jamnuts using lockwire (C-152).
5-10B
To roll a blade, adjust pitch link
NOTE
The blade sweep adjustment is also used to
correct lateral vibration. Refer to step (6).
If maximum blade sweep adjustment fails to
correct rotor vibrations, remove main rotor
hub and blade assembly and align blades
(paragraphs 5-4a and 5-5h).
(c) After each adjustment, test fly
helicopter to observe effect. Continue until vertical
vibration is reduced to acceptable level.
NOTE
The amount of lead in blade bolt shall be
changed only by AVIM to statically balance
uninstalled hub without rotor blades.
(6)
Test fly helicopter through full
airspeed range to check for lateral vibrations. If lateral
vibration is severe enough to require correction, request
AVIM assistance to check alignment of installed blades
(sweep).
Change 71
TM 55-1520-234-23
WARNING
K747 main rotor blades have a tendency to
attain a higher percent RPM during
autorotation than B540 main rotor blades.
DO NOT RIG (adjust length of pitch links)
beyond the limits established in paragraph
54b(13) to obtain a lower main rotor percent
RPM.
(7)
Prior to accomplishment of MWO 551520-244-50-9 check rotor rpm in autorotation. If rotor
overspeeds, shorten both pitch links equally. If rotor
underspeeds, lengthen both pitch links equally. One
turn of barrels will change rotor rpm approximately 3
percent. Lockwire (C152) jamnuts. Repeat flight check
and adjustment as necessary.
(7A) After incorporation of MWO 55-1520244-50-9 check rotor rpm in autorotation. If rotor
overspeeds, shorten both pitch links equally. If rotor
underspeeds, lengthen both pitch links equally. Onehalf turn of tube assemblies will change rotor rpm
approximately 3 percent. Lockwire (C152) jamnuts.
Repeat flight check and adjustment as necessary.
Change 71
5-10C
TM 55-1520-234-23
Figure 5-6B. Vertical and lateral vibration chart for K747 rotor blades
5-10D
Change 71
TM 55-1520-234-23
5-4. Main Rotor Hub and Blade Assembly
The main rotor hub and blade assembly consists of the
main rotor hub assembly and the main rotor blades.
See figure 5-1.
Premaintenance Requirements for Main Rotor
Hub and Blade Assembly
Condition
hoist.
(7)
Position two slings (T21) on main
rotor hub and attach to hoist. Do not wrap lifting sling
cable around sharp corners on rotor hub.
(8)
Attach a tie down assembly on each
rotor blade to use to guide and steady rotor during
removal.
Requirements
Model
Part No. or Serial No.
Special Tools
Test Equipment
Support Equipment
Minimum Personnel
Required
Consumable Materials
(9)
Carefully hoist the main rotor hub
and blade assembly clear of the mast and remove cone
set (12, figure 5-1). Attach cone set halves together and
retain as a matched set.
AH-1S
All
(T8) (T13) (T14) (T21)
(T26) (T34) (T49) (T59)
NA
NA
2
(C51) or (C52) (C136)
(C151)
NA
Special Environmental
Condition
a.
Removal.
(1)
Remove bolts (20, figure 5-1) at each
pitch horn (17). Secure pitch links (14) to mast to
prevent damage to these parts.
(2)
Install grip locks (T59) on each pitch
horn. See figure 5-7.
(3)
Remove bolt (5, figure 5-1) and lock
(4).
(4)
Install socket (T13) (8, figure 5-8) on
mast nut. Position adapter (T14) (9) between stabilizer
bar mounts and ensure that it sits level on top of
trunnion. Position power wrench (T8) (3) onto adapter
(9) and ensure that the through pins on the wrench
reaction arm engage the holes in the adapter. Position
the 3/4 inch square drive bar (4) into the square drive of
the power wrench and turn the ratchet indexer (2)
counterclockwise until the drive bar drops into socket
(8).
Install crank handle (1) and turn in a
counterclockwise direction. Observe the indicator on
the power wrench as crank handle is turned. When
breakaway torque of approximately 650 foot-pounds is
reached, the indicator will reverse as the mast nut
loosens.
Remove the special tools and complete
removal of the mast nut manually.
(5)
Remove three bolts (6, figure 5-1),
sand deflector (7), spacer (8) and spacer (9). Remove
opposite sand deflector in the same manner.
(6)
Install maintenance hoist (T49) or
position other suitable hoist directly over mast. Refer to
Chapter 1 for instructions for installation of maintenance
Change 71
(10) Place adapter plate (T34) on build-up
bench (T26) and position hub and blade assembly on
the bench. Place padded supports under each blade.
See figure 5-17 for view of hub mounted on build-up
bench.
(11) Install deflectors that were removed in
step (5).
b.
Installation.
(1)
Install grip locks (T59) on each pitch
horn if not previously accomplished. See figure 5-7.
(2)
Remove sand deflectors if not
previously accomplished. Refer to step (5) in step a
above.
(3)
Position two slings (T21) on main
rotor hub and attach to hoist. Refer to step (a) above for
description of hoist. Do not wrap lifting sling cable
around sharp corners on rotor hub. Hoist hub and blade
assembly into position above mast. Use blade tie down
assemblies to guide and steady the blades during
hoisting.
CAUTION
Never apply corrosion preventive compound or
any kind of grease on or near teflon bearings.
Teflon bearings are used in the hub, the friction
collet, and the swashplate and support
assembly. This instruction applies regardless of
helicopter status; operation, in storage, or in
preparation for overseas shipment.
NOTE
Do not coat the mast threads or split cone
groove with corrosion preventive compound
(C70). Split cones are installed with equal
end gap spacing. If the split cones touch at
any time after maximum torque is reached,
there is no need to respace them.
5-11
TM 55-1520-234-23
Install split cones as matched set only.
(6)
Align master spline in hub with master
spline on mast. Carefully lower the hub and blade on
the mast splines to avoid damage to mast threads.
Lower the hub assembly slowly until it rests on the cone
set.
(7)
Remove excess corrosion preventive
compound with clean cloths.
NOTE
Before installing retaining nut, be sure cap plug
is installed in mast.
(8)
Remove hoisting slings.
Install
washer (2, figure 5-1) and mast nut (3). Tighten mast
nut snug with socket wrench. Install main rotor mast nut
special installation tools in same manner described in
paragraph a. Turn input crank handle (1, figure 5-8) in a
clockwise direction and observe indicator on power
wrench. Torque to 650 foot-pounds. Continue to
observe the indicator for three full minutes. It will be
normal if the indicator reading decreases. This is
caused by the cone set setting. Do not back-off torque if
the indicator reading decreases.
Figure 5-7. Tool application-grip lock
installation on pitch horn
NOTE
Consult TM 55-1500-322-24, pages 13-6 and
13-7, paragraph 13.57 for inspection of teflon
lined bearings.
(4)
Coat splines of mast (13, figure 5-1)
with compound (C52) or compound (C51).
(5)
Inspect split cones for any nicks,
scratches, indentions and deformities of any type.
Dropping the split cone does not constitute automatic
replacement unless the damage limits shown in figure 59 are exceeded. Place cone set (12, figure 5-1) in
groove of mast upper splines with bevel side up and the
gaps evenly spaced.
(9)
If the indicator reading decreases
during the three minute observation period in the
preceding step, re-torque to 650 foot-pounds and
monitor for one minute. Repeat one additional time if
necessary. After obtaining 650 foot-pound indication
with no loss, turn the crank counterclockwise until the
torque indicator returns to zero (green) to remove
holding force on the wrench.
(10) Remove input handle, power wrench,
drive bar, adapter and socket. Check cone set (12,
figure 5-1) for even gap. Check lock (4) installation to
ensure that it will align. If lock (4) will not align, reinstall
the power wrench and increase torque, but do not
exceed 780 foot-pounds. Use the protractor on the face
of the power wrench (3, figure 5-8) to estimate the
degree of turn required to obtain alignment.
CAUTION
ROTOR HUB MUST BE ALIGNED CAREFULLY TO
AVOID DAMAGING MAST THREADS.
Lockwire
NOTE
5-12
Change 71
(11) Install lock (4, figure 5-1) and bolt (5).
bolt
head.
TM 55-1520-234-23
(12) Remove locks (T59) from pitch horns.
Remove eye bolts shown on figure 5-7 and replace with
trunnion bearing retaining bolts. Torque nuts 160 TO
190 inch-pounds. If more than five threads show at nut,
add a washer under the nut.
244-50-9 torque nuts (26) 800 TO 1000 inch- pounds
and install cotter pin (27). After accomplishment of
MWO 55-1520-244-50-9 torque nuts (26) 1250 TO 1550
inch-pounds and install cotter pin (27). Repeat for other
pitch link.
NOTE
CAUTION
Close-tolerance, high-tensile bolts and special
washers are used in the main rotor flight control
linkage. Refer to TM 55-1520-234-23P for part
numbers.
NOTE
If same rotor and associated parts are being
reinstalled, the pitch links (14, figure 5-1) should
already be installed on the scissors and it will
not be necessary to adjust pitch links to nominal
length. Skip step (a) and (b) and install pitch
links per steps (c) through (f).
(13) Install pitch links (14). If previously
removed, install universal bearing (23) on pitch link (14)
with bolt (24) installed through clevis of pitch link (14)
and universal bearing (23). Then install washer (25) and
nut (26). Prior to accomplishment of MWO 55-1520-
Use only new unused nut referenced for
installation in step (13).
(a)
Measure both pitch links and if
necessary adjust length as shown on figures 5-6 and 56A. Also comply with the equal thread requirement
shown on these figures. Prior to accomplishment of
MWO 55-1520-244-50-9 tighten jamnut nuts at each
end of barrel snug but do not torque at this time. After
accomplishment of MWO 55-1520-244-50-9 tighten
jamnuts at each end of tube assembly finger tight.
NOTE
The adhesive bond can be broken when the
adjustment barrel jam nuts are loosened and
torqued without using a back-up wrench on the
adjustment barrel wrench flats.
Change 71 5-12A/(5-12B blank)
TM 55-1520-234-23
Figure 5-8. Tool application - main rotor mast nut removal/installation
Change 2
5-13
TM 55-1520-234-23
Figure 5-9. Corrosion and damage limits-split cones
mast). The lower bolt (28) will be installed with
the bolt head toward opposite scissors link.
CAUTION
Do not reuse nuts P/N MS17825.
WARNING
(b) Before installing pitch change link check
clevis end of scissor (16) in bushing bore. If recess
between bushing end and clevis outer surface exceeds
0.004 inch a shim is required, refer to paragraph 5-7.e.
(6)(g). Prior to accomplishment of MWO 55-1520-24450-9, install pitch link (14, figure 5-1) on scissors (16) if
not previously accomplished. The universal end of the
connecting tubes (pitch links) is the end that attaches to
the pitch links. Install recessed washer (29) on shear
bolt (28) and install shear bolt. Install recessed washer
(30) and nut (31). Repeat procedure to install opposite
pitch link (14). Torque nut of upper bearing bolt 1250 TO
1550 inch-pounds. Torque nuts on shear bolt (28) 800
TO 1000 inch-pounds and install cotter pins.
NOTE
During installation of pitch link universal
bearing, assure that the bolts are installed and
torqued correctly. The upper bolt (24) will be
installed with the bolt head inboard (towards
5-14
Ensure that the self locking feature of barrel nut
(18, fig.
5-1) has adequate tare torque,
minimum of 32 inch-pounds.
Ensure that the correct recessed washer (19) is
being used and that the recess is turned toward
the bolt head.
Ensure that the safety wire is installed correctly.
Bushing assembly, P/N 209-010-523-107 is
required in modified pitch horn assembly, P/N
209-010-109-109 to properly install old pitch link
assembly, P/N 209-010-523-107 is not used in
modified pitch horn assembly, P/N 209-010-109109 to install new pitch link assembly, P/N 209010-520-103.
Change 71
TM 55-1520-234-23
NOTE
Bolt (20) may not extend past barrel nut (18) the
necessary minimum three threads. As long as
the end of the bolt is visible and flush with the
aft side of barrel nut and proper torque has been
accomplished, this is an acceptable condition.
(c) Prior to accomplishment of
MWO 55-1520-244-50-9, install barrel nut (18, figure 51) if not previously accomplished, in pitch horn (17).
Prior to installing pitch link insert bolt (20), check and
record tare torque (friction torque). Tare torque must be
a minimum of 32 inch-pounds. If tare torque is below
limits replace barrel nut (18) and recheck.
(
d)
Prior to accomplishment of
MWO 55-1520-244-50-9, remove bolt (20) and install
barrel end of pitch link (14) to pitch horn (17). Install
recessed washer (19) with recess toward bolt head) and
install bolt (20). Torque bolt (20) 1250 TO 1550 inchpounds above the tare torque previously recorded. For
example: If tare torque was 100 inch pounds you would
torque the bolt 1350 TO 1650 inch-pounds. Lockwire
(C151) bolt (20) to hole in pitch horn (17). Repeat the
procedure for the opposite side.
WARNING
Ensure that the self locking feature of barrel nut
(18) has adequate tare torque, minimum of 32
inch-pounds.
Ensure that the correct recessed washer (19) is
being used and that the recess is turned toward
the bolt head.
Ensure that the lockwire is installed correctly.
NOTE
Bolt (20) may not extend past barrel nut (18) the
necessary minimum three threads. As long as
the end of the bolt is visible and flush with the
aft side of barrel nut and proper torque has been
accomplished, this is an acceptable condition.
(e) After accomplishment of MWO
55-1520-244-50-9, install barrel nut (18) if not previously
accomplished, in pitch horn (17). Prior to installing pitch
link insert bolt (20), check and record tare torque
Change 71
(friction torque). Tare torque must be a minimum of 32
inch-pounds. If tare torque is below limits replace barrel
nut (18) and recheck.
CAUTION
Assure tangs on bushing assembly (19A, Figure
5-1), are engaged into slots on upper rod end of
pitch link (14).
(f)
After accomplishment of MWO
55-1520-244-50-9, remove bolt (20) and install barrel
end of pitch link (14) to pitch horn (17). Install bushing
assembly (19A) on pitch horn (17) and install bolt (20).
Torque bolt (20) 1250 TO 1550 inch pounds above the
tare torque previously recorded. For example: If tare
torque was 100 inch-pounds you would torque the bolt
1350 TO 1650 inch-pounds. Lockwire (C151) bolt (20)
to hole in bushing (19A). Repeat the procedure to install
opposite pitch link (14).
CAUTION
Do not reuse nut MS17825.
NOTE
During installation of pitch links (14) assure bolts
(24 and 28) are installed and torqued correctly.
The upper bolt (24) will be installed with the bolt
head inboard (towards mast). The lower bolt
(28) will be installed with the bolt head toward
opposite scissors link.
(g) After accomplishment of MWO
55-1520-244-50-9, install pitch link (14) on scissors and
sleeve assembly (16) if not previously accomplished.
The universal bearing end of the connecting tube (pitch
link) is the end that attaches to the pitch link. Install
recessed washer (29) on shear bolt (28) and install
shear bolt. Install recessed washer (30) and nut (31).
Torque nuts on shear bolts (28) 800 TO 1000 inchpounds and install cotter pins.
(h) B540. Adjust length of pitch
links to set main rotor blades to a minimum pitch angle
of 8 1/4 degrees (plus or minus 1/4 degree) as follows:
5-14A
TM 55-1520-234-23
1
Position collective controls to
full down position and center cyclic stick.
Lockwire (C152) upper jamnut to barrel. Lockwire
(C152) barrel and lower jamnut to pitch link tube.
2
Place a protractor on machined
surface of one blade grip near blade retention bolt and
measure pitch angle. Record reading and repeat for
opposite blade. The total reading for both blades should
be 17 degrees (± 1/2 degree). If pitch angle is not within
limits, adjust both pitch links in same direction and in
equal amounts until blade pitch angle is within limits.
WARNING
WARNING
Additional pitch link adjustment may be required
at time of maintenance test flight. It is not
necessary to maintain ex- posed threads equal
within 1/6 inch after initial adjustment, but
threads must show in barrel inspection holes.
3
Prior to accomplishment of
MWO 55-1520-244-50-9 check rod end bearings on both
pitch links to ensure that both are centered (figures 5-6
and 5-10). Adjust upper rod end bearing to obtain
alignment if necessary. After alignment is correct,
torque both jamnuts on barrel 700 inch-pounds.
5-14B
After accomplishment of MWO 55-1520- 24450-9 additional pitch link adjustment may be
required at time of maintenance test flight. It is
not necessary to maintain exposed threads
equal within 0.12 inch after initial adjustment,
however exposed threads shall not exceed 1.00
inch per end.
3A After accomplishment of MWO
55-1520-244-50-9 check rod end bearings on both pitch
links to ensure that both are centered (figures 5-6A and
5-10). Adjust upper rod end bearing to obtain alignment
if necessary. After alignment is correct, torque upper
jamnut 1400 TO 1600 inch-pounds and lower jamnut
1100 TO 1300 inch-pounds. Lockwire (C152)jamnuts to
tube assembly.
Change 71
TM 55-1520-234-23
NOTE: View A applicable after accomplishment of MWO 55-1520-244-50-9.
Figure 5-10. Pitch link assembly
4
Lubricate lower bearing on pitch
links (14, figure 5-1). (See figure 1-2.)
5
Ensure that sand deflectors (7,
figure 5-1) have the same part numbers, including dash
numbers, and install with spacers (8 and 9). If spacers
are not a snug fit in sand deflector, wrap with tape
(C135).
6
Perform maintenance test flight
to ensure that main rotor rigging is satisfactory.
(i)
K747. Adjust length of pitch links to
set main rotor hub grips to a minimum pitch angle of 9
3/4 degrees (± 1/2 degree) as follows:
1
Position collective controls to
full down position. Set cyclic to center position.
2
Place a protractor on machined
surface of one blade grip near blade retention bolt and
measure angle. Record reading and repeat for opposite
grip.
Change 71
The total reading for both blades should be 19 1/2
degrees (± 1 degree). If angle is not within limits, adjust
both pitch links in same direction and in equal amounts
until angle is within limits.
WARNING
Additional pitch link adjustment may be required
at time of maintenance test flight. It is not
necessary to maintain exposed threads equal
within 0.060 inch after initial adjustment, but
threads shall show in barrel inspection holes.
3
Prior to accomplishment of
MWO 55-1520-244-50-9 check rod end bearings on both
pitch links to ensure that both are centered (figures 5-6
and 5-10). Adjust upper rod end bearing to obtain
alignment if necessary. After alignment is correct,
torque both jamnuts on barrel 700 inch-pounds. Lockwire (C152) upper jamnut to barrel. Lockwire (C152)
barrel and lower jamnut to pitch link tube.
5-15
TM 55-1520-234-23
WARNING
After accomplishment of MWO 55-1520-244-50-9
additional pitch link adjustment may be required at time
of maintenance test flight. It is not necessary to
maintain exposed threads equal within 0.12 inch after
initial adjustment, however exposed threads shall not
exceed 1.00 inch per end.
CAUTION
Assure tangs on bushing assembly (19A, Figure 5-1) are
engaged into slots on upper rod end of pitch link (14).
3A
After accomplishment of MWO
55-1520-244-50-9 check rod end bearings of both pitch
links to ensure that both are centered (figures 5-6A and
5-10). Adjust upper rod end bearing to obtain alignment
if necessary. After alignment is correct, torque upper
jamnut 1400 TO 1600 inch- pounds and lower jamnut
1100 TO 1300 inch- pounds. Lockwire (C152) jamnuts
to tube assembly.
4
Lubricate lower bearing on pitch
links (14, figure 5-1) with grease (C70).
WARNING
K747 main rotor blades have a tendency to attain a
higher percent RPM during autorotation than B540 main
rotor blades.
DO NOT RIG beyond the limits
established in paragraph 5-4b (13) to obtain a lower
main rotor percent RPM.
5
Perform maintenance test flight
to ensure that main rotor rigging is satisfactory.
6
If the maintenance test flight indicates the need for rotor adjustment, recheck blade
alignment with blades on helicopter. Using align- ment
scope (T30), make required adjustments. (Maximum
tolerance of alignment between two blades is 0.050
inch.)
5-5. Main Rotor Blade Assembly.
Each blade is attached in the hub with a retaining bolt
assembly and is held in alignment by adjustable drag
braces.
Premaintenance requirements for main rotor blades.
Condition
Model
AH-1S
Part No. or Serial No.
All
Special Tools
(T26) (T31) (T34)
(T44) (T59)
Test Equipment
NA
Support Equipment
Work aid for removal of
blade retaining bolt
Minimum Personnel
Four
Consumable
(C3) (C7) (C11) (C12)
Materials
(C17) (C23) (C24) (C30)
(C37) (C39) (C41) (C44)
(C67) (C76) (C77) (C87)
(C88) (C91) (C10OO0)
(C102) (C112) (C113)
(C124) (C127) (C136)
(C134A) (C149)
Special Environmental
NA
Conditions
a.
Removal.
(1)
Position main rotor hub and blade
assembly on build-up bench. Refer to paragraph 5-4.
Place padded supports under blades so that leading
edge is approximately straight.
(2)
Remove nut(11, figure 5-11) and bolt
(14). Loosen nut (21) and swing drag brace (15) away
from rotor blade. Retain shims (10) for reinstallation.
(3)
Remove locking screw (20, figure 5-
11).
(4)
Remove nut (17, figure 5-11) with
socket wrench (T31).
.
The main rotor blades are metal, bonded assemblies.
5-16
Requirements
Change 71
TM 55-1520-234-23
CAUTION
Avoid blade contact with the drag brace during
removal procedure to prevent possible blade
damage.
(5)
Remove blade retaining bolt (8, figure
5-11). Slowly raise and lower blade tip while tapping
bolt with fiber mallet. If bolt is difficult to remove, use a
bolt removal work aid similar to the one shown on figure
5-12. Remove blade retaining bolt as follows.
(a) Remove threaded plugs from
upper and lower ends of blade retaining bolt. If weights
are present in bolt, retain for reinstallation.
Change 71
5-16A
TM 55-1520-234-23
Figure 5-11. Main rotor hub and blade assembly
5-16B
Change 71
TM 55-1520-234-23
Figure 5-12. Work aid for main rotor blade bolt removal - fabrication instructions (AVIM)
5-17
TM 55-1520-234-23
adequate ventilation. Avoid prolonged breathing of
solvent vapors and contact with skin or eyes.
(2)
Remove stubborn deposits with a
cloth dampened with solvent (C124).
c.
Inspection.
(1)
Inspect blade historical records and
the blade for evidence that the blade has been
subjected to an accident or incident outside the realm of
normal usage. If such evidence exists, perform Special
Inspections outlined in Chapter 1.
(2)
Inspect blade for nick, scratch, dent
and erosion damage. See figure 5-14.
(a) Nicks and scratches anywhere
on the surface of the skins or trailing edge strip that do
not exceed 0.008 inch in depth are acceptable if they
are polished out.
Figure 5-13. Work aid application - removal of main
rotor blade retaining bolt
(b) Position work aid on bolt as
shown on figure 5-13 and also place a piece of hard
rubber or similar material between work aid tube and
grip to prevent marring the grip. Hold puller rod (1,
figure 5-12) and tighten nut (2) to remove blade
retaining bolt.
(c) Remove work aid from blade
retaining bolt. Reinstall weight and plugs in blade
retaining bolt and identify the blade retaining bolt for
reinstallation in the same grip.
(6)
Remove blade from grip and place in
a padded stand.
(7)
same manner.
b.
Remove opposite blade from hub in
Cleaning.
(1)
Clean main rotor blade with cleaning
compound (C41).
NOTE
If a nick or scratch in the skin in excess of 0.008 inch
depth can be polished smooth without leaving the skin in
the polished area so thin that skin can be dented with
fingernail pressure, a patch may be applied over the
area. Refer to paragraph e. for instructions to apply
patch to this type damage.
(b) Dents in the trailing edge strip
that are between 0.020 and 0.040 inch deep are
reparable by working the metal with a plastic mallet.
Nicks and notches in the extreme trailing edge of the
blade that are between 0.040 and 0.120 inch in depth
are acceptable if they are polished and faired out over a
minimum distance of 2 inches on each side of the nick
or notch.
(c)
Any dent in the skin in the
outboard four feet of the blade that does not tear the
skin, produce a void detectable by tapping with a coin,
or affect flight characteristics, is acceptable without
repair.
(d)
Dents in the skin inboard of the
station located four feet inboard of the tip of the blade
that do not exceed 0.060 inch are acceptable without
repair.
Cleaning solvent is flammable and toxic. Provide
5-18
Change 7
TM 55-1520-234-23
Figure 5-14. Main rotor blade
(e)
If a nick or scratch exists in a
sharp dent in the skin, the total depth of both must not
exceed 0.060 inch if scratch or nick does not exceed
0.010 inch. Nicks and scratches must be polished out.
Refer to step (a).
(f)
Nicks or scratches in the
abrasive strips, doublers, grip plates or drag plates that
are not in excess of 0.012 inch in depth are acceptable
if they are polished out.
(g)
If a leading edge abrasive strip
is worn, eroded or damaged so that any holes appear,
the blade must be sent to an overhaul facility for repair.
NOTE
Main rotor blade abrasive strip splice joints (16 and 18,
figure 5-14) may have no covers, may be covered with
polyurethane tape, or may have splice covers (15 and
17 installed).
(h)
edge stations 83 and 216 for loss of filler material and
corrosion. Filler material will be replaced at depot level
of maintenance.
(i)
Application
of
scarf
joint
protection tape. Polyurethane tape bonded in place
over the scarf joint is recommended for dusty (sandy)
environments as protection for the scarf joint filler
material. Apply as follows:
1
Clean
aliphatic in aphtha (C88).
2
seven inches long.
edge
with
Cut a piece of tape (C154)
3
Center the piece of tape over
the scarf joint. Press in place.
4
required.
Inspect scarf joints at leading
Change 38
leading
5-19
Replace tape when worn or as
TM 55-1520-234-23
(j )
If no covers are installed on
splice joints (16 and 18), inspect for los10 of filler
material and for corrosion. Replace filler material at
next higher level of maintenance.
(3)
Inspect blade for void damage.
NOTE
A void is defined as an unbonded area that is
supposed to be bonded. Many sub-definitions
of voids are often given such as lack of
adhesive, gas pocket, misfit, etc. This manual
makes no distinction among these, but groups
them in the one general term of "Void." All
dimensions are in inches.
(a) Voids
between
the
spar
assembly and the adhesive strip outboard of station
100:00.
1
1.0 inch wide maximum void
between abrasive strip and spar at extreme leading
edge is acceptable, to within 1.0 inch of the tip of the
blades.
2
Voids not exceeding 30 square
inches with a maximum of 10 square inches in any
single void are acceptable. If voids come closer than
1.0 inch to each other consider them a single void.
3
Voids within 0.38 inch of edge
of abrasive strip are not acceptable.
(b) Voids
between
the
spar
assembly and the abrasive strip inboard of station
100.00.
1
1.0 inch wide maximum void
between abrasive strip and spar is acceptable. Refer to
step 4.
2
Voids between abrasive strip
and spar not exceeding 10.0 square inches with a
maximum of 2.0 square inches in any single void are
acceptable. Minimum spacing between void center
must exceed 3.0 inches. Refer to step 4.
3
Voids within 0.38 inch of edge
of the abrasive strip are not acceptable except at the
butt end, per step 1 above. Refer to step 4.
4
Voids defined in steps 1 2, and
3 that are apparent at the butt end of the blade must be
sealed with adhesive. Refer to paragraph d. for
instructions to apply adhesive.
(c)
Voids at butt end of blade.
1 Void between trailing edge extrusion
and skin not deeper than 1.0 inch nor wider than 1.0
inch is acceptable if sealed. Refer to step 4.
2
Any other void not longer than
1.0 inch or deeper than 0.35 inch is acceptable if sealed.
5-20
Refer to step 4.
3
Voids are not acceptable within
0.5 inch of the front or rear edge of either grip plate or
grip pads, viewing the "Section" of the butt end, if
sealed. Refer to step 4.
4
Voids defined in steps 1,2, and
3 must be sealed-with adhesive. Refer to paragraphd.
for instructions to apply adhesive.
(d) Voids in the retention area,
inboard of Station 100.00.
1
Singledge voids of 0.06 inch
maximum depth on-leading edge of doublers and 0.10
inch of trailing edge of doublers are acceptable if
sealed. Edge voids are not acceptable in outboard
seven inches of each finger of the doublers. Edge voids
in the outer three inches of the grip plate and outer 1.5
inches of the drag plate are not acceptable. Up to 0.5
inch maximum may be removed from the outboard tip
of the drag plate tang, grip plate tang, or outboard tip of
doublers to eliminate a void. Refer to paragraph 5-5, d,
(6).
2
Voids between the doublers,
doubler and skin, doubler and grip plate, grip plate and
grip pad are not acceptable, except as allowed in steps
(c) 2, (c) 3 and (d) 1 above.
3
Voids between the skin and
core in a five-inch-wide region running adjacent to
trailing edge extrusion, not wider than 1.0 inch
longer than 10.0 inches with a minimum width of
inch of good bond between them, are permissible.
the remaining area, the width of the void must
exceed 0.5 inch. The total area of all voids must
exceed 30 square inches.
the
the
nor
1.0
In
not
not
4
Edge voids between the edge of
the skin and the trailing edge extrusion, that are less
than 0.06 inches wide by any length or less than 0.25
inches wide by 7 inches long are acceptable if they are
sealed with adhesive (C7), adhesive (C11), adhesive
(C12) or adhesive (C17).
5
Voids within one inch of the
main retention bolt, in any bond line, are not
permissible.
6
Other voids between the skin
and the trailing edge extrusion which do not exceed one
third the width of the faying surfaces by 10.0 inches long
are acceptable.
Change 7
TM 55-1520-234-23
(e)
Voids under skin, outboard of
Station 100.0.
1
Voids between the skin and the
trailing edge extrusion shall not exceed one third the
width of the faying surfaces.
2
Voids between the skin and the
core must not exceed 1.0 inch in width chordwise. If two
voids are within 1.0 inch of each other, consider them as
one void.
3
Voids between the skin and the
spar not wider (chordwise) than 1/3 the width of the
mating surfaces are acceptable. Edge voids are not
acceptable.
4
Edge voids between the edge of
the skin and the trailing edge extrusion that are less than
0.06 inches wide by any length or. less than 0.25 inches
wide by 10.0 inches long are acceptable if they are
sealed with adhesive (C7), adhesive (C11), adhesive
(C12) or adhesive (C17).
Change 65
5-20A/(5-20B blank)
TM 55-1520-234-23
NOTE
Where two voids of two different types are closer than
1.0 inch apart, consider them as one void and apply the
more strict limitations. (Example: Voids between skin
and trailing edge extrusion next to a void between the
skin and the core).
Repairs inboard of Station 210 must be
inspected daily for cracks.
(4)
Inspect blade for worn retention bolt
hole and worn drag brace bolt hole.
(a)
1.0 inch of doublers.
(a)
If wear allowance listed in steps
(b), (c), or (d) is exceeded, send blade to Depot Level
Maintenance for repair.
(b) Inboard of Station 216, only one
repair on the same chordline is permitted.
After
cleanup, holes are limited to 2.0 inches maximum
diameter and are restricted to a minimum of 2.0 inches
between repairs.
(b) Main retention bolt hole
oversize when the diameter exceeds 2.505 inches.
is
(c) Drag plate bolt hole is oversize
when the diameter exceeds 0.877 inches.
(d) Polish out any corrosion or
pilting from either bushing (Figure 5-15, Items 12 or 13).
Inside diameter of bushing will not exceed limits of steps
(b) and (c) above after polishing.
(5)
Inspect blade for cracks.
(a) Visually inspect top and bottom
surfaces along entire length of blade for damage. Any
fatigue crack in any location is cause for blade
scrappage. Evaluate cracks caused by strikes and other
damage to other damage criteria.
(b) Penetration through spar
trailing edge strip is cause for blade replacement.
or
(c) Damage penetrating skin and at
least one inch from doublers may be repaired, provided
that after cleanup damage does not exceed two inches
in diameter.
(d) Spanwise cracks penetrating
skin and at least one inch from doublers may be
repaired, provided that after cleanup, using an oblong
hole, damage does not exceed four inches by one inch
and direction of oblong hole falls within 15 degrees of a
line parallel to leading or trailing edge of blade.
(6)
Inspect blade for holes in skin. If any
holes are found, classify them as reparable or nonreparable by patching in accordance with the following
limits: See figure 5-15.
Change 65
No patches are permitted within
(c) Between Station 216 and
Station 240, two holes are permitted on same chordline
on same skin surface. Maximum diameter of holes is
2.0 inches and a minimum spacing of 2.0 inches after
cleanup is required between repairs.
(d) Between Station 240 and
outboard tip of blade, two holes are permitted on same
chordline on same skin surface. Maximum diameter of
holes is 3.0 inches and a minimum spacing of 2.0 inches
after cleanup is required between repairs.
(e) Spanwise
holes
may
be
repaired providing that after cleanup making an oblong
hole, damage does not exceed 1.0 inch wide and 4.0
inches long. Direction of oblong hole must fall within 15
degrees of a line parallel to leading or trailing edge of
blade. Ends of the hole must have a minimum radius of
0.25 inch to break corners.
(f)
Any damage or defect in the
skin that can be polished smooth without leaving the
skin in the area so thin that it can be dented with
fingernail pressure does not require a cut out. In these
cases a patch must be applied as though a hole exists.
Maximum diameter of a patch of this type is 4.0 inches
with a minimum of 0.75 inch of bonded area around the
perimeter of the dent.
(7)
Inspect main rotor blade for the
following defects. If blades are damaged to the extent
described, condemn and demilitarize locally rather than
returning blade to an overhaul facility.
(a) Any
penetration
damage
through spar or trailing edge strip, doublers, grip plates
or drag plates.
5-21
TM 55-1520-234-23
No patches are permitted within one inch of the doublers. spar. trailing edge strip and the tip of the blade.
This is the shaded area in the illustration above. Refer to notes 1 through 4 and the table to accurately define
the patchable area.
NOTES:
1.
On blades P/N 540-011-001-5, P/N 540-011-250-1, and 540-015-001-1, -3 and -5 the spar tapers
uniformly between blade stations 80.0 and 140.0. The spar width is constant either side of these stations.
2.
On blade P/N 540-011-001-5 the trailing edge strip tapers uniformly between blade stations 61.5 and
150.0. Trailing edge strip width is constant either side of these stations.
3.
On blade P/N 540-011-250-1 and 540-015-001-1, -3 and -5 the trailing edge strip tapers uniformly
between station 95.0 and 220.0. Trailing edge strip width is constant either side of these stations.
4.
All dimensions are in inches.
Chordwise limits of patchable area at various stations for blades P/N 540-011-001-5,
540-011-250-1 and 540-015-001-1, -3 and-5.
BLADE
P/N
BLADE
STATION
AFT OF
LEADING EDGE
FORWARD OF
AFT EDGE
540-011-001-5
540-011-001-5
540-011-001-5
540-011-001-5
80.0
140.0
61.5
150.0
9.76
6.70
---
-4.145
2.245
540-011-250-1
540-011-250-1
540-011-250-1
540-011-250-1
80.0
140.0
95.0
220.0
9.76
6.70
---
--4.145
2.245
540-015-001-1, -3 and -5
540-015-001-1, -3 and -5
540-015-001-1. -3 and -5
540-015-001-1. -3 and -5
80.0
140.0
95.0
220.0
9.76
6.70
--
--4.145
2.245
540011-142-1B
Figure 5-15. Main rotor blade authorized patch area (Sheet 1 of 2) (AVIM)
5-22
TM 55-1520-234-23
Figure 5-15. Main rotor blade authorized patch area (Sheet 2 of 2) (AVIM)
5-23
TM 55-1520-234-23
(b) Skin penetration in any area
larger than 2.0 inches in diameter after cleanup or
oblong penetration larger than 1.0 inch by 4.0 inches
after cleanup.
(c)
inspection procedure.
(e). Inspect studs for looseness and
distortion. Loose or distorted studs are unacceptable
and is cause for blade removal.
Water in honeycomb core.
d.
(d) Voids
between
skin
honeycomb core larger than 30 square inches.
(e) Edge voids deeper than 0.50
inch in tip end of any of doublers or grip plates.
(f)
Edge voids in the leading edge
of the doublers that exceed 0.060 inch in depth and at
the trailing edge of the doublers that exceed 0.10 in
depth.
(g)
entirely through skin.
Any corrosion that penetrates
(h) If one or more cracks develop
and extend from a previously repaired area.
(i)
More than one patch on the
same chordline on the same side.
(j)
Obvious deformation of blade.
(8)
Inspect main rotor blade trim tab for
the following defects:
(a)
Distortion that can be repaired
BY straightening.
(b) Cracks, tears, rips and holes.
This type damage must be repaired by replacement of
the trim tab.
(9)
(1)
Polish out nick and scratch damage in
skin that is within limits stated in inspection paragraph c.
Use 320 grit sandpaper (C112) to polish out damage.
Use fine aluminum wool (C24) or Scotchbrite (C113) to
finish polish the area.
Rub spanwise to remove
sandpaper marks and polish to a finish of 32 RMS.
Touch-up repair area with chemical film (C37), primer
(C102) and paint to match surrounding area. Refer to
paragraph e.
(2)
Polish out nick and scratch damage in
the abrasive strips, doublers, grip plates and drag plates
that is within limits stated in inspection paragraph c.
Use 400 grit sandpaper (C112) or equivalent. Steel
wool (C127) may also be used providing that no
aluminum parts are touched with it. Touch up repair
area with primer and paint to match surrounding area.
(3)
Repair minor dent damage in the
trailing edge strip that is within limits stated in inspection
paragraph c. Use a plastic mallet to work the metal
slightly.
(4)
Repair nick, scratch, and notch
damage in the trailing edge strip that is within limits
stated in inspection paragraph c. Use varying grades of
sandpaper to polish out damage and fair out over a
distance of 2.0 inch minimum on each side of damage.
Touch up repair area with primer and paint to match
surrounding area.
(5)
(AVIM) Repair hole damage and nick
or scratch damage in skin that is within limits stated in
inspection paragraph c as follows:
Tip balance weight inspection.
(a)
Repair.
and
Remove tip cap from rotor
blade.
for distortion.
removal.
(b)
Visually inspect lead weights
Distorted weights are cause for blade
(c) Inspect for loose weights.
Loose weights alone are not cause for blade removal.
Inspect stud retention nuts for looseness by applying 30
inch pound torque. Torque loose stud retention nuts to
130 to 145 inch pound.
(d) Striking of rotor blades to check
for loose balance weights is not an acceptable
5-24
Provide adequate ventilation when using methyl-ethylketone (C87).
Avoid breathing vapors and avoid
prolonged contact with skin.
(a) Remove paint from repair area with methylethyl-ketone (C87). Dry with a clean cloth. Do not allow
cleaner to enter the blade.
Change 41
TM 55-1520-234-23
NOTE
If a nick or scratch in the skin in excess of 0.008
inch depth can be polished smooth with- out
leaving the skin in the polished area so thin that
skin can be dented with fingernail pressure,
apply a patch over the area without cutting a
hole. Comply with step (a). Skip steps (b) and
(c) and proceed with step (d).
(b) Draw a circle around the
damaged area just large enough to encompass damage.
(c) Remove skin just inside the
circled area, disturbing the honeycomb core as little as
possible. Heat the cut out disk to 200°F maximum and
lift out the disk of skin while heated.
(d) Deburr edges of hole and polish
out scratches and nicks.
Change 41
5-24A/(5-24B blank)
TM 55-1520-234-23
(e) Prepare a patch to cover the hole that will
overlap by 0.75 inch. Fabricate patch from 2024 T3
aluminum 0.020 inch thick (item 4, table 2-1) large
enough to overlap the hole at least 0.75 inch all around
the perimeter. Deburr and blend out edges. Sand the
bond area of the patch and blade with 400 grit paper
(C112).
(7)
Repair distorted main rotor blade trim
tab that is within limits stated in inspection paragraph c.
as follows:
(a) Straighten the trailing edge of
the main rotor trim tab with a mallet and a heavy backup block.
(b) Set trim tab to trail with tab
bending tool (T51) and tab bending gage (T44).
(8)
(AVIM) Remove damaged main rotor
blade trim tab and install new trim tab if replacement is
indicated by inspection paragraph c.
Provide adequate ventilation when using
methyl-ethyl-ketone (C87).
Avoid breathing
vapors and avoid prolonged contact with skin.
(a) Cut through the trim tab (8,
figure 5-15) at a line approximately one-eighth inch aft,
and parallel to blade trailing edge.
(f)
Clean bond area on patch and
blade with methyl-ethyl-ketone (C87). Dry with a clean
cloth.
(b) Drill out all rivets attaching trim
tabs to rotor blade, if existing.
(c) Apply heat to tab with a heat
gun, but do not exceed 200°F. Start at outer corner of
trim tab and peel tab off blade in spanwise direction.
Area must be clean, dry and free of grease, oil
and wax.
(d) Mask blade area around trim
tab; allow one-half inch border from trim edge for
squeeze-out.
(g) Apply adhesive (C12) or
adhesive (C17) to patch area around hole and to patch.
Apply patch to blade and move slightly under pressure
to expel air and prevent voids in bond. Blend out
excess adhesive.
(e) Remove old adhesive in
masked area by sanding spanwise with 180 grit
sandpaper (C111). Use progressively finer sandpaper
320 and 400 grit to obtain a smooth finish.
(h) Hold patch in place with rubber
bands, made from inner tube, or other mechanical
means while curing. Allow adhesive (C12) to cure at 70
to 90 degrees F for 24 hours or 145 to 155 degrees for
30 minutes until completely firm. Allow adhesive (C-17)
to cure at 75°F minimum for 5 days, or at 180°F for 60
minutes.
Provide adequate ventilation when using
methyl-ethyl-ketone (C87).
Avoid breathing
solvent vapors and avoid prolonged contact with
skin.
(f)
Clean trim tab area of blade
with cloths moistened with methyl-ethyl-ketone (C87).
Dry the area with dry, clean cloth.
(g)
Fill rivet holes in trim tab area
of rotor blade, if existing, with adhesive (C12).
(h) Drill nine 0.129-0.132 inch
diameter holes in new trim tab as illustrated on figure 516. If the previous trim tab was riveted, reverse top and
bottom hole locations in trim tab.
(6)
Repair voids up to a maximum of
0.50 inch from tips of drag plate tang, grip plate tang, or
outboard tip of doublers as follows:
(a) Cut material from doubler,
maximum of 0.50 inch, following the same radius as
original tip. Use extreme care to avoid cutting into
adjacent parts.
(b)
After cutting, debur and break
sharp edges.
(c) Refinish
procedures in paragraph 5-5, e.
in
accordance
with
5-25
TM 55-1520-234-23
Figure 5-16. Main rotor blade trim tab installation
(i)
Position trim tab on rotor blade in the install
position and, using holes in trim tab as template, drill
corresponding 0.129-0.132 inch diameter holes to a
maximum depth 0.125 inch in rotor blade.
methylethyl-ketone (C87); then dry surfaces with clean,
dry cloth.
(m) Spread a thin film of adhesive
(C12) to mating areas of rotor blade and trim tab.
(])
Remove trim tab from rotor
blade, sand and smooth areas around drilled holes in
trim tab
and blade.
(n) Position and secure trim tab in
install position on rotor blade, with holes in trim tab
aligned with corresponding holes in rotor blade.
(k) Sand inside mating sides of trim
tab with 200 grit sandpaper (C112) and finish with 400
grit sandpaper (C112).
(o) Install a rivet (CR2263-4-1) in
each of nine holes drilled in trim tab and blade; dip each
rivet in adhesive (C12) before installation.
Provide adequate ventilation when using
methyl-ethyl-ketone (C87).
Avoid breathing
solvent vapors and avoid
prolonged contact with skin.
(p) Use
two
wooden
blocks
approximately the same size as trim tab and two
sections of hard rubber one sixteenth inch thick and
approximately the same size as the wooden blocks to
use as pressure pads. Place the rubber sections next to
the trim tab bend area and place the wooden blocks
over the rubber sections, apply two to ten PSI pressure
on the trim tab bend area and maintain for a minimum
of 24 hours at 70°°to 90°F.
blade
and
trim
(I)
Clean mating surfaces of rotor
tab with cloth, moistened with
5-26
Change 48
TM 55-1520-234-23
NOTE
Curing time may be accelerated by application
of heat of 145 TO 155 degrees F (63 TO 69
degrees C) for approximately 30 minutes, or
until squeeze-out resists fingernail penetration.
(q) Remove pressure pads after
curing time and smooth squeeze-out with 180, 320, and
400 grit sandpaper (C112).
Provide adequate ventilation when using
methyl-ethyl-ketone. Avoid breathing solvent
vapors and avoid prolonged skin contact.
(r)
Clean up adhesive squeeze out
in trim tab area with methyl-ethyl-ketone (C87) and dry
with clean, dry cloth.
(s) Apply chromic
(C3) and dry with clean, dry cloth.
acid
solution
(t)
Apply one coat of primer (C100)
to trim tab, and adjacent blade area and allow to dry for
period of 30 minutes to four hours.
Provide adequate ventilation when using
methyl-ethyl-ketone. Avoid breathing solvent
vapors and avoid prolonged contact with skin.
1
Use sharp plastic scrapers and
small amount of methyl-ethyl-keytone (C87) to remove
the tape.
2
Remove residual parts of tape
with 100 grit or finer sandpaper (C112). Sand in
spanwise direction.
Provide adequate ventilation when using
naphtha. Avoid breathing solvent vapors and
avoid prolonged contact with skin.
3
Clean leading edge of rotor
blade, in area where tape will be applied, with clean
cloths dampened with aliphatic naphtha (C88).
(b) Cut a piece of polyurethane
tape (C134.1) seven inches long.
(u) Apply lacquer (C76) to trim tab
and adjacent blade area.
(c) Center
polyurethane
tape
chordwise over the splice joint, press into place and
force out all air bubbles. If necessary, make pin hole in
tape to allow trapped air to escape.
NOTE
Adhesion difficulty will be encountered if acrylic
lacquer is not applied within a four hour period.
(10) Remove damaged splice cover (15 or
17, figure 5-14) and install new splice cover if
replacement is indicated by inspection paragraph c.
(9)
Remove damaged polyurethane tape
at splice joints (16 and 18, figure 5-14) and install new
tape if replacement is indicated by inspection paragraph
c.
NOTE
Polyurethane tape bonded in place over the
scarfjoint is recommended for dusty (sandy)
environments as protection for the scarf joint
filler material.
(a) Remove old polyurethane tape
from rotor blade leading edge.
Change 7
Do not exceed 200 degrees F (93 degrees C)
during splice cover removal or damage to rotor
blade may result.
(a) Heat splice cover to 200
degrees F (93 degrees C) maximum with a heat gun.
Maintain temperature during removal procedure.
(b) Carefully remove old splice
cover with a putty knife or chisel. Do not damage rotor
blade.
5-27
TM 55-1520-234-23
(c) After splice cover has been
removed, clean old adhesive, paint, and other
contaminants from the abrasive strips in the area where
the splice cover will be installed. Use 100 grit or finer
sandpaper (C 112). Sand in spanwise direction.
spatula and rub adhesive around on the splice cover
and the blade to assure complete "wetting" of the mating
surfaces.
(12) Bond new splice cover to rotor blac as
follows:
(11) Prepare a new splice cover P/N 204015-011-1 and the rotor blade for installation of the
splice cover as follows:
(a) Wear clean, dry gloves when
handling parts that have been prepared for bonding.
Avoid contaminating parts with oil, grease, or mold.
(b) Mask the area of the rotor blade
around the splice cover installation area. Leave a one
inch border between the splice cover and the masking
tape. Use tuck tape (C134.2).
(a) Position splice cover (15 or 17,
figure 5-14) as applicable on rotor blade with the inboard
end of the cover 0.650 TO 0.850 inch inboard of the
most inboard end of the splice joint as illustrated.
(b) Move splice cover back and
forth slightly to expel air pockets.
(c)
Wipe off excess adhesive and
fair in adhesive at edges of cover.
(d) Place a sheet of peel ply (C95.1)
over splice cover to prevent adhesion to bands to be
installed in the following step.
(e) Hold splice cover in place with
heavy rubber bands or bungee cords.
Provide adequate ventilation when using
methyl-ethyl-ketone. Avoid breathing vapors
and avoid prolonged contact with skin.
(c) Clean the area of the rotor
blade where the splice cover will be installed. Use 100
grit or finer sandpaper (C112). Sand in spanwise
direction and remove old adhesive, paint, and other
contaminants.
Clean the area with clean cloths
dampened with methyl-ethyl-keytone (C87) and wipe dry
with clean cloths.
(d) Remove peel ply from the
inside surface of the splice cover.
(e) Lightly sand the cured adhesive
on the splice cover with 300 grit sandpaper (C112).
Wipe with clean cloth to remove residue.
(f)
Allow adhesive to cure for 24
hour at room temperature or for 30 minutes at 145 TO
180 degrees F (63 TO 82 degrees C).
(g)
Refer to paragraph e.
e.
Touch up finish on rotor blade.
Refinish. (AVIM)
(1)
Remove the tip cap assembly and
plug the holes in the end of the spar and inertia weight
to keep out paint.
(2)
Degrease with naphtha (C88) or any
good degreasing solvent.
EA 9340 adhesive is the only adhesive of the
consumable materials (C12) group that is authorized for
splice cover installation.
Provide adequate ventilation when using methyl-ethylketone. Avoid breathing solvent vapors and avoid
prolonged skin contact.
(f)
Apply thin coat of EA 9340
adhesive (C12) to the inside of the splice cover and to
the mating surface of the rotor blade. Use wooden
(3)
If entire blade is to be refinished, strip
old paint from blade with methyl-ethyl-ketone (C87).
5-28
Change 38
TM 55-1520-234-23
(4)
If skins are pitted or eroded in the
area just behind the abrasion strip, polish out the pits
with 320 grit sandpaper (C112). Use fine aluminum
wool (C24) and 320 grit sandpaper (C112) to finish
polish the damaged area. Rub spanwise to remove
burnishing or sandpaper marks and all traces of pitting.
Finish must be minimum of 32 RMS. If the depth of the
repaired area is no greater than 0.008 inch the repair is
satisfactory.
NOTE
Prior to refinishing, blade must have all
scratches, nicks, dents, etc., repaired as shown
under repair of nicks, dents, scratches, notches
and bent trim tab.
(5)
Using abrasive cloth (C44) or
equivalent, remove all surface oxides and all aged
chemical conversion coatings from all bare aluminum
surfaces.
(6)
Wash blade with compound (C39) or
equivalent. Achieve water break free surface, which will
be evident by continuous unbroken film of water on the
surface after thoroughly rinsing the soap from the
surface.
sealant (C116) around cap.
f.
Preparation for Storage or Shipment.
NOTE
The following instructions cover storage or
shipment of main rotor blades in either
cardboard or metal containers.
(1)
Condemn and demilitarize locally any
blade which has incurred non-reparable damage. Refer
to inspection paragraph c, step (7).
(2)
Thoroughly remove foreign matter
from entire exterior surface of blade. Use cleancheese
cloth dampened with naphtha (C88).
(3)
Tape (C134) all holes in the blade
such as bullet damage, tree damage, foreign object
damage, etc. to protect the interior of the blade.
(4)
Apply a coating of wax (C149) to all
exterior surfaces of the blade, except the main retention
bolt hole and the drag brace retention hole. If nonsiliconized composition wax is not available, coat
exterior blade surfaces with oil (C91).
NOTE
From completion of this step through final paint,
surfaces of blades should not be handled with
bare hands.
(7)
On all bare aluminum, apply brush or
spray application of chemical film (C37).
If not
available, use application of phosphoric solution (C23)
or a ten percent solution of chromic acid (C3).
(8)
Thoroughly dry the cleaned surfaces.
Apply a 0.3 to 0.5 mils thick coat of primer(C100). Allow
to air dry from 45 minutes to 4 hours before next step.
(9)
Apply first coat of lacquer (C74A) to
the upper and lower surface of the blade. Allow one
hour minimum drying time, then apply second coat.
Allow one hour minimum drying time before putting any
other paint over the second coat. Spray only the
repaired areas.
(10) Deleted.
(5)
Apply grease (C67) to main bolt hole
and drag brace retention bolt hole.
(6)
Wrap blade with barrier material
(C30), shiny side next to the blade, at all locations when
the blade will contact the hog hair supports (5 places)
and secure with tape (C136).
(7)
Secure contours to the blade at the
paper wrapped areas.
(8)
Attach a properly filled out DD Form
15772 (Unserviceable/Reparable Tag) directly to the
blade.
(9)
blade
to
shock
mounted
(10) Secure lid. If cardboard container is
used, band container shut with 0.50 inch steel bands. If
metal container is used, install top half of container, with
top cushions attached, on lower half of container and
secure with camlock fasteners.
(11) Unplug holes in end of spar and
inertia weight and install tip cap assembly. Apply
Change 22
Secure
support.
5-28A
TM 85-1520-234-23
g.
Installation.
(1)
Support main rotor hub on a build-up
bench. Refer to paragraph 5-4. Check that locating pin
(6, figure 5-11) is installed in upper surface of each grip
(5) at inboard side of retaining bolt hole.
Do not intermix main rotor blades P/N 540011-001-5 with main rotor blades P/N 540011-250-1/540-015-001-1
on
the
same
helicopter because of chordwind moment
differences. Main rotor blades P/N 540-011250-1 and main rotor blade P/N 540-015-0011 may be intermixed on the same helicopter.
NOTE
Corrosion preventive compound (C52) shall be
applied to blade retaining bolts and holes, drag
brace bolts, clevis holes, and to blade butts.
(2)
Insert blade (9) in grip. Place washer
(7) on retaining bolt (8). Align bolt holes carefully and
insert bolt from top. If bolt binds, move tip of blade up
and down slowly to find position which allows bolt to
5-28B
pass through without binding. Seat bolt and washer with
notches on locating pin (6). Apply corrosion preventive
compound (C52) to blade retaining bolts and holes in
hub grip and blade butts.
(3)
Place padded support under blade
approximately one third blade length inboard from blade
tip.
Install washer (16) with counterbore up facing
grip.
(4)
Install washer (16) with counterbore
up as illustrated and install nut (17). Do not tighten nut
at this time.
(5) Align clevis of drag brace (15) on bolt
holes of the blade drag plate. Install shims (10) equally
between clevis and upper and lower drag plates, to
obtain 0.000 to 0.005 inch clearance. Install bolt (14)
and secure with two washers (13 and 12) and nut (11)
on lower side as illustrated. Do not tighten nut (11) at
this time. Apply corrosion preventive compound (C52)
to drag brace bolt and clevis holes.
Change 48
TM 55-1520-234-23
Figure 5-17. Tool application - alignment of main rotor hub and blades
(6) Install opposite blade in the same manner.
(7) If the blades are to be aligned in the hub,
proceed to paragraph h.
(8) If the blades are not to be aligned in the
hub, torque nuts (11) 125 TO 150 foot-pounds. Torque
barrel jam nuts 150 TO 200 foot-pounds.
(9) Use wrench (T31) to tighten nuts (17) to a
torque of 475 TO 525 foot pounds. Align a notch in the
nut with a hole in the bolt. Install locking screw (20) with
head in a direction so that centrifugal force will keep the
locking screw in. In some cases, this may mean the
locking screw may be installed from the inside of the
bolt. Install washer (19) and nut (18).
(10)
Install grip locks (T59) on each pitch
horn if not previously accomplished. See figure 5-7.
h. Alignment.
(1) Install grip locks (T59) on each pitch horn if
not previously accomplished. See figure 5-7.
(2) Place adapter plate (T34) on buildup bench
(T26). Place main rotor hub and blade assembly on
buildup bench. Place a support equipped with wheels
under each blade to support the blades at a precone
angle of 2 1/2 degrees up. The flap stops (T42) may be
installed with 540 side down to ensure stability (figure 535). Place a bubble protractor on the machined surface
adjacent to blade retaining bolts. Both blades should be at
zero degrees chordwise. Remove grip locks (T59) and set
blades to zero degrees chordwise.
(3) Position scope support (T42) over elastomeric
bearings as shown on figure 5-17.
(4) Install scope assembly (T30) on support. Zero
crosshair on an object approximately 50 feet away. Draw a
vertical line on the object. Loosen clamp screws and rotate
scope tube 180 degrees on tube axis in scope clamp and
repeat check. If vertical crosshair does not align with drawn
line, draw a second vertical line on object that will align with
vertical crosshairs. Measure one half the distance between
drawn lines, one and two, and draw a third vertical line.
Adjust screw on side of scope to align vertical crosshairs
with third drawn line. Rotate scope 180 degrees on tube
axis in scope clamp and check that crosshair aligns with
third drawn line. Repeat above steps until satisfactory
adjustment is accomplished.
(5) Locate alignment drive screw (item 8, figure 514 or item 17, figure 5-17a). Sight through the scope
assembly, installed in step (4), and determine whether the
alignment drive screw is lined up with the scope crosshair
within 0.000 inch forward to 0.100 inch aft. If drive is not
aligned within tolerance, adjust drag brace (15, figure 5-11)
to move blade tip and bring alignment drive screw within
tolerance. Be sure that the wheels under the stand
supporting the blade are free to roll when the drag brace is
adjusted. After the blade is aligned, torque jam nuts on the
drag brace 150 TO 200 foot-pounds and recheck to ensure
that blade alignment is still within limits.
Change 71
5-29
TM 55-1520-234-23
(6) Reverse scope to check and adjust alignment of
opposite blade.
Maximum tolerance of alignment
between two blades is 0.050 inch difference or one-half
rivet head diameter.
between stations 213.5 and 260.0. An almost invisible
seam may be detected at station 213.5. The leading edge
erosion guard is completely composed of estane material.
No other material com- position is used.
(7) Torque nuts (11, figure 5-11) on both blades
125 TO 150 foot pounds after blades are aligned.
(2) K747-003-205/309 Deviation implements the
use of fluorocarbon leading edge erosion guard material
between stations 213.5 and 260.0. This is a higher impact
resistant material covering implemented IBWR features. A
very slight difference in sheen between the estane and
fluorocarbon material may be detected. A very slight seam
may be visible at station 213.5. If these differences are not
readily ap- parent, the blade log component DA Form 240816 must be consulted.
(8) Torque nuts (17, figure 5-11) on both blades
475 TO 525 foot-pounds after blades are aligned. Use
socket wrench, T101414 to tighten nuts. Select a notch
in the nut with a hole in the bolt and install locking screw
(20) with head in a direction so that centrifugal force will
keep the lockscrew in. In some cases, this may mean
the locking screw may be installed from the outside of
the bolt. Install washer (19) and nut (18).
(9) Verify blade alignment.
(10) Remove grip locks (T59) and flap stops
(T42).
5-5A.
K747 MAIN ROTOR BLADES
5-5B.
DESCRIPTION-K747 MAIN ROTOR BLADES.
NOTE
After incorporation of drag strut K747-082-1 and
root fitting K747-083-1, per MWO 55-1520-24450-11, main rotor blades K747-003-205, -209,
and -303, become K747-003-309, -401, and 403, respectively.
a. The K747-003 improved main rotor blade
(figures 5-17A and 5-17B) is an advanced technology
composite structure which offers improved performance,
reliability, maintainability, and reduced radar cross
section. It is a glass fiber epoxy resin bonded assembly
with an elastomeric erosion guard. The blade is attached
in the hub with a retaining bolt assembly (root fitting)
and is held in alignment by a drag strut.
b. Difference Between Models. K747-003 series
main rotor blades have the following part numbers and
differences as noted. (See table 5-1A).
(1) K747-003-205/309
incorporates
an
improved blade weight retention (IBWR) feature not
implemented in earlier K747 blades. (See figures 5-17B
and 5-17C). This change alters physical appearance in
that a slightly raised area is visible on the top and
bottom of the leading edge (LE) erosion guard surfaces
(3) K747-003-209/401 is visually the same as the 205/309 blade with the exception of the leading edge
erosion guard. A full fluorocarbon guard is used instead of
a full guard composed of estane materials. If there is any
doubt as to material composition, the blade log component
DA Form 2408-16 and blade ID plate must be consulted.
CAUTION
K747-003-303/-403 blades shall only be flown with
other -303/-403 blades. They shall not be flown
with K747-003-205/-309, -205/-309 deviation, 209/-401, or -303/-403 field modified blades. The
K747-003-303/ -403 blade can be easily identified
by the stainless steel erosion guard installed on the
outboard leading edge. Do not mistake a -303/-403
field modified blade (stainless steel guard removed
and screws and shields installed) for -303/-403
blade.
(4) K747-003-303/-403
incorporates
all
the
improvements of the -209/-401 blade and adds a stain- less
steel erosion guard over the fluorocarbon guard on the
blade outer leading edge. Stainless steel guard location is
between stations 217.5 and 261.0. The -303/-403 blade
cannot be mixed with other blade models for flight. All
skin/core repairs and repair kits used to repair -205/-309, 205/-309 deviation, and -209/-401 blades can be used to
repair the -303/-403
blade.
CAUTION
K747-003-303/-403 field modified blades shall not
be flown with K747-003-303/-403 blades. K747003-303/-403 field modified blades are easily
identified by the absence of a stainless steel
erosion guard and by six shields secured by selflocking screws located on the outboard area of the
fluorocarbon erosion guard.
5-30 Change 71
TM 55-1520-234-23
Figure 5-17A. K747 Main Rotor Blade (Part Numbers K747-003-205, -209, -303)
Change 65
5-30A/(5-30B blank)
TM 55-1520-234-23
Figure 5-17A.1. K747 Main Rotor Blade (Part Numbers K747-003-309, -401, -403)
Change 65
5-30C
TM 55-1520-234-23
TABLE 5-1A. DIFFERENCE BETWEEN MODELS
5-30D
Change 65
TM 55-1520-234-23
Figure 5-17B.
External Appearance Changes to K747-003-205/-309, -209/-401, and -303/-403 Blades Resulting
From Improved Weight Retention Features.
Change 65 5-30E
TM 55-1520-234-23
Figure 5-17C.
Internal Modifications Incorporated in K747-003 -205/-309, -209/-401, and -303/-403 Blades for
Improved Weight Retention.
5-30F Change 65
TM 55-1520-234-23
(5) K747-003-303/-403
field
modification
removes the stainless steel erosion guard from the 303/-403 blade. This modification makes the -303/403
blade compatible for use with the -205/-309, -205/-309
deviation, -209/-401, and other -303/-403 field modified
blades only. All skin/core repairs and repair kits used to
repair -205/-309, -205/-309 deviation, and -209/-401
blades can be used to repair -303/-403 field modified
blades.
5-5C.
f. Remove blade from grip and place in a padded
stand.
g. Remove opposite blade from hub in same manner.
5-5D.
CLEANING - K747 MAIN ROTOR BLADES.
a. Clean main rotor blade with one part cleaning
compound IC41) and nine parts water.
REMOVAL - K747 MAIN ROTOR BLADES.
a. Position main rotor hub and blade assembly on
build-up bench (paragraph 5-4). Place padded supports
under blades so that leading edge is ap- proximately
straight.
Cleaning solvent is flammable and toxic.
Provide adequate ventilation. Avoid prolonged
breathing of vapors and contact with skin or
eyes.
b. Remove locking screw (20, figure 6-11.
c. Remove nut (17) and washer (16) with blade
bolt wrench (T31).
CAUTION
d. Remove nut (11), washers (12 and 13), and bolt
(14). Loosen nut 121) and swing drag brace (15) away
from rotor blade. Retain shims (10) for reinstallation.
The erosion boot is very susceptible to
solvents. Use care to prevent spillage or runoff of solvents onto the boot.
CAUTION
Avoid blade contact with the drag brace and
hub during removal procedure to prevent
possible blade damage.
e. Remove blade retaining bolt (8) and washer 17).
Slowly raise and lower blade tip while tapping bolt with
fiber mallet. If bolt is difficult to remove, use a bolt
removal work aid similar to the one shown in figure 512. Remove blade retaining bolt as follows:
(1) Remove threaded plugs from upper and
lower ends of blade retaining bolt. If weights are present
in bolt, retain for reinstallation.
(2) Position work aid on bolt as shown in figure
5-13, and also place a piece of hard rubber or similar
material between work aid tube and grip to prevent
marring the grip. Hold puller rod and tighten hexagon
nut to remove blade retaining bolt.
(3) Remove work aid from blade retaining bolt.
Reinstall weight and plugs in blade retaining bolt and
identify the blade retaining bolt for reinstallation in the
same grip.
b. Remove stubborn deposits with a cloth dampened
with solvent (C124) except the boot shall be cleaned only
with detergent (C59BI or one part cleaning compound (C41)
and nine parts water.
5-5E.
INSPECTION - K747 MAIN ROTOR BLADES.
a. Inspect blade historical records and the blade for
evidence that the blade has been subjected to an accident,
overspeed or incident outside the realm of normal usage. If
such evidence exists, perform special inspections outlined
in paragraph 1-31.
b. Inspect blades for damage. Classify damage as
acceptable or repairable, using the limits in table 5-1B and
table 5-1C. Acceptable damage shall not be repaired.
c. The top of the painted surface shall be used to
measure dents, cuts and scratches by using a dial indicator
with a probe. The fibers of the basketweave may appear to
be raised or rough; this is not cause for rejection.
NOTE
Change 65
Faces of the spar wrap fittings with cracks at the 6
o'clock and 12 o'clock positions.
5-30G
TM 55-1520-234-23
plus or minus 30°, are acceptable. Cracks on all
eight faces and across the bolt hole on one
blade are acceptable. Multiple cracks on one
face at the 6 and/or 12 o'clock position are
acceptable providing no metal is lost.
blade must be inspected. Inspect for crescent-shaped,
raised areas or circular delaminations of the erosion guard
from the spar. These will appear as circular raised areas
approximately 2.0 inches in diameter, not to exceed 0.060
inch in height. (See figure 5-17F.)
d. Inspect root fitting for damage in accordance
with figure 5-17D and figure 5-17D.1.
b. If the discrepancy noted in step a above is
detected, replace the blade.
NOTE
5-5G. REPAIR OR REPLACEMENT - K747 MAIN
ROTOR BLADES.
Ensure the cotter pins are installed in the
castellated nuts on the attaching hardware
of the root fitting. Root fitting attaching
hardware may turn by hand. This is not an
indication of a loss of torque and is an
acceptable installation.
e. Inspect drag strut for damage in accordance
with figure 5-17E and figure 5-17E.1.
f. Tap test blade spar area for cracks between
station 70 to 90 from leading edge erosion guard to back
side of spar. The tap sound from uncracked area to
cracked area is from a solid sound to a highly muffled
sound.
5-5F. INSPECTION - K747 MAIN ROTOR BLADE
TIP WEIGHT RETENTION.
a. Inspect blades in tip area of leading edge of
erosion guard. Both upper and lower surfaces of the
5-30H
WARNING
The following protective equipment must be
used when working with fiberglass repair kits:
Respirator, Chemical Cartridge
Respirator, Disposable Half-Mask
Gloves, Rubber: Acid, Alkali Resistant, Black
Apron, Impermeable: Duck, Rubber Coated
Goggles, Industrial for Chemical Handling
Faceshield, Industrial, Hinged Window
CAUTION
Use only tools specified for repair of K747 main
rotor blades.
a. Main rotor blades meeting all of the following
requirements shall be repaired.
Change 75
TM 55-1520-234-23
Table 5-1B. Classification of Damage - K747 Main Rotor Blades
Change 65 5-30J
TM 55-1520-234-23
Table 5-1B. Classification of Damage - K747 Main Rotor Blades (Continued)
5-30K
Change 49
TM 55-1520-234-2'
Table 5-1B. Classification of Damage - K747 Main Rotor Blades (Continued)
Change 49
5-30L
TM 55-1520-234-23
Table 5-1B. Classification of Damage - K747 Main Rotor Blades (Continued)
5-30M
Change 65
TM 55-1520-234-23
Table 5-1C. Classification of Damage - K747 Main Rotor Blades
TYPE OF DAMAGE
COMPONENT
a. Root fitting
A-Acceptable. Do not repair. U -Unacceptable. Replace. ALL DIMENSIONS
R - Reparable if within requirements of paragraph 5-30
ARE IN INCHES
NICKS, SCRATCHES, SKIN EROSION
CORROSION
See figure 5-17D
See figure 5-17D
b. Drag strut
See figure 5-17E
See figure 5-17E
c. Root fitting bolt
A
To 0.050 deep on hex and exposed
thread area.
d. Cheek plate assembly
A
Inboard of sta 48 to 0.015, no limit
on length or number.
R
To 0.050 deep on hex head
and exposed thread area (by
polishing).
R
To 0.035 deep.
e. Cheek plate fitting
A
To 0.015 deep on exposed surfaces.
R
To 0.030 deep on exposed
surfaces (by polishing).
A
Cracks at 12 & 6 o'clock positions on
faces with root fitting removed.
R
0.015-0.030
deep
on
surfaces(by polishing).
f. Balance weight covers
exposed
A
Nicks, scratches, dents, and bends to
0.035 deep.
R
Bends and distortion over 0.035 deep
(by mechanical straightening).
g. Trailing edge fitting
A
To 0.015 deep on exposed surfaces.
Change 49
5-30N
R
To 0.030 deep on exposed
surfaces (by polishing).
TM 55-1520-234-23
Table 5-1C. Classification of Damage - K747 Main Rotor Blades (Continued)
TYPE OF DAMAGE
COMPONENT
A-Acceptable. Do not repair. U -Unacceptable. Replace. ALL DIMENSIONS
R - Reparable if within requirements of paragraph 5-30
ARE IN INCHES
NICKS, SCRATCHES, SKIN EROSION
CORROSION
R
0.015-0.030 deep on exposed surfaces
(by polishing). No limit on length or
number.
h Tip cap forward
(see note)
A
Erosion:
To 0.030 deep.
R
0.030-0.060 deep (by polishing).
A
Until loss of weight causes
helicopter vibration.
R
Gap around tip cap. Sealer
missing. Apply sealant
(C107).
i. Tip weight cover
No limit on length or number.
Erosion:
A
To 0.030 deep.
A
To 0.030 deep.
R
0.030-0.060 deep (by polishing).
NOTE
The balance weights consist of several metal plates that are free to move within the cavity of the tip-balance weight cap.
If the cavity is not full, the balance weight plates are able to move, producing the clicking noise during the static vertical
movement of the blade. The clicking noise is to be considered normal, unless the tip balance weight cap is not properly
secured to its attaching pocket bracket.
5-30P
Change 49
TM 55-1520-234-23
Figure 5-17D. Damage Limits - Root Fittings (P/N K747-061-5) (K747 Blade)
Change 65 5-30P.1
TM 55-1520-234-23
Figure 5-17D.1. Damage Limits - Root Fittings - P/N K747-083-1 (K747 Blade)
5-30P.2
Change 65
TM 55-1520-234-23
Figure 5-17E. Damage Limits - Drag Strut - P/N K747-072-1 (K747 Blade)
Change 65
5-30Q
TM 55-1520-234-23
Figure 5-17E.1. Damage Limits - Drag Strut - P/N K747-082-1 (K747 Blade)
5-30R
Change 65
TM 55-1520-234-23
Figure 5-17F. Inspection of K747-003-205/-309, 209/-401, and -303/-403 Blade for Loss
of Blade Weight Retention Integrity.
Change 65 5-30S
TM 55-1520-234-23
(1) Blade shall have only damage that is listed
as repairable or acceptable in table 5-1B and table 5-1C,
and damage shall not be in any area previously
repaired. Damage listed as acceptable shall not be
repaired.
d. Remove drag strut (15) and set aside.
NOTE
Drag strut (15) is not interchangable between
rotor blades and must be reinstalled to match
up with the rotor blade from which it was
removed.
(2) All required repairs shall be within the
proximity limits shown in figure 5-17G.
(3) Blade shall contain sufficient existing
balance weight to permit adjustment of blade balance as
shown in figure 5-17H.
(4) Blades showing evidence of blade weight
retention failure as defined in paragraph 5-5F shall not
be repaired.
b. Main rotor blades not meeting the requirements
of step a. above, shall be replaced.
e. Remove cotter pins (16), nuts (17), washers (18, 19,
and 21), and bolt (22) to remove root fitting (23). Do not
remove bushings (20)
f.
Replace root fitting (23) with a serviceable root fitting.
g. To install root fitting (23), position on blade, using a
soft-headed mallet.
h. Insert two locator pins, to assure holes are in alignment.
Dowel pins are acceptable.
K747 main rotor blades shall not be
intermixed with main rotor blades of any
other type on the same helicopter, because
of performance difference.
i. Remove top locator pin and install washers (21) as
required on bolt (22). Drive bolt into position using softheaded mallet.
j. Remove lower locator pin and install washers (21) as
required on bolt (22). Drive bolt into position using softheaded mallet.
k. Install washers (18 and 19) and nut (17) on each bolt
(22).
NOTE
The following repairs can be made on the
top and bottom of main rotor blade while
blades are installed on helicopter. When
repair limits are questioned, proceed to next
critical repair procedure.
5-5H. REMOVAL AND REPLACEMENT OF ROOT
FITTING - K747 MAIN ROTOR BLADES (K747-003205, -209, AND -303 BLADES).
a. Remove K747 main rotor blade assembly
(paragraph 5-5C).
l. Draw nuts up tight to ensure bolt heads are properly
seated. This prevents any possibility of false readings
when torquing nuts to specification.
m. Back off nuts for torquing. Torque nuts 30 to 150 inchpounds. Adjust to ensure cotter pin hole alignment. Shim
as necessary to gain hole alignment within torque range.
Use washers (19) for shimming.
n. Install new cotter pin (16) in bolts. Trim cotter pin
length. Bend long cotter pin length over end of bolt and
short length towards bolt head.
b. Remove cotter pin (1, figure 5-17J), nut (2),
washers (3), bushing (3.1), and bolt (4).
o. Position drag strut (15) on blade. Align aft inboard and
outboard bushings of drag strut with bushing holes in
trailing edge wrap fitting.
c. Remove cotter pins (5 and 10), nuts (6 and 11),
washers (7, 12 and 12.1), and bolts (9 and 14). Using
tapered pin, gently drive bushings (8 and 13) aft to
remove.
p. Install aft inboard bushing (8) and outboard bushing
(13). Use a soft-headed mallet, if necessary. Install
bushing in aft to forward direction (trailing edge).
5-30T
Change 65
TM 55-1520-234-23
Figure 5-17G. Proximity Limits for Patches - K747 Main Rotor Blades (Sheet 1 of 3)
Change 49
5-30U
TM 55-1520-234-23
Figure 5-17G. Proximity Limits for Patches - K747 MAIN ROTOR BLADES (Sheet 2 of 3)
5-30V
Change 49
TM 55-1520-234-23
Figure 5-17G. Proximity Limits for Patches - K747 Main Rotor Bldes (Sheet 3 of 3)
Change 49
5-30W
TM 55-1520-234-23
Figure 5-17H. Balance Adjustment for Patches (K747 Blade) (Sheet 1 of 2)
5-30X
Change 49
TM 55-1520-234-23
Figure 5-17H. Balance Adjustment for Patches (K747 Blade) (Sheet 2 of 2)
Change 56
5-30Y
TM 55-1520-234-23
Figure 5-17J. Root Fitting Assembly (K747-205, -209, -303 Blades)
5-30Z
Change 65
TM 55-1520-234-23
Figure 5-17J.1. Root Fitting Assembly (K747-309, -401, -403 Blades) (Sheet 1 of 2)
Change 65
5-30Z.1
TM 55-1520-234-23
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
19.
20.
21.
22.
23.
24.
25.
26.
27.
28.
29.
COTTER PIN
NUT
WASHER
WASHER
BOLT
COTTER PIN
NUT
WASHER
BOLT
COTTER PIN
NUT
WASHER
WASHER
BOLT
COTTER PIN
NUT
WASHER
WASHER
BOLT
BUSHING, INBOARD
BUSHING, OUTBOARD
BUSHING, SLIP FIT
DRAG STRUT
ROOT FITTING
WASHER
WASHER
WASHER
WASHER
PLATE IDENTIFICATION
Figure 5-17J.1. Root Fitting Assembly (K747-309, -401, -403 Blades)
(Sheet 2 of 2)
5-30Z.2
Change 65
TM 55-1520-234-23
Figure 5-17J.2. Drag Strut Assembly (K747-309, -401, -403 Blades)
Change 65
5-30Z.3/(5-30Z.4 blank)
TM 55-1520-234-23
q. Install washer (7) on inboard bolt (9). Insert bolt
(9) through trailing edge wrap fitting and drag strut (15)
in an aft to forward direction. Use a soft-headed mallet.
Install washer (7) and nut (6) on bolt (9).
r. Install washers (12 and 12.1) on outboard bolt
(14). Insert bolt (14) through drag strut (15) and trailing
edge wrap fitting in forward to aft direction. Use a softheaded mallet. Install washer (12) and nut (11) on bolt
(14).
e. Drive two bushings (20 and 21) out of drag strut
and trailing edge wrap fitting in aft direction, supporting
drag strut assembly during bushing removal, using a
phenolic, aluminum, or wooden dowel and a soft-headed
hammer. Keep dowel axis in line with axis of bushing being
removed.
f.
Remove drag strut from blade and set aside.
NOTE
s. Install bushing (3.1) in root fitting (23). Install
washer (3) on bolt (4). Insert bolt (4) through root fitting
(23) and drag strut (15). Install washer (3) and nut (2).
t. Draw nuts (2, 6, and 11) up tight to ensure bolt
heads are properly seated. This prevents any possibility
of false readings when torquing nuts to specification.
u. Torque nut (6) to 480 to 540 inch-pounds.
Ensure cotter pin hole alignment. Torque nuts (2 and
11) to 120 to 150 inch-pounds. Ensure cotter pin
alignment.
v. Install new cotter pins (1, 5, and 10). Trim
cotter pin length. Bend long cotter pin of bolt and short
length towards bolt head.
5-5H.1.
REMOVAL AND REINSTALLATION OF
ROOT FITTING-K747 MAIN ROTOR BLADES (K747003-309, -401, -403 BLADES).
a. Remove K747 main rotor blade assembly
(paragraph 5-27).
b. Remove cotter pins (1 and 6, Figure 5-24.2),
nuts, (2 and 7), washers (3, 4, 8, 26, 27), and bolts (5
and 9), from drag strut at trailing edge wrap fitting.
c. Remove cotter pin (10), nut (11), washers (12,
13, and 28), and bolt (14), securing drag strut to root
fitting.
d. Remove bushing (22) from leading edge of drag
strut at root fitting.
CAUTION
Do not drive inboard aft drag strut bushing out in
direction of leading edge. Irreparable damage to the
drag strut could result. Support drag strut assembly
during bushing removal.
Change 71
Drag strut (23) is not interchangeable between rotor blades
and must be reinstalled to match up with the rotor blade
from which it was removed.
g. Remove cotter pins (15), nuts (16), washers (17,
18, 25), and bolts (19), to remove root fitting.
h. Replace root fitting (24) only if required. Refer to
Table 5.1.1 and Figure 5-19.2.
i. To install root fitting (24), position on blade using a
soft-headed mallet.
j. Insert two locator pins, to ensure holes are in
alignment. Dowel pins are acceptable.
CAUTION
Beveled ID on washer (18) must face bolt head. Do not
damage bolt threads or bushing bores when installing bolts.
k. Apply corrosion preventive compound, Brayco 599
(C44.1) to bolt shank. Remove top locator pin and install
washer (18), beveled ID toward bolthead, on bolt (19).
Drive bolt into position using a soft-headed mallet.
l. Apply corrosion preventive compound, Brayco 599
(C44.1) to bolt shank. Remove lower locator pin and install
washer (18), beveled ID toward bolthead, on bolt (19).
Drive bolt into position using a soft- headed mallet.
m. Install washer (17) under nut (16) on each bolt. If
required, add washers (17 and/or 25) under nut (16) on
each bolt to obtain cotter pin hole alignment.
n. Draw all nuts up tight to ensure boltheads are
properly seated. This prevents any possibility of false
readings when torquing nuts to specification.
5-30AA
TM 55-1520-234-23
Figure 5-17K. Application of Skin Patch (K747 Balde) (sheet 1 of 2)
5-30AB Change 67
TM 55-1520-234-23
o. Back off nuts for torquing. Torque nuts to 30 to
150 inch-pounds. Adjust to ensure cotter pin hole
alignment.
p. Insert two cotter pins (15) in bolts. Bend long
cotter pin length over end of bolt and bend short length
toward bolthead.
q. Position drag strut (23) on blade. Align aft
inboard and outboard bushings with bushing holes in
trailing edge spline wrap fitting.
r. Apply corrosion preventive concentrate, Brayco
599 (C44.1), to the outside diameters and install aft
inboard and outboard bushings (20 and 21) in an aft-toforward direction, using a soft-headed mallet, if
necessary.
s. Install washer (3), with bevel to bolthead, on
larger bolt (5). Apply corrosion preventive concentrate,
Brayco 599 (C44.1), to bolt shank and insert bolt
through drag strut and trailing edge wrap fitting in an aftto-forward direction. Check for bind- ing; none allowed.
Install washer (4) and castellated nut (2) on bolt (5). If
required, add washers (4 and/or 26) under nut (2) to
obtain cotter pin hole alignment.
NOTE
This stack-up may be capable of being turned
by hand after proper installation.
t. Install washer (8) on smaller outboard bolt (9).
Apply corrosion preventive concentrate Brayco 599
(C44.1) to bolt shank and insert bolt through drag strut
and trailing edge spline wrap fitting in for- ward-to-aft
direction. Install washer (8) and castellated nut (7) on
bolt. If required, add washers (8 and/or 27) under nut
(7) to obtain cotter pin hole alignment.
stalled, then install washer (12), with bevel to bolt- head,
and washer (13) on bolt (14). Apply corrosion preventive
concentrate Brayco 599 (C44.1) to bolt shank and insert
bolt through root fitting and drag strut from root fitting top
side. Install washer (13) and castellated nut (11) on bolt. If
required, add washers (13 and/or 28) under nut (11) to
obtain cotter pin hole alignment.
w. Ensure that head of bolt (5) and washer (3) are
flush against the aft surface of the drag strut. If necessary,
tap bolthead forward using a soft headed mallet.
NOTE
To improve fatigue-loading capabilities of the drag
strut at the inboard connection to blade trailing
edge, there shall be a mini- mum of 0.010 inch
clearance between washer(s) and forward surface
of fitting after torquing. (Figure 5-24.3)
x. Torque large inboard nut (2, Figure 5-24.2) to 480
to 540 inch-pounds to include torque drag. Ensure cotter
pin hole alignment.
y. Torque smaller outboard nut (7) to 120 to 150 inchpounds to include torque drag. Ensure cotter pin hole
alignment.
z. Torque root fitting nut (11) to 300 to 420 inchpounds to include torque drag. Ensure cotter pin hole
alignment.
aa. Install new cotter pins (1, 6, and 10). Trim cotter
pin length. Bend long cotter pin length over end of bolt and
bend short length toward bolt head. Ensure that no sharp
edges are exposed.
5-5J. POLISHING AND CORROSION TREATMENTK747 MAIN ROTOR BLADES (AVUTM).
NOTE
A loose slip fit of this installation stack-up is
acceptable.
u. Apply corrosion preventive concentrate, Brayco
599 (C44.1), to outside diameter of slip fit bushing
(22) and install bushing in top clevis of root fitting.
v. Ensure that slip fit bushing (22) has been in-
Change 71
a. Polish out nicks, scratches, and corrosion on
exposed metallic parts with No. 320 or finer sandpaper
(C112) and touchup as required in accordance with
paragraph 5-5AC.
b. Repairs requiring removal of drag strut (15, figure
5-17A) may be made by removing attaching hardware.
Following repairs as specified in step a. above, reinstall
drag
strut
using
figure
5-17A
as
a
guide.
5-30AB.1
TM 55-1520-234-23
5-5K. APPLICATION OF SKIN PATCH-K747 MAIN
ROTOR BLADES (AVIM).
patch kits are available in the following sizes.
Kit No.
Patch Diameter
K747-201-1
K747-201-3
K747-201-5
3 inch
5 inch
9 inch
CAUTION
Blade must contain sufficient balance weight to
permit adjustment of blade balance after repair.
Refer to paragraph 5-5G before starting any
repair.
Grease or lead pencils will not be used. Only
ball point pens will be used to make lines as
shown. Marks other than those specified can
weaken the repair.
a. Obtain blade repair tool set P/N K747-401-1
(T91).
f. Damage passing through both skins with core
damage of less than 1 inch diameter shall be repaired by
applying a skin patch to both top and bottom sides of blade.
g. Place the template (kit) on the blade. Position the
inner circle to enclose the damage. Hold the template from
slipping, and draw a line around the outer circle of the
template (View A, figure 5-17K).
CAUTION
b. Position blade for access to damaged area
(figure 5-17K). Support blade to prevent movement and
droop.
c.
Measure diameter of damage.
d. Obtain adhesive package (C7A) or alternate
(C17).
e. Obtain skin patch repair kit no larger than
necessary to overlap damage 1 inch all around. Skin
5-30AB.2
Excessive sanding will weaken blade skin. Sand
only until yellow color is removed.
h. Starting with 120 grit and finishing with 220 grit
abrasive paper (kit), sand the paint and yellow primer from
the blade from the area within the guide circle. Sand only
until yellow color is removed. Do not sand skin fibers.
Also, sand off any damaged material raised above normal
contour of blade (View A, figure 5-17K).
Change 71
TM 55-1520-234-23
View B. Application of adhesive and positioning of patch.
Figure 5-17K. Application of Skin Patch (K747 Blade) (Sheet 2 of 2)
Change 67 5-30AC
TM 55-1520-234-23
Cleaning solvent is flammable and toxic.
Provide adequate ventilation.
Avoid
prolonged breathing of vapors and contact
with skin or eyes.
CAUTION
Care shall be taken to prevent MEK from
entering core area of blade. Spillage shall
be avoided. MEK can damage leading edge
erosion guard.
i. Put on cotton gloves (kit), then plastic gloves
(kit). Leave on until completion of step r. Dampen
cheesecloth (kit) with MEK (C 87).
j.
Wipe off all sanding dust.
k.
Use template to redraw guide circle.
Adhesive contains toxic ingredients. Provide
adequate ventilation and protect the skin and
eyes from contact with uncured resins or
curing agent. Wash off uncured resins and
curing agent from skin with warm water and
soap. Avoid use of solvents for cleaning the
skin.
NOTE
Never mix less than a complete two-part package
of adhesive (C7A). Mix the full batch and then
discard the excess after the repair is completed.
n. Mix
adhesive
(C7A)
per
manufacturer's
instructions. Stir with wooden spatula until color is uniform
and all streaks have disappeared. Adhesive (C17) may be
used as an alternate.
NOTE
l. Cut short lengths of the masking tape (kit) and
mask around the outside of guide circle (View B, figure
5-17K).
Cleaning solvent is flammable and toxic.
Provide adequate ventilation.
Avoid
prolonged breathing of vapors and contact
with skin or eyes.
CAUTION
Care shall be taken to prevent MEK from
entering core area of blade. Spillage shall
be avoided. MEK can damage leading edge
erosion guard.
m. Dampen clean cheesecloth (kit) with MEK (C87)
and clean inside masked area. Wipe with clean dry
cheesecloth before dampness evaporates.
5-30AD
Pot life of adhesive (C7A) is 15 minutes at 75
degrees F (23.8 degrees C). It is shorter at higher
temperatures. Always check package dates to
make sure the maximum adhesive life time limit of
1 year is not exceeded. Pot life of adhesive (C17)
is approximately 1/2 to 1 hour at 75 degrees F
(23.8 degrees C).
o. Using clean one inch brush (kit), apply a light coat
of adhesive to blade skin, within guide circle, and to
underside of skin patch (View B, figure 5-17K).
p. Center skin patch within guide circle, with stenciled
arrow pointing outboard (spanwise), and press firmly into
place. Slide patch back and forth slightly under hand
pressure to even adhesive. Use light hand pressure to
squeeze the patch from the center to edge to work out any
air bubbles.
Cleaning solvent is flammable and toxic.
Provide adequate ventilation. Avoid prolonged
breathing of vapors and con- tact with skin or
eyes.
Change 49
TM 55-1520-234-23
Figure 5-17L. Curing Patch with Blade Repair Fixture (K747 Blade)
Change 49 5-30AE
TM 55-1520-234-23
power and relieve air pressure by lifting center portion of
relief valve.
CAUTION
Care shall be taken to prevent MEK from
entering core area of blade. Spillage shall
be avoided. MEK can damage leading edge
erosion guard.
t.
Remove repair fixture from blade.
u. Refinish repair area.
(1) Remove peel-ply and masking tape from blade.
q. Using clean cheesecloth (kit) dampened with
MEK (C87), temporarily lift edges of peel-ply and wipe
off excess adhesive.
r. Place masking tape over edge of patch in four
places to prevent movement of patch. Place two long
pieces of masking tape at right angles, centered over
the patch spanwise and chordwise and extending
beyond the dimensions of the blade repair fixture
bladder.
s.
Install blade repair fixture (T88) (figure 5-17L).
CAUTION
Sanding skin fibers can weaken blade.
(2) Using 220 or finer grit abrasive paper (kit),
feather edge of adhesive squeeze-out around patch.
(3) Paint repaired area in accordance with
paragraph 5-5AC.
v. Adjust blade balance weights as required by figure
5-17H.
(1) Install from trailing edge side of blades only.
(2) Center bladder over repair area and secure.
(3) Center pad opposite bladder and secure.
CAUTION
w. K747 blade repairs are required to be logged in DA
Form 2408-13 and -16. A permanent record must be
maintained to determine the minimum spacing requirement
between repairs. Once a repair has been made, it is not
possible to determine which type of repair has been
applied.
Tightening of locking knobs so that metal
skirt around bladder is closer than 0.125
inch to blade can damage blade.
5-5L. INSTALLATION OF PLUG PATCH - K747 MAIN
ROTOR BLADES (AVIM).
(4) Tighten fixture channel locking knobs until
metal skirt around bladder is approximately 0.125 inch
from blade skin.
CAUTION
(5) Actuate hand pump to obtain 4 psi minimum
reading on pressure gage. Disconnect pump hose
clamp from air valve.
Blade must contain sufficient balance weight to
permit adjustment of blade balance
after
repair. Refer to paragraph 5-5G before starting
any repair.
a. Position blade for access to damaged area.
Support blade to prevent movement and droop.
NOTE
During curing, it may be necessary to
periodically reconnect hose and to actuate
pump to maintain 4 psi minimum.
(6) Connect 110 volt ac electrical power for
curing time shown in table 5-1D.
b. Measure diameter and depth of damage.
figure 5-17M.)
c. Obtain plug patch repair kit no larger than
necessary to replace damage. A core void 1 inch or less in
diameter is permitted after repair. Plug patch kits are
available as shown in table 5-1D.
(7) At end of curing time, disconnect electrical
5-30AF
(See
Change 49
TM 55-1520-234-23
d. Damage not more than 1.750 inches deep can
be repaired with a single patch. Damage that passes
completely through blade and is larger than 1 inch in
diameter, will be repaired by installing plug patches
from both top and bottom sides of blade. Install larger
diameter and depth plug patch first figure (5-17N).
e. Obtain required number of adhesive packages
(C7A) as shown in table 5-1D. If adhesive (C17) is
used, obtain an equal amount.
Grease or lead pencils will not be used. Only
ball point pens will be used to make lines as
shown. Marks other than those specified can
weaken the repair.
f. Place the template (kit) on the blade. Position
the inner circle to enclose the damage. Hold the
template from slipping and draw lines around the inner
and outer circles of the template (View A, figure 517M).
Excessive sanding will weaken blade skin.
Sand only until yellow color is removed.
g. Starting with 120 grit and finishing with 220
grit abrasive paper (kit), sand the paint and the yellow
primer from the blade from the area between circles A
and B. Sand only until yellow color is removed. Do not
sand skin fibers (View A, figure 5-17M).
Cleaning solvent is flammable and toxic.
Provide adequate ventilation.
Avoid
prolonged breathing of vapors and con- tact
with skin or eyes.
Care shall be taken to prevent MEK from
entering core area of blade. Spillage shall be
avoided. MEK can damage leading edge
erosion guard.
h. Put on cotton gloves (kit), then plastic gloves
(kit). Dampen cheesecloth (kit) with MEK (C87). Wipe
off sanding dust.
i. Redraw circle A.
guideline.
This circle is the routing
Table 5-1D. Plug Patch Data
Kit
Part
No.
Plug
Dia.
Plug
Depth
K747-201-7
K747-201-9
K747-201-101
K747-201-103
K747-201-105
K747-201-107
K747-201-109
K747-201-111
3 in.
3 in.
3 in.
3 in.
7 in.
7 in.
7 in.
7 in.
0.250 in.
0.500 in.
1.250 in.
1.750 in.
0.250 in.
0.500 in.
1.250 in.
1.750
Adhesive
Pkg.
Req.
(32 gm each)
See note
0.333
0.333
0.666
1.0
1.0
1.250
2.0
2.500
Patch
Over
Core
Only
15
15
30
30
15
15
45
45
Minutes Cure
Patch
Over
Core/Spar
Core/Trailing Edge
30
30
30
30
30
20
45
45
NOTE: 32 gm of mixed alternate adhesive (C17) bulk materials will be equal to one adhesive kit (C7A).
Change 67
5-30AG
TM 55-1520-234-23
Figure 5-17M. Installation of Plug Patch (K747 Blade (Sheet 1 of 4)
5-30AH
Change 67
TM 55-1520-234-23
Figure 5-17M. Installation of Plug Patch (K747 Blade ) (Sheet 2 of 4)
5-30AJ
Change 67
TM 55-1520-234-23
Figure 5-17M. Installation of Plug Patch (K747 Blade (Sheet 3 of 4)
Change 49
5-30AK
TM 55-1520-234-23
Figure 5-17M. Installation of Plug Patch (K747 Blade ) ( Sheet 4 of 4)
5-30AL
Change 67
TM 55-1520-234-23
Figure 5-17N. Typical Double Plug Patch Repair (K747 Blade)
Change 49
5-30AM
TM 55-1520-234-23
m. Wipe off all cuttings, sanding dust, etc. from
repair area.
n. Use template to redraw circle B.
Disconnect router cord from outlet before
changing or installing bits or end mills, or
making adjustments.
Ensure router switch is in off position before
connecting router to electrical power.
o. Cut short lengths of masking tape (kit) and
mask around the outside of circle B (View D, figure
5-17M.
p. Put on cotton gloves (kit), then plastic gloves
(kit). Leave on until completion of step y.
Keep hands and fingers away from rotating
bits and end mills.
Guide router with both hands on router grip.
Use
personal
protection
respirator, goggles, apron, etc.
equipment;
During all routing operations, long dimension
of route base shall be kept in spanwise
direction.
End mills will burn out if used to cut skin.
It is absolutely necessary to take every
precaution not to damage the spar and trailing
edge during routing. The spar in the leading
edge and trailing edge can be located by using
the instructions in figure 5-17G.
j. Insert rasp-type bit, P/N 4-BR, in router collet.
Set router depth of cut for 0.1875 inch. Rout a complete
circle through the skin, inside of, and following circle A
(View B, figure 5- 17M).
k. Using duckbill pliers, lift the edge of the cut
circle of skin and peel the cut circle of skin off core
(View B, figure 5-17M). After removing skin, check
depth of core at trailing edge of circle. Core thickness at
trailing edge side less than depth of plug selected will
require use of more shallow plug or a double plug patch
repair.
I. Insert end mill in router collet. Set router depth
of cut to match depth of plug plus thickness of wafer
(kit) (View C, figure 5-17M). Rout out core. First rout a
complete circle, following inside circle A. Then rout out
remainder of core moving router in chordwise direction
(View D, figure 5-17M.
5-30AN
Cleaning solvent is flammable and toxic.
Provide
adequate
ventilation.
Avoid
prolonged breathing of vapors and contact
with skin or eyes.
Care shall be taken to prevent MEK from
entering core area of blade. Spillage shall be
avoided.
MEK can damage leading edge
erosion guard.
Surfaces to be bonded must be clean, dry,
and free of finger prints and all foreign matter.
q. Dampen clean cheesecloth (kit) with MEK
(C87) and clean skin inside masked area. Also, clean
both sides of wafer (kit) and underside of plug patch
flange. Wipe with clean,
dry cheesecloth before
dampness evaporates.
Adhesive contains toxic ingredients. Provide
adequate ventilation and protect the skin and
eyes from contact with uncured resins or
curing agent. Wash off uncured resins and
curing agent from skin with warm water and
soap. Avoid use of solvents for cleaning the
skin.
NOTE
Never mix less than a complete two- part package
of adhesive (C7A). When less than a full batch is
required, mix the full batch and then discard the
excess after the repair is completed.
Change 49
TM 55-1520-234-23
Pot life of adhesive (C7A) is 15 minutes at 75
degrees F (23.8 degrees C). It is shorter at
higher temperatures. Always check package
dates to make sure the adhesive life limit of 1
year is not exceeded. Work without delay. Pot
life of adhesive (C17) is approximately 1/2 to 1
hour at 75 degrees F (23.8 degrees C).
r. Mix adhesive (C7A) per manufacturer's
instructions. Stir with wooden spatula until color is
uniform and all streaks have disappeared. Repeat if
more than one package is needed. Transfer adhesive to
plastic coated paper cups.
s. Using clean one inch brush (kit), apply a
liberal coat of adhesive to one side of wafer (kit) (View
D, figure 5-17M).
t. If repair is on top of blade,
routed cavity with adhesive side down.
place wafer in
u. If blade is installed on helicopter and repair is
on bottom of blade, place adhesive side of wafer
against plug (kit) with open ends of plug core up.
Cleaning solvent is flammable and toxic.
Provide adequate ventilation.
Avoid
prolonged breathing of vapors and con- tact
with skin or eyes.
Care shall be taken to prevent MEK from
entering core area of blade. Spillage shall be
avoided. MEK can damage leading edge
erosion guard.
y. Using clean cheesecloth (kit) dampened with
MEK (C87), temporarily life edges of peel- ply and wipe
off excess adhesive.
z. Place two long pieces of masking tape at right
angles,
centered over the patch spanwise and
chordwise and extending beyond the dimensions of the
blade repair fixture bladder.
aa. Install blade repair fixture (figure 5-17L).
Adhesive should not be packed into cells of
blade core or plug patch. Excess adhesive
can cause blade to be out of balance.
v. Using spatula or brush (kit), apply a liberal
coat of adhesive to walls of cavity in blade core.
w. Using brush (kit),
adhesive to:
(1)
apply a light coat of
Blade skin in masked off area around core
cavity.
(2)
Plug patch flange surrounding plug.
(3)
Outside diameter of plug.
(4)
Second side of wafer.
(1) Install from trailing edge side of blade only.
(2) Center bladder over repair area and secure.
(3) Center pad opposite bladder and secure.
Tightening of locking knobs so that metal
skirt around bladder is closer than 0.125 inch
to blade can damage blade.
(4) Tighten fixture channel locking knobs until
metal skirt around bladder is approximately 0.125 inch
from blade skin.
(5) Actuate hand pump to obtain 4 psi minimum
reading on pressure gage. Disconnect pump hose
clamp from air valve.
x. Position plug patch in cavity with stenciled
arrow pointing outboard (spanwise) and press firmly into
place. Use light hand pressure to squeeze patch area
overlapping blade skin to expel excess adhesive and air
bubbles.
Change 49
NOTE
During curing, it may be necessary to
periodically reconnect hose, and to actuate
pump to maintain 4 psi minimum.
5-30AP
TM 55-1520-234-23
(6) Connect 110 volt ac electrical power for
curing time shown in table 5-1D.
(7) At end of curing time, disconnect electrical
power, and relieve air pressure by lifting center portion
of relief valve.
ab.
Remove repair fixture from blade.
ac.
Refinish repair area.
Remove only red/purple colored filler. Do
not chip toward skin surface as damage to
bonded skin may result. Do not penetrate
the gold/brown substrate under the
red/purple filler.
d. Carefully enlarge damaged area at deepest
point of damage, exposing gold/brown substrate.
(1) Remove peel-ply and masking tape from
blade.
e. Inspect gold/brown substrate for cracks. If
there are cracks in gold/brown substrate, blade is not
repairable. If there are no cracks, proceed as follows.
Sanding skin fibers can weaken blade skin.
f.
(C17).
(2) Using 220 grit abrasive paper (kit), feather
edge of adhesive squeeze out around plug patch.
(3) Paint repaired area in accordance with
paragraph 5-5AC.
Obtain adhesive package (C7A) or alternate
g. Remove paint from surface of damaged area
by hand abrading with 220 grit abrasive paper.
ad. Adjust blade balance weights as required by
figure 5-17H.
ae. K747 blade repairs are required to be logged
in DA Form 2408-13 and -16. A permanent record must
be maintained to determine the minimum spacing
requirement between repairs. Once a repair has been
made, it is not possible to determine which type of
repair has been applied.
5-5M. REPAIR OF TRAILING EDGE
FILLED AREAS - K747 MAIN
ROTOR BLADES (AVIM).
a. This repair is for cracks, chips or missing
pieces of trailing edge filler substance which has been
applied to the trailing edge between stations 48.0 and
66.0.
b. Determine depth of damage at deepest point.
If depth of damage is less than 0.150 inch, no repair is
necessary. If depth of damage is greater than 0.150
inch but less than 0.250 inch, proceed to step f. If
depth of damage is greater than 0.250 inch, proceed as
follows.
c. Remove paint from surface of damaged area
by hand abrading with 220 grit abrasive paper to expose
red/purple filler and clear skin bonding resin.
5-30AQ
Adhesive contains toxic ingredients. Provide
adequate ventilation and protect the skin
and eyes from contact with uncured resins
or curing agent. Wash off uncured resins
and curing agent from skin with warm water
and soap. Avoid use of solvents for cleaning the skin.
NOTE
Never mix less than a complete two- part
package of adhesive (C7A). When less than a
full batch is required, mix the full batch and
then discard the excess after the repair is
completed.
Pot life of adhesive (C7A) is 15 minutes at 75
degrees F (23.8 degrees C). It is shorter at
higher temperatures. Always check package
dates to make sure the adhesive life limit of 1
year is not exceeded. Work without delay. Pot
life of adhesive (C17) is approximately 1/2 to 1
hour at 75 degrees F (23.8 degrees C).
h. Mix adhesive (C7A) per manufacturer's instructions. Stir with wooden spatula until color is
uniform and all streaks have disappeared. Transfer
adhesive to plastic coated paper cup.
Change 49
TM 55-1520-234-23
Figure 5-17P. Application of Trailing Edge Doubler Patch (K747 Blade)
Change 67
5-30AR
TM 55-1520-234-23
i. Using a wooden spatula, fill the damaged area
with adhesive.
j. Allow adhesive to cure at room temperature for
8 hours.
k. Use 220 grit abrasive paper and hand abrade
adhesive to the contour of the trailing edge.
l. Refinish repair
paragraph 5-5AC.
area
in
accordance
e. Starting with 120 grit and finishing with 220
grit abrasive paper (kit), sand the paint and the yellow
primer from the blade from the area within the guide
lines on both sides of blade and along trailing edge.
Sand only until yellow color is removed. Also sand off
any material that may be raised above the normal
contour of the blade at edges of damage. Do not sand
undamaged skin fibers (View A, figure 5-17P).
with
5-5N. APPLICATION OF TRAILING EDGE
DOUBLER PATCH - K747 MAIN
ROTOR BLADES (AVIM).
Cleaning solvent is flammable and toxic.
Provide adequate ventilation. Avoid prolonged
breathing of vapors and con- tact with skin or
eyes.
Blade must contain sufficient balance
weight to permit adjustment of blade
balance after repair. Refer to paragraph 55G before starting any repair.
a. Position blade for access to damaged area.
(See figure 5-17P.)
b. Support blade to prevent movement and
droop.
c. Obtain trailing edge doubler patch repair kit
P/N K747-201-113,
and adhesive package (C7A).
Adhesive (C17) may be used as an alternate.
Grease or lead pencils will not be used. Only
ball point pens will be used to make lines as
shown. Marks other than those specified can
weaken the repair.
d. Place the template (kit) on the blade,
centering it spanwise over the damage. Hold the
template from slip- ping and draw a line around the
template on both the top and bottom of the blade (View
A, figure 5-17P).
Spillage of MEK shall be avoided. MEK can
damage leading edge erosion guard.
f. Put on cotton gloves (kit), then plastic gloves
(kit). Leave on until completion of step o. Dampen
cheesecloth (kit) with MEK (C87).
g. Wipe off all cuttings, sanding dust, etc., from
repair area.
h. Use template to redraw guide lines (View A,
figure 5-17P).
i. Cut lengths of masking tape (kit) and mask
around the outside of the guide lines (View B, figure 517P).
Cleaning solvent is flammable and toxic.
Provide adequate ventilation. Avoid prolonged
breathing of vapors and con- tact with skin or
eyes.
Spillage of MEK shall be avoided. MEK can
damage leading edge erosion guard.
Excessive sanding will weaken blade skin.
only until yellow color is removed.
5-30AS
Sand
Surfaces to be bonded must be clean, dry and
free of finger prints and all foreign matter.
Change 67
TM 55-1520-234-23
j. Dampen clean cheesecloth (kit) with MEK
(C87) and clean skin inside masked area. Wipe with
clean dry cheesecloth before dampness evaporates.
prolonged breathing of vapors and contact
with skin or eyes.
Spillage of MEK shall be avoided. MEK can
damage leading edge erosion guard.
Adhesive contains toxic ingredients. Provide
adequate ventilation and protect the skin and
eyes from contact with uncured resins or
curing agent. Wash off uncured resins and
curing agent from skin with warm water and
soap. Avoid use of solvents for clean- ing
the skin.
n. Using clean cheesecloth (kit) dampened with
MEK (C87), temporarily lift edges of peel- ply and wipe
off excess adhesive.
o. Place masking tape over edges of patch to
prevent movement of patch. p. Install blade repair
fixture (figure 5-17L).
NOTE
Never mix less than a complete two- part
package of adhesive (C7A). When less than a
full batch is required, mix the full batch and
then discard the excess after the repair is
completed.
Pot life of adhesive (C7A) is 15 minutes at 75
degrees F (23.8 degrees C). It is shorter at
higher temperatures. Always check package
dates to make sure the adhesive life limit of 1
year is not exceeded. Work without delay. Pot
life of adhesive (C17) is approximately 1/2 to
1 hour at 75 degrees F (23.8 degrees C).
k. Mix adhesive (C7A) per manufacturer's
instructions. Stir with wooded spatula until color is
uniform and all streaks have disappeared. Transfer to
plastic coated paper cup.
I. Using clean one inch brush (kit), apply a light
coat of adhesive to inside surfaces of doubler patch
(View B, figure 5-17P) and to skin.
m. Center doubler patch within guide lines and
press into place. Slide patch back and forth slightly
under hand pressure to even adhesive. Push patch
firmly against trailing edge and center within guide lines.
Use light hand pressure to squeeze the patch from the
center to edges to work out any air bubbles.
(1) Install from trailing edge side of blade only
with bladder side on blade upper surface.
(2) Position bladder over repair area and secure.
(3) Center pad opposite bladder and secure.
Tightening of locking knobs so that metal
skirt around bladder is closer than 0.125 inch
to blade can damage blade.
(4) Tighten fixture channel locking knobs until
metal skirt around bladder is approximately 0.125 inch
from blade skin.
(5) Actuate hand pump to obtain 4 psi minimum
reading on pressure gage. Disconnect pump hose from
air valve.
NOTE
During curing, it may be necessary to periodically
reconnect hose and to actuate pump to maintain
4 psi minimum.
(6) Connect 110 volt ac electrical power for 30
minutes, or 2 hours if adhesive (C17) is used.
(7) At end of 30 minutes, disconnect electrical
power and relieve air pressure by lifting center portion of
relief valve.
q. Remove repair fixture from blade.
Cleaning solvent is flammable and toxic.
Provide adequate ventilation. Avoid
Change 49
5-30AT
TM 55-1520-234-23
A allowables. Fiber separations may be filled with
sealer (C116) or adhesive (C17) as desired.
r. Refinish repair area.
(1)
Remove peel-ply and masking tape from
blade.
Sanding fibers can weaken blade skin.
(2)
Using 220 grit abrasive paper (kit), feather
edge of adhesive squeeze-out around patch.
(3)
Paint repaired area in accordance with
paragraph 5-5AC.
s. Adjust blade balance weights as required
by figure 5-17H.
t. K747 blade repairs are required to be logged in
DA Form 2408-13 and -16. A permanent record must
be maintained to determine the minimum spacing
requirement between repairs.
5-5P. FIBER SEPARATION AND RESIN
CRACKS - TRAILING EDGE
SPLINE - K747 MAIN ROTOR
BLADES.
a. The trailing edge spline, located at station
49.00 to 41.00, is made up of Kevlar fibers in a matrix
of cured resin. Fiber separations may give the false
appearance of a crack. Fiber separations filled and
unfilled with resin are acceptable to the standards
specified below. (See figure 5-17Q.)
NOTE
There is no limit on length, location, or
closeness of separations in each respective
area.
b. Area A. Fiber separations not filled with resin
are acceptable to a depth of 0.060 inch. Fiber
separations with a depth of 0.060 to 0, 200 inch shall be
filled with sealer (C116) or adhesive (C17). Fibers are
oriented spanwise in this area, therefore, separations
are generally oriented spanwise, too.
c. Area B. Fiber separations not filled with resin
are acceptable to a depth of 0.025 inch. Fiber
separations with a depth greater than 0.025 inch have
penetrated into the area A type composite material and,
therefore, fall under the area
5-30AU
d. K747 blade repairs are required to be logged in DA Form 2408-13 and -16. A permanent record
must be maintained to determine the minimum spacing
requirement between repairs. Once a repair has been
made, it is not possible to determine which type of
repair has been applied.
5-5Q. REBONDING DELAMINATED
LEADING EDGE EROSION GUARD
- K747 MAIN ROTOR BLADES
(AVIM).
a. Position blade for access to delaminated
erosion guard. Support blade to prevent move- ment
and droop. (See figure 5-17R.)
b. Obtain erosion guard patch kit P/N K747-201119 and epoxy resin (C107).
Cleaning solvent is flammable and toxic.
Provide
adequate
ventilation.
Avoid
prolonged breathing of vapors and contact
with skin or eyes.
Isopropyl alcohol can damage leading edge
erosion guard. Avoid spillage.
c. Prior to cleaning both the erosion guard and
the blade surface,
peel back the erosion guard
approximately 0.5 inch to insure that total void or
delaminated area is identified for repair. Using cotton
tipped swab (kit) dipped in isopropyl alcohol (C23)
solvent, clean surfaces to be bonded.
d. Using masking tape (kit), mask blade along
trailing edge of boot to prevent squeezed- out adhesive
from coming in contact with the exposed blade surface.
e. Put on cotton gloves (kit), then plastic gloves
(kit).
Adhesive contains toxic ingredients. Provide
adequate ventilation and protect the skin and eyes
from contact
Change 49
TM 55-1520-234-23
Figure 5-17Q. Spline Repair (K747 Blade)
Change 49
5-30AV
TM 55-1520-234-2
Figure 5-17R. Rebonding Delaminated Leading Edge Erosion Guard (K747 Blade)
5-30AW
Change 49
TM 55-1520-234-23
with uncured resins or curing agent. Wash off
uncured resins and curing agent from skin
with warm water and soap. Avoid use of
solvents for cleaning the skin.
Protective equipment must be used when
performing these repairs.
Both erosion guard and blade leading edge
surfaces must be clean, dry, and free of finger
prints and foreign matter.
f. Mix 100 parts/weight of epoxy resin (C107)
with 12 parts/weight of DTA activator (C107) in a clean
glass, metal, polyethylene, or plastic coated paper
container.
NOTE
Pot life of adhesive is 15 minutes at 75 degrees
F (23.8 degrees C). It is shorter at higher
temperatures. Always check package dates to
make sure the adhesive life limit of 1 year is
not exceeded. Work without delay.
g. Using clean 0.25 inch brush (kit), apply a
light coat of adhesive (C107) to both surfaces to be
bonded.
5-5R. REPAIR OF ESTANE EROSION GUARD
— K747-003-205/-309 MAIN ROTOR
BLADES (AVIM1).
a. This repair is for station 213.5 inboard,
however, may be used outboard of station 213.5 (on
estane material only) when
time requirements
dictate a need for quick repair. Obtain leading edge
erosion guard patch kit P/N K747-201-119.
Excessive heat will cause the estane to
loose its properties and bulge. Use care to
keep at low heat.
b. The cuts and nicks are repaired by softening
the guard with sealing iron (T63).
c. Set the temperature at the minimum. Use just
enough heat to cause the estane guard material to be
soft. Keep the iron moving.
d. Small damage can be repaired by moving the
material from each side of the damage with the sealing
iron.
e. A 0.25 inch damage can be repaired by
adding slivers of estane guard (kit).
h. Using finger pressure, press erosion guard
to blade while working out excess adhesive from under
the erosion guard. Wipe away excess adhesive with
clean cheesecloth (kit) to pre- vent adhesive from
running off the masking tape onto the exposed blade
surface.
Cleaning solvent is flammable and toxic.
Provide adequate ventilation.
Avoid
prolonged breathing of vapors and con- tact
with skin or eyes.
i.
Lay teflon parting blanket (kit) over repair.
Place masking tape (kit) over edges of parting blanket to
prevent movement.
f. Using a clean 1-inch brush, apply one coat of
MEK (C87(. Do not overbrush the same area more than
three times.
j. Obtain two wooden blocks approximately 0.75
x 2 x 6 inches and a C clamp (8 inch opening by 6
inches deep). Place 0.25 x 2 x 6 inches strip of rubber
between block and parting blanket. Place remaining
block and rubber strip on opposite surface and, using C
clamp, apply light pressure to rebonded area.
g. Allow to air dry for a minimum of 1 hour.
k. At end of four hours, at room temperature,
remove clamp, blocks, rubber strip, parting blanket
and masking tape.
Change 65
NOTE
The coating will develop optimum durability
in approximately 6 to 8 hours. Flying in rain
conditions with less drying time may cause
rapid erosion of resurfaced area.
5-5S. REPAIR OF FLUOROCARBON
EROSION GUARD - K747 MAIN
ROTOR BLADES (AVIM).
5-30AX
a. This procedure uses kit P/N K747-207-1 for
the repair of nicks and cuts which involve less than one
square inch of damaged area in the following areas:
TM 55-1520-234-23
at each end of major axis of cut out repair area. The two
holes should be made with a 5/64 inch diameter drill bit
and be large enough to accept syringe supplied with kit
(1) K747-003-205/-309, between stations 213.5
and 260.0 where fluorocarbon leading edge erosion
guard replacement has been accomplished. Check
blade DA Form 2408-16 to determine this fact.
(5) Turn tape over (mating surface toward blade
and press into position as shown in figure 5-17S.
Ensure good tape adhesion is made. Do not press tape
mold into void.
(2) K747-003-209/-401, and -303/-403,
leading edge erosion guard surface.
any
Do not cut into the spar when cutting away
damaged leading edge erosion guard areas.
b. Remove damaged leading edge erosion guard
in area being repaired. Cut material away using a razor
blade or equivalent to form an oval area. The sides of
cut must slope inward toward center of damaged area.
(See figure 5-17S.)
Adhesive contains toxic ingredients. Provide
adequate ventilation and protect the skin
and eyes from contact with uncured resins
or curing agent. Wash off uncured resins
and curing agent from skin with warm water
and soap. Avoid use of solvents for cleaning the skin. Wear polyethylene gloves
while mixing
and injecting adhesive
material.
NOTE
Cleaning solvent is flammable and toxic.
Provide adequate ventilation.
Avoid
prolonged breathing of vapors and con- tact
with skin or eyes.
c. Wipe exposed area with a cheesecloth (C36)
dampened with MEK (C87).
d. Form a mold to contain injected adhesive.
Proceed as follows:
(1) Cut a piece of #Y8412 tape (kit) in a rectangle which is approximately 1/2 inch larger than repair
area.
(2) Cut a second piece of #Y8412 tape (kit) one
inch larger in length and width than the piece cut above.
(3) Position smaller piece of tape (mating the
adhesive surfaces) in the center of the larger one. A
1/2 inch exposed adhesive border will result. (See
figure 5-17S.)
(4) Locate tape mold centrally over damaged
area so mating surface (adhesive border) faces outward.
Poke two holes in the tape, one
5-30AY
Pot life of adhesive is approximately 15 to 20
minutes. Do not use out-of- date adhesive.
Work without delay.
e. Mix adhesive filler in plastic syringe by
following manufacturer's instructions.
f. Inject adhesive filler into one hole until it
seeps from second hole. Reverse action by injecting
into second hole until the adhesive filler seeps from first
hole. Ensure no air pockets are left beneath tape mold.
g. Cure repair using one of the following
procedures:
(1)
Three days at room temperature.
(2)
Four hours at room temperature follow- ed
by 4 hours at 160 + 10 degrees F (71 + 6 degrees C).
(Heat can be applied by lamps or a hot air gun.)
h. Cut excess adhesive protruding from injection
holes with a razor blade.
i. Remove tape mold and abrade any high spots
or excess material from erosion guard surface. Blend
patch to match contour of surrounding areas using 240
grit abrasive paper
Change 65
TM 55-1520-234-23
Figure 5-17S. Typical Repair of Fluorocarbon Erosion Guard Nicks and Cuts Using Kit, PN. K747-207
Change 49
5-30AZ
TM 55-1520-234-2
Figure 5-17T. Application of Loading Edge Erosion Guard Patch (K747 Blade)
5-30BA
Change 67
TM 55-1520-234-23
f. Using 180 or 220 grit sandpaper (kit), abrade
area of erosion guard inside guidelines and underside
surface of patch.
5-5T. APPLICATION OF LEADING EDGE
EROSION GUARD PATCHK747-003-205/-309 MAIN ROTOR
BLADES (AVIM)
a. This repair is for station 213.5 inboard, however, may be used outboard of station 213.5 (on estane
material only) when time requirements dictate a need for
quick repair.
CAUTION
Blade must contain sufficient balance
weight to permit adjustment of blade
balance after repair. Refer to paragraph 55G before starting any repair.
b. Position blade for access to damaged leading
edge erosion guard. Support blade to prevent movement and droop.
c. Obtain leading edge erosion guard patch kit
P/N K747-201-119 and estane contact cement. Prepare
estane contact cement for estane erosion guards as
follows:
(1) Obtain a piece of estane material and some
MEK (C87).
(2) Cut estane material into very small slivers
approximately 0.025 inch in size.
(3) Combine estane and MEK. Suggested mix
ratio is 15 grams of estane to 85 grams of MEK. Allow to
stand for 24 hours. Agitate intermittently throughout this
24 hour period to ensure that the estane is totally
dissolved.
(4) The solution is now ready to use and should
be agitated before any such use.
CAUTION
Grease or lead pencils will not be used. Only
ball point pens will be used to make lines as
shown. Marks other than those specified
can weaken the repair.
d. Place the template (kit) on the erosion guard
centering it over the damage. Hold template from
slipping and mark outline of template on erosion guard.
(See figure 5-17T.)
g. Put on cotton gloves (kit) and then plastic
gloves (kit). Leave on until completion of step 1.
WARNING
Cleaning solvent is flammable and toxic.
Provide adequate ventilation.
Avoid
prolonged breathing of vapors and contact
with skin or eyes.
h. Using clean cheesecloth (kit) and isopropyl
alcohol (C23) solvent, wipe surfaces to be bonded.
i. Using clean one inch brush (kit), apply a light
coat of estane contact cement within masked area of
erosion guard and to the underside surface of the patch.
Allow to air dry for five minutes.
CAUTION
Patch will adhere to erosion guard on contact. Make certain that patch is correctly
aligned before making contact.
j. Starting towards leading edge of blade, install
patch, working it carefully into place with fingers, using
extreme care not to entrap air under patch. Press all
areas of patch firmly into contact with erosion guard.
k. Remove masking tape.
l. Using one inch brush (kit), apply a medium
thick coat of estane contact cement to the patch,
extending over the edges of the patch to blend into
adjacent area of the erosion guard.
m. Adjust blade balance weights as required by
figure 5-17H.
n. Allow to air dry 12 hours.
5-5U. APPLICATION OF LEADING EDGE
EROSION GUARD-K747 MAIN
ROTOR BLADES (AVIM).
a. This repair is for use between stations 213.5
and 260.0 using kit P/N K747-206-1.
e. Cut lengths of masking tape (kit) and mask
around outside of the guide lines.
Change 71
5-30BB
TM 55-1520-234-2
Figure 5-17U. Placement of Erosion Guard Replacement Part (Kit K747-206-1) And Method for Determining
Current Boot Material And Thickness (K747 Blade Series)
5-30BC
Change 49
TM 55-1520-234-23
b. Remove the leading edge erosion guard
between stations 215.5 and 260.0.
(1)
Using a ball point pen, locate and mark
station 224.0 on leading edge of blade.
Grease or lead pencils will not be used.
Use only a ball point pen.
(2) At station 260.0 joint of forward tip cap and
blade, measure distance from leading edge of tip cap to
leading edge of blade. This dimension should be 0.234
to 0.279 inch. If dimension is not within the range,
proceed to next step. If dimension is within the range,
proceed to step d.
(1) Mark a straight line chordwise on the top,
bottom, and leading edge erosion guard surfaces at
station 215.5 for defining the cut line. (See figure 517U.)
Care must be exercised when cutting,
lifting, or peeling leading edge erosion
guard from skin, spar and filler surfaces.
Damage to the spar or graphite/doubler
components could result in scrapping of
blade.
(2) Cut with a sharp knife and lift the leading
edge erosion guard edge at station 215.5 on blade top
surface using a one inch chisel or sharpened file blank.
Peel leading edge erosion guard away in the outboard
direction, working it loose by hand with a sharpened file
blank.
(3) Repeat preceding step for blade bottom
surfaces. Observe the necessary caution.
(3) Measure distance of 0.250 inch from leading
edge of tip cap and mark blade top and bottom surface.
Draw a line on the blade top and bottom surface using a
ball point pin and a straight edge between station 224.0
leading edge and set back dimension at station 260.0.
(4) Abrade a flat vertical surface of the leading
edge filler along the two lines just drawn. Use 50 to 80
grit abrasive paper wrapped on a wooded block.
(5) Radius the flat
edge filler just abraded.
using a 5/64 radius gage.
from the existing radius
radius at station 260.0.
vertical surface of the leading
First radius at station 260.0
Then radius the leading edge
at station 224.0 to the 5/64
(6) Lay a straight edge with edge coated with
chalk, against leading edge of blade between the
station 260.0 indent and station 224.0 as marked. There
should be no gap greater than 0.010 inch between them.
To bring any gaps within tolerance, proceed as follows:
(a) For any high spots, hand abrade us- ing 50
and 80 grit abrasive paper wrapped on a wooden block
for a straight edge contour.
NOTE
A new thicker erosion guard repair part is
supplied in kit K747-206-1. Some blades may
still have a thinner original or replacement
fluorocarbon erosion part in position. It is
necessary to check each blade prior to
application of a repair part to determine if the
part being removed is the early thin or later
thick dimension. All blades having a full
estane or early fluorocarbon erosion guard,
or a thin replacement fluorocarbon outboard
guard section will require recontouring of the
blade leading edge filler material.
c. Check dimensions of blades as shown in
figure 5-17U. If it is determined that the erosion guard
just removed was one of the early thin parts, it will be
necessary to recontour the leading edge filler to
accommodate the new thicker replacement part. Do so
using the following procedure.
Change 49
(b) For any low spots, it will be necessary to fill
with leading edge filler. Refer to paragraph 5-5W for
mixing filler resin (C64A). Refer to figure 5-17U for
attaching masking tape mold. Using wooden spatula,
apply a sufficient amount of filler at low spot to do the
repair. Wrap masking tape over repair surface and let
dry for 15 minutes. Abrade as in step (a) above for a
straight edge contour.
Use protective equipment over eyes and mouth
when abrading. Be careful not to abrade or nick
spar surfaces.
5-30BD
TM 55-1520-234-23
(7) Check leading edge filler on both top and
bottom surfaces for any depressions between stations
260.0 and 224.0 and fill as in step (b) above. Using a
right angle air motor and a 3 inch x 80 grit abrasive disc,
abrade surfaces to a smooth contour.
(a)
area.
Remove any loose uralite from damaged
(b) Place bag sealant,
if needed around
damaged uralite area to act as a dam as shown in figure
5-17V.
(8) Using a vacuum or a clean, dry, oil- free
cloth, clean both blade surfaces of any abrading dust.
d. Prepare blade surfaces for application of
leading edge erosion guard part.
Care must be exercised not to abrade spar
surfaces.
Removal of paint from blade surface any
distance greater than 1/4 inch aft of erosion
guard trailing edge will destroy blade
lightning protection.
(1) Abrade faying surfaces of blade to remove
any adhesive residue left from erosion guard removal.
Use a disc sander and a 3 inch x 120 grit abrasive disc.
Grease or lead pencils will not be used. Use
only a ball point pen.
(2) Mark a straight line parallel to cut line on the
top, bottom, and leading edge erosion guard surfaces
at station 213.5 for defining the scarf line. (See figure 517U.)
(3) Scarf remaining leading edge erosion guard
from surface at station 213.5 down to spar at station
215.5. Use marks made in step (2) above as guidelines.
Check scarfed surface with a straight edge. Use a disc
sander and a 3 inch x 24 grit, then 80 grit abrasive disc.
(4) Abrade blade bottom surface to accept a 1 x
6 inch test sample. Use 120 grit abrasive paper. (See
figure 5-17X.)
(5) Restore any uralite surfaces (over weight
retention bolts) which may have been damaged during
leading edge erosion guard removal using the following
method. If no damage occurred in this area, proceed to
step (6).
5-30BE
Filler resin contains toxic ingredients.
Provide adequate ventilation and protect
skin and eyes from contact with uncured
resins or curing agent. Wash off uncured
resins or curing agent from skin with warm
water and soap. Avoid use of solvents for
cleaning skin.
NOTE
Filler resin should be thoroughly mixed until
all streaks in the mixture are eliminated. Do
not whip any air into mixture.
Blade may be tilted to level resin in the dam
around the repair area.
(c) Mix enough potting resin (C98B) part A 100
parts/weight with part B 40 parts/weight to restore area
being repaired.
(d) Pour resin into and fill area dammed by bag
sealant (if used) or wipe in place with a squeegee.
(e) Allow resin to cure for a minimum of 12
hours at 75 degrees F (23.8 degrees C).
Do not abrade any surface surrounding
uralite potting resin.
Damage would be
inflicted that would require depot repair.
(f) Remove bag sealant dam surrounding the
cast resin (if used). Abrade resin flush with blade
contour using a 3 inch x 24 grit then 80 grit sanding disc.
Cleaning solvent is flammable and toxic.
Provide
adequate
ventilation.
Avoid
prolonged breathing of vapors and contact
with skin or eyes.
Change 49
TM 55-1520-234-23
Figure 5-17V. K747 Blade Uralite Repair
Change 49
5-30BF
TM 55-1520-234-23
(6) Wipe any foreign material from blade repair
area using a cheesecloth (C36) dampened with solvent
(C142). Repeat this step two more times, wiping
solvent dry before it evaporates. Allow 15 minutes for
solvent evaporation.
(7) Wipe replacement erosion guard on mating
and outside surfaces with MEK (C87). Repeat this step
two more times,
wiping solvent dry before it
evaporates. Allow 15 minutes to dry.
(8) Apply 2 inch wide nylon tape to both top and
bottom blade surfaces as shown in figure 5-17W, detail
A. The tape should be 0.250 inch aft of erosion guard
trailing edge on blade top and bottom surfaces. This will
provide an area to include a test specimen under
vacuum bagging on blade bottom surface as shown in
figures 5-17W, detail A, and 5-17X.
NOTE
Do not remove paper backing on bag sealant
or double faced tape.
Keep adhesive
bonding surfaces clean.
(9) Apply bag sealant and tabs of double faced
tape (C135). Secure vacuum hoses and vacuum gage
as shown in figure 5-17W, detail B. A vacuum hose
should be applied to both blade sides.
(10) Pre-fit 2.50 inch wide bleeder cloth to the
blade as shown in figure 5-17W, detail C, and remove.
5-17X.
(e) Remove erosion guard and the masking
tape securing it.
(12)
Abrade mating surfaces of replacement
erosion guard and test specimen as follows:
Abrasion should be done on smooth, hard
surface at low rpm.
(a) Abrade inboard end of replacement erosion
guard part that will be fitted between stations 213.5 and
215.5. Taper guard material from stations 215.5 to
213.5 to fit scarfed surface of existing guard. Use a disc
sander and 3 inch x 60 grit abrasive disc.
Care must be exercised not to rip or gouge
material when abrading replace- ment guard
part.
(b) Hand abrade remaining replacement part
mating surface. Use 180 grit abrasive paper.
(c) Abrade mating surface of test specimen.
Use a disc sander and a 3 inch x 80 grit abrasive disc.
(11) Pre-fit replacement erosion guard part to blade.
(a) Trim extreme ends of replacement guard at
flash lines.
NOTE
Guard material may be slightly stretched to
fit.
(b) Fit erosion guard to blade leading edge.
Temporarily secure in position with strips of masking
tape. (See figure 5-17W, detail D.)
(c) Trim extreme ends of replacement guard
part at station 213.5 and then at station 260.0 to fit, if
necessary.
(d) Apply a wrap of 1 inch wide nylon tape to one
end of 1 x 6 inch test specimen as shown in figure
5-30BG
Cleaning solvent is flammable and toxic.
Provide adequate ventilation. Avoid prolonged
breathing of vapors and contact with skin or
eyes.
(13)
Wipe replacement erosion guard on
mating and outside surfaces with MEK (C87). Repeat
this step two more times. Wipe solvent dry before it
evaporates. Allow 15 minutes to dry.
(14)
Wipe blade repair area surfaces with
cheesecloth (C36) dampened with solvent (C142).
Repeat this step two more times, wiping solvent dry
before it evaporates. Allow 15 minutes to dry.
Change 49
TM 55-1520-234-2
Figure 5-17W. Application of Vacuum Bagging Materials and Placement of Erosion Guard (Sheet 1 of 4)
Change 49
5-30BH
TM 55-1520-234-2
Figure 5-17W. Application of Vacuum Bagging Materials and Placement of Erosion Guard (Sheet 2 of 4)
5-30BJ
Change 49
TM 55-1520-234-23
Figure 5-17W. Application of Vacuum Bagging Materials and Placement of Erosion Guard (Sheet 3 of 4)
Change 49
5-30BK
TM 55-1520-234-23
Figure 5-17W. Application of Vacuum Bagging Materials and Placement of Erosion Guard (Sheet 4 of 4)
5-30BL
Change 49
TM 55-1520-234-23
Figure 5-17X. Repair Parts and Specimen Orientation (Kit, P/N K747-206)
Change 65
5-30BM
TM 55-1520-234-23
(2)
Using a 3 inch paint roller, apply a uniform
coat of adhesive to the glass cloth just positioned.
NOTE
Wear clean white gloves from this point in
repair until the vacuum bagging material is in
place.
This will aid in preventing
contamination of bond surface preparation.
(15) Inspect repair for cleanliness and surface
preparation.
g. Install leading edge erosion guard and test
specimen in position on blade.
(1) Using a 3 inch paint roller, apply a uniform
light coat of adhesive to mating surface of the
replacement guard.
(2) Apply a light coat of adhesive to mating
surface of test specimen.
Adhesive contains toxic ingredients. Provide
adequate ventilation and protect the skin and
eyes from contact with uncured resins or
curing agent. DTA can cause blindness and
burns. Wash off uncured resins and curing
agent from skin with warm water and soap.
Avoid use of solvents for clean- ing the skin.
Wear polyethylene gloves over cotton gloves
for this task.
NOTE
Pot life of adhesive is approximately 30
minutes at 75 degrees F (23.8 degrees C). It is
shorter at higher temperatures. Always check
package dates to make sure the adhesive life
limit is not exceeded. Work without delay.
Record time at which adhesive was mixed as
an aid in determining pot life.
(4) Position guard replacement on the blade and
hand work it around the leading edge chordwise for full
span length of the part. It may be necessary to stretch
replacement part for a good spanwise fit.
(5) Remove any adhesive on the outside surface
of the leading edge erosion guard with solvent (C 142).
(6) Temporarily tape leading edge guard and
test specimen in place with masking tape. Apply 1 inch
wide nylon tape to hold part and test specimen in
position. Remove masking tape carefully. (See figure
5-17W, detail D.)
h. Vacuum bag repair in accordance with the
following procedures after reviewing figures 5-17W and
5-17X.
(1) Remove backing of double faced tape and
bag sealant.
e. Mix and apply adhesive(C13A).
(1) Mix 100 parts/weight of epoxy resin (Epon
826) with 10 parts/weight of Versamid 125 and 6
parts/weight DTA activator. Stir until streaks disappear.
Do not induce air bubbles while stirring.
(2) Apply a uniform light coat of adhesive to
blade bottom surface, leading edge, and top surface.
Use a 3 inch paint roller for adhesive application.
f. Apply the #120 glass cloth,
guard, and test specimen to blade.
(3) Place test specimen in position on blade
bottom surface in areas shown in figure 5-17X.
leading edge
(1) Place #120 glass cloth in position as shown
in figure 5-17X. Smooth into place by hand. Using
scissors, trim glass cloth to fit exactly.
5-30BN
(2) Apply bleeder cloth as shown in figure 517W, detail C.
(3) Install vacuum bagging and
position on bag sealant. Smooth out any
bagging material. Keep tension on bagging
it is pulled into position. Ensure a good seal.
5-17W, detail D.)
press into
wrinkles in
material as
(See figure
(4) Apply vacuum to vacuum hose and obtain
20 inches Hg vacuum under bagging.
(5) Coat bagging material with petrolatum (C96)
to act as a lubricant for the squeegee.
Change 49
TM 55-1520-234-23
Figure 5-17Y. Vacuum Bagging For Installation of Erosion Guard Repair Kit, PN K747-206
Change 49
5-30BP
TM 55-1520-234-23
(6) Squeegee leading edge erosion guard along
leading edge and test specimen. Roll any excess
adhesive outboard toward bleeder cloth eliminating any
air pockets under leading edge erosion guard. (See
figure 5-17Y.) Do top surface first, then bottom.
(7) Inspect leading edge erosion guard surface
for any soft spots including the test specimen. Soft
spots are an indication of air pockets. Rework bagging
using a squeegee to remove air pockets, as required,
by doing bottom surface, then top.
(8) Check for any vacuum leaks by press- ing
vacuum bagging firmly against vacuum sealant and
eliminating any leaking bag creases.
i.
Cure repair using one of three following time
temperature sequences.
(1) Room temperature (75 degrees F (23.8
degrees C)) for 24 hours minimum. Vacuum bagging,
bleeder and masking material may be removed from
blade after 12 hours (optional 24 hours).
Do not heat any estane leading edge erosion
guard material if it exists (in- board of
station 213.5) above 140 degrees F (60
degrees C).
(2) Room temperature (75 degrees F (23.8
degrees C)) for 16 hours minimum followed by 130 + 10
degrees F (54.4 + 6 degrees C) for 2 hours minimum.
Remove bagging, bleeder and masking materials.
exercised not to frictionally heat the guard
material. Use short, light, quick strokes
with the sanding tools.
l. Feather peripheral seams of repair area and
test specimen. Use a disc sander and 3 inch x 80 grit
abrasive disc. Follow with 120 grit if necessary. Do not
feather trailing edge of leading edge erosion guard.
NOTE
A peel test will be performed 24 hours after
the cure cycle is completed. Do not exceed
limits of fish scale during the peel test.
m. Perform a peel test on test specimen attached to blade surface.
(1) Securely attach a C clamp to taped end of
test specimen.
(2) Attach a fish scale (capable of measuring 15
or 20 pounds) to C clamp.
(3) Pull specimen back across its longitudinal
axis. Record on component DA Form 2408-16, the
amount of force measured on the scale to peel
specimen from blade. The minimum accepted peek
strength is 6 pounds which ensures a good adhesive
bond. This step will be verified by a QA/QC inspector.
(4) If peel test results are less than 6 pounds,
the bond is unacceptable. The new leading edge
erosion guard must be replaced by repeating this
procedure.
(3) 130 ±10 degrees F (54.4±6 degrees C) for 4
hours minimum.
Remove bagging,
bleeder and
masking materials.
n. Remove the test specimen (if it did not come
off during peel test) and any residue left by test
specimen. Use a disc sander with 3 inch x 80 grit
abrasive disc. Follow with 120 grit if necessary.
j. Remove all vacuum bagging materials from
blade. Use care not to disturb test specimen.
o. Restore blade finish in accordance with
paragraph 5-5AC.
k. Remove any excess adhesive from leading
edge erosion guard and blade areas using putty knife,
solvent (C142), and scotch brite (C113).
5-5V. BONDING OF LEADING EDGE EROSION
GUARD STATION 213.5 OUTBOARD K747-003-303/-403 AND -303/-403 FIELD I
MODIFIED MAIN ROTOR BLADES
(AVIM).
When feathering chordwise seam at station
213.5, extreme caution must be
5-30BQ
a. Make up six bonding intensifiers by attaching
together two AN960-1416 washers, one
Change 65
TM 55-1520-234-23
AN970-5 washer, and one AN960-1416 washer in that
sequence. Attach by spot welding or using a suitable
adhesive.
When using the squeegee to eliminate any
air pockets under leading edge erosion
guard, extra care should be exercised to
press erosion guard securely around
circumferences of posts,
both blade
surfaces.
b.
Install the fluorocarbon erosion guard in
accordance with paragraph 5-5U.a thru h. Be sure to
cover both ends of posts with circular tabs of masking
tape to prevent entry of adhesive into post holes.
c.
After vacuum bagging has been completed,
install the six intensifiers, made in step a above.
(1) Position an intensifier, with the two AN9601416 washer side towards blade surface, over opposite
post humps on both blade surfaces. Make sure the two
intensifiers are seated evenly around humps. Install a C
clamp around them and tighten clamp only until
intensifiers are bottomed around post circumferences.
Repeat for other two posts.
(2)
Squeegee area around post to remove any
air bubbles that may have formed when securing
clamps.
d. Cure repair in accordance with paragraph 55U.i. Remove clamps,
intensifiers,
and vacuum
bagging materials.
e. Cut holes in fluorocarbon erosion guard over
ends of posts, both blade surfaces, as follows:
(1) Manufacture a 2 x 6 inch hole cutting
template out of any suitable material. Drill three 1 inch
diameter holes with their centers spaced 2 inches apart.
(2) Place template over post hole area with
template holes seated evenly around the humps. Scribe
circles at humps using a lead pencil.
(3) With template removed and using a utility
knife with a sharp cutting edge,
cut around
circumference of marked circles with cutting plane of
blade slanted towards posts. Ensure that erosion guard
and fiberglass layer is completely severed and pry off
the rubber plugs. Remove masking tape over ends of
posts.
Change 65
(4) Using a right angle air motor with a 1-1/2
inch x 80 grit abrasive disc, chamfer edge of rubber
around holes.
(5) Repeat steps (2) thru (4) above for other
side of blade.
f.
Proceed with paragraph 5-5U.k thru p.
5-5W. REPAIR OF DAMAGED PORTION OF
LEADING EDGE FILLER - K747 MAIN
ROTOR BLADES (AVIM).
NOTE
Repair of leading edge filler is limited to replacement of the filler, not to exceed 6
inches in length and only one repair per
blade.
a. This repair is to be used for damage between
stations 224.4 and 260.0.
b. Position blade, top surface up, in fixed blade
rack.
Provide adequate ventilation to remove any
vapor concentrations in the area.
Wear polyethylene gloves over cotton
gloves to prevent skin contact with solvent
and eliminate possibility of getting skin oils
on blade surfaces.
Wear splash-proof goggles when working
with solvent.
c. Remove any contamination by wiping repair
area with a cheesecloth (C36) dampened with solvent
(C142). Repeat this step three times, wiping solvent dry
before it evaporates.
d. Apply several strips of 2 inch wide masking
tape (to form a mold) to blade bottom surface. (See
figure 5-17Z.) Do not attach tape to blade top surface.
NOTE
Both parts of two-part urethane filler must
be at room temperature (70 degrees F (21
degrees C)) before use.
Two - part urethane filler must be
5-30BR
TM 55-1520-234-23
Figure 5-17Z. Improvised Mold For Casting A Small Section Of Leading Edge Filler
5-30BS
Change 49
TM 55-1520-234-23
k. Refinish by painting affected areas in
accordance with paragraph 5-5AC.
thoroughly mixed in order to preserve repair
strength.
l. K747 blade repairs are required to be logged
in DA Form 2408-13 and -16. A permanent record must
be maintained to determine the minimum spacing
requirement between repairs. Once a repair has been
made it is not possible to determine which type of repair
has been applied.
Surfaces to be bonded must be clean, dry,
and free of finger prints and all foreign
matter.
The resin compound used in casting leading
edge filler is fast setting lap- proximately 15
minutes between mix- ing and gelling).
Work without delay.
5-5X. REMOVAL OF STAINLESS STEEL
EROSION GUARD - K747-003-303/-403
MAIN ROTOR BLADE.
e. Mix filler resin (C64A).
(1) Stir base and catalyst components in their
respective supply containers until thoroughly mixed.
a. Inspect erosion guard for damage. Refer to
table 5-1B for limits.
(2) Weigh out equal amounts by weight (as
required) of catalyst and base material in separate paper
cups.
b. Remove six screws (3,
figure 5-17AA)
securing erosion guard (1) to blade posts (4), us- ing a
No. 4 Philips screwdriver. Remove two screws (2)
securing erosion guard to forward tip cap, using a
standard Philips screwdriver. Retain screws (2 and 3)
for later use.
(3) Combine catalyst into base and stir with a
wooden spatula. Ensure a thorough mixture is obtained.
f. Pour filler into the masking tape mold applied to blade surface. Fold masking tape over blade
leading edge and attach to blade top surface. (See
figure 5-17Z.)
g. Remove masking tape mold after 15 to 20
minutes. Rough sand the filler, removing high spots
and excess material. Use right angle air motor and 3
inch x 60 grit adhesive disc.
NOTE
To prevent scoring of primer coating on
inside surface of stainless steel erosion
guard and surface of fluorocarbon erosion
guard when prying them apart, file a radius
on working edges and corners of putty knife.
NOTE
It is advisable to wear a leather glove to
cushion the palm for this task.
A straight edge used as a guide and flat
wood block covered by 50 grit abrasive
paper provides a means of obtaining
finished
contour
and
leading
edge
acceptable limits.
c. Insert putty knife under lip of stainless steel
erosion guard, with rounded underside edge against
fluorocarbon erosion guard. Start- ing at outboard end
of guard and working towards inboard end, apply a firm
wedging force to pry stainless steel guard loose from
sealant.
h. Fine sand filler to match blade outline and
contour. Place a straight edge across the leading edge
to gage leading edge outline. Use a wooden block
(approximately 10 to 12 inches in length) covered with
50 grit abrasive paper.
d. Turn blade over in rack and repeat step b
above for loosening stainless steel guard on underside
of blade. Remove guard when completely broken free
of sealant.
i. Repeat previous task for both top and bottom filler surfaces. Always use the straight edge as a
guide in obtaining proper contour.
e. The sealer residue can be removed from
fluorocarbon erosion guard by rubbing with the open
palm or a cloth. Lightly abrade both top and bottom
surfaces of fluorocarbon erosion
j. Install leading edge
accordance with paragraph 5-5U.
erosion
guard
in
Change 65
5-30BT
TM 55-1520-234-23
Figure 5-17AA. Repair Parts Orientation Use for Removal and Installation of Stainless Steel erosion GuardK747-003-303/-403 and-303/-403 Filed Modified Blades
5-30BU
Change 65
TM 55-1520-234-23
guard with a layer of screen cloth (C44B) wrapped
around a wooden block only to give erosion guard
surface a dull finish.
f. Remove any foreign material and residue
from blade surfaces. Use a vacuum or clean, dry, oilfree cloth.
Cleaning solvent is flammable and toxic.
Provide adequate ventilation.
Avoid
prolonged breathing of vapors and con- tact
with skin or eyes.
g. If the stainless steel erosion guard is to be
reinstalled, lightly wipe inside surface with cheesecloth
dampened with MEK (C87) to remove sealant residue.
5-5Y. INSTALLATION OF STAINLESS STEEL
EROSION GUARD - K747-003-303/-403
MAIN ROTOR BLADES (AVIM).
a. Prepare blade for installation of stainless steel
erosion guard.
(1) Locate stainless steel erosion guard on
fluorocarbon erosion guard to align mounting holes.
Install screws (2 and 3, figure 5-17Z) and tighten until
they bottom out.
(2) Mask erosion guard area using one inch
nylon tape (C135A) positioned 1/4 inch from periphery
of stainless steel erosion guard, both surfaces of blade.
(See figure 5-17AB.)
(3) Remove the eight self-locking screws from
stainless steel erosion guard and remove guard from
blade. Discard screws.
Cleaning solvent is flammable and toxic.
Provide adequate ventilation.
Avoid
prolonged breathing of vapors and con- tact
with skin or eyes.
(4) Wipe both surfaces of fluorocarbon erosion
guard with cheesecloth (C36) dampened with MEK
(C87) to remove contamination in area framed by nylon
tape.
Change 65
Surfaces to be bonded must be clean, dry,
and free of finger prints and all foreign
matter. Wear cotton gloves when handling
these surfaces.
b. Prepare the replacement stainless steel
erosion guard for installation by masking out- side
surface using masking tape (C134) to protect the
surface from sealant. Using a sharp knife, cut out the
eight mounting holes in tape.
Sealant contains toxic ingredients. Provide
adequate ventilation to remove any vapor
concentrations
in
the
area.
Wear
polyethylene gloves over cotton gloves to
prevent
skin
contact
with
sealant
ingredients and eliminate possibility of
getting skin oils on repair area surfaces.
Wear splash-proof goggles when working
with sealant.
NOTE
Pot life of sealant is 30 minutes at 75 degrees
F (23.8 degrees C). It is shorter at higher
temperatures. Work without delay.
c. Prepare Proseal 890 (C105).
Mix 1
part/weight of accelerator to 10 parts/weight of base in a
plastic coated paper container. Use a wooden stirrer
and mix sealant to a uniform color and consistency.
d. Apply sealant with a stiff-bristled brush, to
inner surface of stainless steel erosion guard and
framed area of fluorocarbon erosion guard. Apply only
to area shown in figure 5-17AC.
NOTE
Exercise care to prevent sealant from entering
threaded area of post holes.
e. Position stainless steel erosion guard over
fluorocarbon erosion guard, until mounting holes are
aligned.
5-30BV
TM 55-1520-234-23
Figure 5-17AB. Preparation of K747-003-303/-403 Blade for Application of Sealant
5-30BW
Change 65
TM 55-1520-234-23
The self-locking screws can be used on
once. New screws are required for each
installation.
f. Install six NAS1189E5P8B screws (3, figure 517AA) in blade posts (4) and two NAS1189E4P6B screws
(2) in forward tip cap, using a No. 4 Philips screwdriver
and a standard Philips screwdriver, respectively. Hand
tighten all screws to achieve clamping of stainless steel
erosion guard. QA/QC inspection is required.
g. Manufacturer a holding fixture out of plywood
stock and install over stainless steel guard as shown in
figure 5-17AD. The strips of industrial tape around the
three leading edge contour block details and trailing edge
of blade are to be wrapped tight enough to seat blocks
evenly and securely against leading edge contour of
guard. Tighten the six 6 inch C clamps sufficiently to seat
the two spanwise details securely against top and bottom
surfaces of guard.
h. Cure Proseal 890 (C105) at room temperature (70
degrees F (21 degrees C)) for 24 hours. An alternate
method is to apply heat using a sufficient number of
infrared lamps to provide 140 to 180 degrees F for one
hour.
i. Upon completion of cure, remove holding fixture
and tape from blade surfaces.
j. Remove masking tape from stainless steel
erosion guard and feather sealant at trailing edge of guard
using a layer of screen cloth (C44B) wrapped around a
wooden block.
5-5Z. CHANGING K747-003-303/-403 BLADE TO-303/403 FIELD MODIFIED BLADE (AVIM).
a. Remove the stainless steel erosion guard in
accordance with paragraph 5-5X.
The shields and
installation instructions which come with the-303 blade are
always retained for use when converting.
The self-locking screws can be used only
once. New screws are required for each
installation.
b. Insert three K747-209-11 shields (5, figure 5-17Z)
into the fluorocarbon erosion guard holes on the top
blade surface. Install three new NAS1189E5P8B screws
(3) into the blade posts (4) using a No. 4 Philips
screwdriver. In- stall one new NAS1189E4P6B screw (2)
into the forward tip cap using a standard Philips
screwdriver. The four screws are to be hand tightened
only. Turn blade over in rack and repeat procedure for
bottom blade surface.
5-5AA. CHANGING
K747-003-303/-403
MODIFIED BLADE TO-303/-403 BLADE (AVIM).
FIELD
a. Remove screws (2 and 3, figure 5-17Z) and six
shields (5) from their locations on top and bottom blade
surfaces. Discard screws.
b. Install K747-210-11 stainless steel erosion guard
(1) in accordance with paragraph 5-5Y.
c. Retain shields and installation instructions for use
during any future conversion to K747-003-303/-403 field I
modified blade.
5-5AB. REPLACING SECTIONS OF EROSION GUARD
- K747-003-205/-309 MAIN ROTOR BLADE (AVIM).
a. This repair is for station 213.5 inboard. Obtain kit
P/N K747-201-119.
b. Position blade for access to damaged leading
edge erosion guard. Support blade to prevent movement
and droop.
c. Using sharp knife, remove all damage from
guard including separated guard. Cut the guard in such a
pattern that can be duplicated with a like patch (circle,
square, rectangle).
d. Use 180 to 240 grit abrasive paper (C112) to
remove guard adhesive. Avoid removing any of the spar.
This will appear as white dust.
Isopropyl alcohol is flammable.
Keep
away from heat and open flame. Provide
adequate ventilation when using. Avoid
breathing vapors and prolonged contact
with skin.
Change 65 5-30BX
TM 55-1520-234-23
Figure 5-17AC. Application of Sealant to Stainless Steel Guard for K747-003-303/403 Blade.
Change 65 5-30BY
TM 55-1520-234-23
Figure 5-17AD. Stainless Steel Erosion Guard Holding Fixture
Change 49 5-30BZ
TM 55-1520-234-23
NOTE
Pot life of adhesive is 15 minutes at 75
degrees F (23.8 degrees C). It is shorter at
higher temperatures. Work without delay.
Isopropyl alcohol can damage leading edge
erosion guard. Avoid spillage.
e. Using cotton tipped swab (kit), dip in isopropyl
alcohol (C23) solvent. Clean surfaces to be bonded.
f. Using masking tape (kit), mask around cut out
section to protect guard from solvents and adhesive.
g. Make a pattern from cut out section. This pattern
will be used to make the replacement patch. Use the 4 x
8 inch patch from kit. Fit patch to mate removed section.
i.
Alternate method.
(1) The alternate method to secure the guard
patch is to use epoxy resin EA828 (C107). The patch can
be moved into position after contacting the adhesive.
(2) Mix 100 parts/weight of resin EA828 with 10
parts/weight of DTA activator (C107) in a clean glass,
metal, polyethylene, or plastic coated paper container.
(3) Using 0.25 inch brush (kit), apply a light
coat of adhesive to both surfaces to be bonded.
Adhesive contains
toxic ingredients.
Provide adequate ventilation and protect
the skin and eyes from contact with
uncured resins or curing agent. DTA can
cause blindness and burns.
Wash off
uncured resins and curing agent from skin
with warm water and soap. Avoid use of
solvents for clean- ing the skin. Wear
polyethylene gloves over cotton gloves for
this task.
Once the contact cement and the patch
come in contact, the patch cannot be
moved if it is mislocated.
It will be
necessary to enlarge the repair sections.
(4) Using finger pressure, press erosion guard
to blade while working out excess adhesive from under
the erosion guard. Wipe away excess adhesive with
clean cheesecloth (kit) to prevent adhesive from running
off the masking onto the exposed blade surface.
(5) Lay teflon parting blanket (kit) over repair.
Place masking tape (kit) over edges of parting blanket to
prevent movement.
(6) Obtain two wooden blocks approximately
0.75 x 2 x 6 inches and a C clamp (8 inch opening by 6
inches deep). Place 0.25 x 2 x 1 inch rubber thick strip
(or suitable substitute) between block and parting blanket.
Place remaining block and rubber strip on opposite
surface and, using C clamp, apply light pressure to
rebonded area.
NOTE
Surfaces to be bonded must be clean, dry,
and free of finger prints and all foreign
matter.
h. Bonding the guard patches.
The preferred
method to secure the guard patch to the guard is to use
contact cement. Suggested mix ratio is 15 grams of
estane to 85 grams of MEK (C87). Cure time is 30
minutes.
Pressure can be applied by vacuum or strips of
rubber around the blade. These methods should be
used on the leading edge where clamps would not be
practical.
(7) After four hours at room temperature,
remove clamp, blocks, rubber strip, parting blanket,
and masking tape.
NOTE
Patch should overlap by 0.50 inch.
Change 49 5-30CA
TM 55-1520-234-23
(8) After cure cycle, bond the patch to the
guard by applying heat with a sealing iron as described in
paragraph 5-5R.
5-5AC. REFINISHING
BLADES (AVUM).
PAINT-K747
MAIN
ROTOR
NOTE
When actual operational emergencies
require immediate use of the helicopter,
touchup painting may be deferred until
termination of the actual emergency.
b. Thinning (for spraying).
The mixed epoxy
polyamide primer shall be reduced for spraying with one
volume of thinner (C142A) to two volumes of mixed
primer. The thinned primer shall be stirred thoroughly,
strained, and allowed to stand for about 30 minutes prior
to use. The thinning ratio may be varied slightly to obtain
the proper spraying viscosity. The 30 minute standing
period is necessary to:
(1)
Permit the chemical components to partly
(2)
Shorten the drying time.
(3)
Reduce cratering.
react.
(4)
migrating.
The only paint refinishing authorized is the
touchup of repaired areas and areas
immediately adjacent to repaired areas.
This restriction is necessary to maintain
lightning protection and radar signature
characteristics of the blade. Refinishing
paint must not be applied to leading edge
erosion guard.
Only material from the same kit shall be
mixed, except that two or more kits may be
mixed in the same vessel, provided the kits
are all manufactured by the same vendor.
Established mixing ratios must be followed
closely; other- wise, the primer will exhibit
unsatisfactory film properties,
such as
poor adhesion, poor chemical resistance,
or inadequate drying. Component II shall
always be added to component I.
NOTE
The epoxy polyamide primer is supplied as
a two component kit. Pot life is limited and
only that amount which can be used in less
than 8 hours should be mixed.
a. Mix component I and II in a one to one ratio by
volume. Each component shall be well agitated and shall
be poured separately into the proper container. The
material temperature should be at least 70 degrees F (21
degrees C). Component I shall be poured into the empty
container, then component II shall be slowly poured into
component I with constant stirring.
5)
Preclude component II from sweating out or
Allow any bubbles (formed while stirring) to
escape.
c. Feather edges of finish next to repair area with
400 grit abrasive paper.
d. Remove sanding dust using clean cheesecloth
(C36) dampened with thinner (C 142A).
e. Mask off touchup area.
f. Wipe area with 50/50 mix of MEK (C87) and
lacquer thinner (C140).
g. Apply primer (C100) slightly overlapping repair
area. Allow to dry approximately 5 minutes.
h. Apply a cross coat of primer and allow to dry
about 30 minutes. If temperature is below 70 degrees F
(21 degrees C), allow to dry about 2 to 3 hours. Do not
apply below 50 degrees F (10 degrees C).
i. Brush application. Mix one volume of component
I to one volume of component II. If thinning is required,
use thinner (C142A). Apply only one brush coat of
primer. The same temperature limitations in step h above
apply.
j. Mix component I and II of polyurethane (C98) in
the correct ratio according to manufacturer's instructions.
k. Cross spray polyurethane (C98) over the primer
to a thickness of 0.0010 to 0.0015 inch.
Change 49 5-30CB
TM 55-1520-234-23
NOTE
Avoid
overspray
onto
existing
polyurethane finish. Polyurethane will not
adhere properly to a previously coated area.
1. Allow to dry approximately 6 hours before releasing
helicopter for flight.
5-5AD. MASKING AND REFINISHING LEADING EDGE
EROSION GUARD- K747 MAIN ROTOR BLADES
(AVIM).
NOTE
This procedure is for use at station 213.5
outboard after installation of repair kit P/N
K747-206.
a. Ensure blade is properly grounded.
b. Mask area for application of primer.
(1) Apply masking tape and brown paper at the
forward side of the erosion guard trailing edge on both
blade top and bottom surfaces. Brown paper will be
applied to cover unaffected blade surfaces and erosion
guard area from primer and polyurethane overspray.
(See detail A, figure 5-17AE.)
(2) Apply a second strip of masking tape overlapping
the first strip applied. The trailing edge of the second
masking being applied must be along a line 0.100 inch aft
of the erosion guard trailing edge. (See detail B, figure
5- 17AE.)
When abrading the polyurethane finish, care
must be exercised to only scuff the surface to
allow primer adhesion. Do not remove the
full coating thickness.
c. Abrade the polyurethane top coat on exposed blade
top and bottom surfaces from station 205.5 outboard.
Use 360 grit abrasive paper and just scuff the surface
lightly.
Primer and thinner mixture contains toxic
ingredients. Provide adequate ventilation
and protect
the skin and eyes from contact with
uncured resins or curing agent. Wash off
any primer from skin with warm water and
soap. Avoid use of solvents for cleaning
the skin.
NOTE
Epoxy polyamide primer is supplied as a
two component kit. Pot life is limited to 8
hours from mixing. Mix only the amount
needed. Material temperature shall be at
least 70 degrees F (21 degrees C) before
mixing.
d. Mix and prepare primer for use.
(1) Stir component I and II in their respective supply
containers. Ensure thorough mixing occurs.
(2) Combine an equal volume of component II into
component I. Stir constantly while combining mixing. Do
not induce air bubbles in the mixture.
NOTE
Thinning ratio of mixed primer and thinner
may be varied slightly to obtain a proper
spraying viscosity.
(3) Thin the mixed primer for spraying using a
volume ratio of 1 part primer to 2 parts thinner (C142A).
Stir the thinned mixture thoroughly. Do not induce air
bubbles. Strain the mixture and allow to stand for 30
minutes prior to use.
Surfaces to be primed must be clean, dry
and free of finger prints and all foreign
matter.
e. Wipe the scuffed area with a 50/50 mixture of MEK
(C87) and lacquer thinner (C140).
Primer and thinner mixture contains toxic
ingredients. Provide adequate ventilation
and protect the skin and eyes from contact
with uncured resins or curing agent.
Wash off any primer from skin with warm
water and soap. Avoid use of solvents for
cleaning the skin.
Change 49 5-30CC
TM 55-1520-234-23
Figure 5-17AE. Masking for Paint Touch-up After Installation of Kit K747-206.
Change 49 5-30CD
TM 55-1520-234-23
f. Apply a full coat of primer to the bare area next to the
second masking tape applied. Apply a mist coat of primer
to the areas scuffed. Allow 1 hour drying time prior to
application of polyurethane paint.
g. Remove the second piece of masking tape applied
which extends 0.100 inch aft of the erosion guard trailing
edge.
h. Hand abrade the ridge where the epoxy primer ends.
Use 360 grit abrasive paper.
i. Mix components I and II of the paint (C98) as per
manufacturer's instructions. Heed the manufacturer's
warnings and cautions.
j.
NOTE
Tape all holes in the blade (bullet damage,
tree damage, foreign object damage, etc.)
to protect the interior of the blade.
(2) Thoroughly clean root fitting. Apply grease (C67)
to root fitting bolt hole, drag brace bolt hole, and all
exposed unpainted surfaces.
(3) Wrap blade with barrier material (C29), shiny
side next to blade, at all locations where blade will
contact the molded hair supports (5 places) and secure
with pressure sensitive tape (C136).
(4) Attach a properly filled out DD Form 1577-2
(Unserviceable/Repairable) tag directly to the blade.
Dust sanding residue from the blade surface.
k. Apply a mist coat of paint (C98) to the blade areas
framed by masking tape. Allow the mist coat to dry for 30
minutes.
(5) Place blade in container.
l. Apply a full coat of paint (C98} to the misted area. Do
not exceed 0.0010 to 0.0015 inch thick.
(7) Secure lid.
m. Remove remaining masking and allow to dry for 6
hours minimum before blade is used for flight.
5-5AE. REPAIR OF AFT
ROTOR BLADES (AVIM).
TIP
CAP-K747
MAIN
a. Cracks in aft tip cap may be sanded and routed to a
depth of 0.060 inch.
b. Apply adhesive EA934NA (C17). Smooth to contour
of cap by sanding.
c. Paint repair area in accordance with paragraph 517AC.
5-5AF. PREPARATION
FOR
STORAGE
SHIPMENT-K747 MAIN ROTOR BLADES.
OR
a. The following instructions cover storage or shipment
of main rotor blades in container P/N K747-001.
(1) Thoroughly remove foreign matter from entire
exterior surface of blade using clean cheesecloth (C36).
(6) Secure blade to shock mounted support.
(8) Secure blade log in container log compartment.
5-5AG. INSTALLATION-K747
BLADES.
a. Obtain a balanced
6A).
MAIN
ROTOR
main rotor hub (paragraph 5-
b. Support main rotor hub on a build-up bench in
accordance with paragraph 5-4a. Check that locating pin
(6, figure 5-11) is installed in upper surface of each grip
(5) at inboard side of retaining bolt hole.
c.
Install drag strut (15, figure 5-17A).
d. Remove preservative grease from blade grip bore
and retaining bolt.
e. Apply corrosion preventive compound (C53) to blade
retaining bolt, hub grip, blade butts, and washer with
notches on locating pin (6). Slide blade (9, figure 5-11)
gently into grip (use of sling is optional). Place washer (7)
on retaining bolt (8). Align bolt holes carefully and insert
bolt from top. If bolt binds, move tip of blade up and
down slowly to find position which
Change 65 5-30CE
TM 55-1520-234-23
allows bolt to pass through without binding. Seat bolt and
washer with notches on locating pin (6).
k. If the blades are not to be aligned in the hub, torque
both nuts (11) 125 to 150 foot-pounds.
f. Place padded support under blade approximately one
third blade length inboard from blade fin.
l. If the blades are not to be aligned in the hub, use
wrench (T31) to tighten nuts (17) to a torque of 475 to 525
foot-pounds. Align a notch in the nut with a hole in the
bolt. Install locking screw (20) with head in a direction so
that centrifugal force will keep the locking screw in. In
some cases, this may mean the locking screw may be
installed from the inside of the bolt. Install washer (19)
and nut (18).
Install washer (16) with counterbore up
facing grip.
The erosion guard is a polycarbonate
material which will cut easily upon impact
with a rigid structure. Seal all openings
immediately with sealing iron.
m. Install grip locks (T59) on each pitch horn,
previously accomplished (figure 5-7).
5-5AH. ALIGNMENT-K747 MAIN ROTOR BLADES.
g. Install washer (16) with counterbore up as illustrated
and install nut (17). Do not tighten nut at this time.
Refer to paragraph 5-4.
5-6.
h. Preset drag brace (15) length to 14.750 inches, hole
center to hole center. Align clevis of drag brace (15) on
bolt hole of the blade drag plates. Install shims (10)
equally between clevis and upper and lower drag plates to
obtain 0.000 to 0.005 inch clearance. Install bolt (14) and
secure with two washers (12 and 13) and nut (11) on
lower end. Do not tighten at this time.
i.
Install opposite blade in the same manner.
j. If the blades are to be aligned in the hub,
instructions in paragraph 5-4.
if not
follow
Main Rotor Hub Assembly. (AVIM)
The Main Rotor Hub major components are the yoke,
trunnion extensions, blade grips, drag braces, pitch
horns, and elastomeric bearings. See figures 5-18 and 519. The elastomeric bearings (6, figure 5-19) are
composed of alternating layers of an elastic material
(elastomer) with concentric cylindrical metal laminations
molded to steel inner and outer housings. The bearing
outer housing is bolted to the rotor yoke. The bearing
inner housing is bolted to the trunnion. Movement of the
yoke and blades on the flapping axis is by flexing of the
bearing elastomer.
Change 65 5-30CF
Premaintenance Requirements for Main Rotor Hub
Condition
Requirements
Model
Part No. or Serial No.
Special Tools
Test Equipment
Support Equipment
Minimum Personnel
Required
Consumable Materials
Special Environmental
Condition
AH-1S
540-011-101-17
(T26) (T31) (T34) (T42)
(T43) (T45) (T47) (T56)
(T59) (T79) (T80) (T81)
NA
Work aid for removal of
blade retaining bolt
Two
(C9) (C12) (C17) (C32)
(C37) (C44) (C45) (C62)
(C82) (C87) (C88) (C99)
(C112) (C113) (C116)
(C136) (C151)
NA
(8) Identify hub components which will reach
retirement time prior to next scheduled inspection for
replacement.
Refer to Overhaul and Retirement
Schedule.
(9) Inspect sand deflectors for cracks and
deflection doublers for corrosion.
(10) Remove two sand deflectors, if installed,
from each main rotor hub to be inspected.
(11) Check the clearance between each
extension and grip assembly by inserting the 0.100 to
0.125 inch thick strap between them. Move the strap
through the full range of gap between grip and extension.
(a) The hub and metal blade angle set in
accordance with paragraph 5-4.b. (13) (d).
a. Inspection-Assembled Main Rotor Hub.
(1) Inspect exposed surfaces of assembled
main rotor hub for nicks scratches and corrosion. See
figures 5-21, 5-23, 5-24, 5-25, 5-26, 5-26A, 5-27, 528 and 5-29 for damage limits.
(2) Inspect open bolt holes for scratches,
gouges and corrosion.
(3) Inspect elastomeric bearings for elastomer
squeeze-out and delamination.
Crazing and slight
cracking of elastomer due to weather exposure is not
cause for replacement.
(4)
TM 55-1520-234-23
(7) Inspect hub historical records and the hub
for evidence that the hub has been subjected to an
accident or incident outside the realm of normal usage. If
such evidence exists,
perform special inspections
outlined in Chapter 1.
Inspect for scuffing due to contact between
parts.
(b) The hub and fiberglass blade angle set
in accordance with paragraph 5-4.b. (13) (e).
(c) Disconnect
paragraph 5-4.a.(1).
the
pitch
links
per
(d) Rotate hub and blade and check for
clearance with a total positive angle of attack of 17
degrees on hub with metal blades or 20 degrees on hub
with fiberglass blades.
(e) Rotate hub and blades and check for
clearance with zero angle of attack.
Interference will not occur in normal operation, but can
occur at extreme control positions during ground
operation of controls with external hydraulic power applied
while main rotor is static.
(f) Main rotor hubs which do not allow the
0.100 to 0.125 inch thick strap to pass should be
disassembled for visual check for interference rubbing or
other signs of contact between grip and extension. Inspect
carefully the outboard end of extension barrel and grip for
interference. If there are no visual signs of contact on
grip and extension the hub may be reassembled. All
parts with signs of interference must be replaced.
(5) Inspect trunnion for damaged splines. See
figure 5-28 for allowable damage limits.
(12) Replace extension assembly if any clearances
cannot be met.
NOTE
(6) If any damage is present for which no limits
are specified and/or there is damage beyond limits shown
on figure 5-28, replace the affected part.
b. Disassembly.
(1) Position main rotor hub on build-up bench (T26)
equipped with adapter plate (T34) if not previously
accomplished. Refer to paragraph 5-4, for procedure.
(2) Remove main rotor blades from hub if not
previously accomplished. Refer to paragraph 6-5.
Change 71 5-30CG
(3) Identify blade retaining bolt assemblies (6,
figure 5-18) for reinstallation in the same grip. Use paint
or felt tip pen. Remove both bolts. Use socket wrench
(T31) to remove nuts from blade retaining bolts. Use
work aid shown on figure 5-12 to remove blade retaining
bolts if necessary. Refer to paragraph 5-5 for procedure.
NOTE
Units operating AH-1 aircraft may remove the
sand deflectors P/N 540-011- 174-11, NSN
1615-00-116-7110 from the rotor head. Upon
removal, deflectors should be inspected for
serviceability and repaired as required. The
deflectors will be retained as part of the aircraft
mission equipment. Deflectors will be installed
in extreme sand/dust conditions or in arctic
areas where there are extreme ice/snow
conditions. Both deflectors must be either
installed on the aircraft or removed from the
aircraft.
(4) Remove three bolts (2, figure 5-18), sand
deflector (1), and spacers (3). Remove opposite sand
deflector in the same manner.
TM 55-1520-234-23
(12) Remove yoke and trunnion from buildup
stand and place on a work bench with supports under the
flat portion of the yoke.
(13) Identify elastomeric bearings (6, figure 519) trunnion (9) and yoke (12) with felt-tipped marker so
the bearings and trunnion can be reinstalled in the same
position on the yoke.
(14) Remove four bolts (1,
figure 5-19),
washers (2) and retainer (3) from each side of trunnion.
(15) Remove two nuts (11,
figure 5-19),
washers (10), bolts (4) and washers (5) to free one
elastomeric bearing from yoke. Thread a bolt with 1/2 -x
20 UNF threads into tapped hole in elastomeric bearing
(6). Tighten bolt so that it bears against disk (8) and pulls
elastomeric bearing free of yoke and trunnion. Remove
opposite elastomeric bearing in the same manner and
remove trunnion from yoke.
(16) Remove shims (7, figure 5-19) from both
sides of trunnion and identify for reinstallation in the same
location.
c.
Cleaning.
(5) Remove bolt (18, figure 5-18) and drag
brace (15).
Remove opposite drag brace in same
manner.
WARNING
(6) Remove bolts (19 and 21, figure 5-18) and
pitch horn (23). Remove opposite pitch horn in the same
manner.
Use solvent (C-124) in a well-ventilated area.
Avoid prolonged breathing of vapors and do not
use in an area with open flame or high
temperature.
(7) Remove cotter pin (57, figure 5-18) nut
(55) and washer (56) from bolt (42). Remove bolt (42),
clamp (43) and lock (44). Remove dome nut (45).
(1) Clean all metal parts with solvent (C124)
and dry with compressed air.
WARNING
(8) Remove nuts (48, figure 5-18) and washers
(47) from bolts (60). Remove bolts. Remove blade grip
(41) from yoke extension (28). Use care to prevent
damage to threads on fitting (52) and damage to dust seal
(51). The dust seal (51) bearing (50) and strap indexing
ring (49) should be bonded to the grip near the outboard
end.
Provide adequate ventilation when using methylethyl-ketone (C87). Avoid breath- ing solvent
vapors and avoid prolonged con- tact with skin.
(9) Remove the opposite grip in the same
manner outlined in steps (7) and (8).
Do not allow methyl-ethyl-ketone (C87) to saturate the
teflon bearings or contact the elastomer portion of
elastomeric bearings.
(10) Remove nuts (40, figure 5-18) washers
(29), bolts (27) and washers (26). Remove yoke
extension (28) and housing (38) from yoke (25). Remove
housing (38) from extension. Remove opposite yoke
extension in the same manner.
(2) Clean teflon bearings in housing (38, figure
5-18) and grip (41) with clean cloths dampened with
methyl-ethyl-ketone (C87).
(11) Clean sealant from retainer rings (31 and
34, figure 5-18) with a sharp plastic scraper. Remove pin
(33) and strap (54). Remove opposite strap in the same
manner.
(3) Clean old sealant and zinc chromate primer
from spindles of trunnion and inner metal housing of
elastomeric bearings (6, figure 5-18). Use a sharp plastic
scraper and cloths moistened with methyl-ethyl-ketone
(C87). Do not allow the
CAUTION
Change 71 5-31
TM 55-1520-234-23
Figure 5-18. Main rotor hub yoke extension and grip assembly (Sheet 1 of 3)
Change 77 5-32
TM 55-1520-234-23
Figure 5-18. Main rotor hub yoke extension and grip assembly (sheet 2 of 3)
5-33
TM 55-1520-234-23
Figure 5-18. Main rotor hub yoke extension and grip assembly (Sheet 3 of 3)
methyl-ethyl-ketone to contact the elastomer portion of
the elastomeric bearings.
NOTE
Mating surfaces are the surfaces of a
component that come in contact with another
component when the hub is assembled.
d. Inspection-Disassembled Main Rotor Hub.
(1) Inspect teflon bearings in housings (38,
figure 5-18) and grips (41) for wear and damage.
Compare the bearings with the examples shown on figure
5-20 to determine whether the bearings are suitable for
further service. Also inspect teflon bearings for secure
bonding of fabric.
(a) Mechanical and corrosion damage limits on
mating surfaces are 0.020 inch in depth and one-half of
the quadrant of area around bolt holes. Closer tolerances
apply when specified in inspection instructions for
individual components; i.e., the limit for yoke extension
bearing sleeves is 0.002 inch as shown on figure 5-21.
(2) Inspect yoke extension for damaged, worn
or loose bearing sleeves (30 and 35, figure 5-18). If the
bearing sleeves are loose or have damage in excess of
limits shown in figure 5-21, the yoke extension must be
replaced. If the bearing sleeves have any superficial
marks, polish out the marks with Scotchbrite (C113) and
reinspect the bearing sleeves.
(b) All hub components must assemble without
misalignment or cocking after polishing out mechanical
and corrosion damage.
(5) Inspect the holes in the hub components
illustrated on figure 5-22 for wear in excess of limits.
(3) Inspect yoke extensions for worn or missing
buffer pads (32, figure 5-18).
(6) Inspect seals (36, figure 5-18) for damage which
would affect function and for secure bonding to housings.
(4) Inspect mating surfaces of the hub
components for damage in excess of the following limits:
5-34
TM 55-1520-234-23
Figure 5-19. Main rotor hub yoke and trunnion for main rotor hub
Change 22 5-35
TM 55-1520-234-23
Figure 5-20. Main rotor hub teflon bearing wear patterns
5-36
TM 55-1520-234-23
Figure 5-21. Damage limits - main rotor yoke extension.
Change 7 5-37
(7) Inspect grip bearing dust seals (51, figure
5-18) for damage which would affect function and for
secure bonding to grips.
TM 55-1520-234-23
(11) Inspect retainers (3, figure 5-19) for
obvious damage.
(12)
(8) Inspect radius rings (29, figure 5-18) for
damage and wear. Replace radius rings if the following
limits are exceeded.
(a) Cracks in the carbon face running
from the inside diameter to the outside diameter.
(b) Grooves in the carbon face which
reveal uneven contact with the seal.
(c)
Any chips of carbon missing from the
(d)
Unbonded abrasion shield.
carbon face.
NOTE
Partially unbonded abrasion shield, discoloration
of the shield and/or less than 10 cracks in the
shield are not cause for replacement.
(9) Inspect straps (54, figure 5-18) for damage
and wear in excess of the following maximum limits:
(a) Fifty loose wire ends protruding
through the urethane coating of strap in any one corner
and/or 400 loose ends over the entire strap assembly. If
a lesser number of wire ends are found, record the serial
number of the strap and the number of wire ends found in
the historical record of the main rotor hub.
(b) Use a ten power magnifying glass to
check for cracks in flanges of strap bushings and
urethane wedges. A crack in these parts is cause for
rejection of the strap.
Inspect disks (8, figure 5-19) for obvious
damage.
(13) Inspect sand deflectors (1, figure 5-18)
for cracks, abrasion damage, oversize bolt holes and
corroded,
worn,
damaged,
separated or missing
washers. Replace sand deflectors having any cracks
greater than two inches in length in any portion. Cracks
less than two inches in length are acceptable if stop
drilled, provided crack does not permit material fallout.
(14) Inspect elastomeric bearing housing for
cracks by the magnetic particle methods, TM 43-0103
(figure 5-29). If there is any doubt about serviceability of
elastomeric bearings after inspection,
forward the
bearings to higher level of maintenance for evaluation.
Inspect elastomeric bearings with a 10 power magnifying
glass for delamination of the concentric metal laminations
and the elastomer. Also inspect for delamination of the
bear- ing inner and outer housings. Crazing or cracking of
the elastomer or shedding off of small scraps at end of
bear- ing due to weather exposure is not cause for
rejection. Use a standard 0.005 inch, blunt-end, feeler
gage to check for delamination in excess of the following
limits:
Gage Penetration
% Area/Laminate
0.000 TO 0.250
50%
0.000 TO 0.500
25%
Replace any bearing found to have a cracked concentric
metal laminations. Cracked concentric metal laminations
may be detected by either visual inspection or by running
a fingernail along the edge of the shim.
(15) Inspect hub components illustrated on
figures 5-21, 5-22, 5-23, 5-24, 5-25, 5-26, 5-27, 5-28
and 5-29 for damage in excess of limits.
(c) Severe rupture of the urethane coating
is cause for rejection due to difficulty of installation of the
damaged strap in the yoke extension.
(16)
Deleted.
e. Repair.
(d)
Displacement of urethane wedges
between bushing and inner surface of wire bundle.
(e) Pins (33 and 58) for presence of
retaining rings (31, 34 and 53, 59).
NOTE
A permanent set twist in the strap and/or a slight
bulging of wire cross section is normal and not
cause for rejection of the strap.
(10) Inspect attaching
washers for damage and corrosion.
bolts,
nuts
and
(1) Polish out all traces of corrosion and
mechanical damage on hub components. Polish out
corrosion dam- age on aluminum parts to twice the depth
of the pit. Use fine to medium grades of abrasive cloth
(C44) or fine in- dia stone (C128). Blend the edges of the
repair into the surrounding area with a smooth contour.
Make final cleanup with crocus cloth (C45) to obtain a
smooth, scratch-free surface. If damage exceeds limits
specified in paragraph c. scrap the part.
(2) If cadmium plate is removed, touch up with
a brush coat of cadmium plate (C32).
Change 54 5-38
TM 55-1520-234-23
Figure 5-22. Damage limits - main rotor hub holes (Sheet 1 of 2)
Change 71 5-39
TM 55-1520-234-23
Figure 5-22. Damage limits - main rotor hub bolt holes (Sheet 2 of 2)
Change 71 5-40
TM 55-1520-234-23
Figure 5-23. Damage Limits - Main Rotor Hub Yoke
Change 65 5-41
TM 55-1520-234-23
Figure 5-24. Damage limits - main rotor hub grip
Change 38 5-42
TM 55-1520-234-23
Figure 5-25. Damage limits - main rotor hub extension and strap fitting
Change 56 5-43
TM 55-1520-234-23
Figure 5-26. Damage limits - main rotor hub pitch horn and inboard bearing housing
(Prior to accomplishment of MWO 55-1520-244-50-6) and pitch horn bushing
(Sheet 1 of 2)
Change 71 5-44
TM 55-1520-234-23
Figure 5-26. Damage limits - main rotor hub pitch horn and inboard bearing housing
(Prior to accomplishment of MWO 55-1520-244-50-6) and pitch horn bushing
(Sheet 2 of 2)
Change 71 5-44A
TM 55-1520-234-23
Figure 5-26A. Damage limits - main rotor hub pitch horn and inboard bearing housing
(Prior to accomplishment of MWO 55-1520-244-50-6) and pitch horn bushing
(Sheet 1 of 2)
Change 71 5-44B
TM 55-1520-234-23
Figure 5-26A. Damage limits - main rotor hub pitch horn and inboard bearing housing
(After accomplishment of MWO 55-1520-244-50-6 and MWO 55-1520-244-50-9 (Sheet 2 of 2)
Change 71 5-44C/(5-44D blank)
TM 55-1520-234-23
Figure 5-27. Damage limits - main rotor drag brace
Change 7 5-45
TM 55-1520-234-23
Figure 5-28. Damage Limits - Main Rotor Hub Trunnion.
Change 29 5-46
TM 55-1520-234-23
Figure 5-29. Damage limits - main rotor hub elastomeric bearing
Change 7 5-47
(3) Touch up rework areas on aluminum parts
with chemical film (C37).
(4) Replace damaged or missing buffer pads
(32, figure 5-18) as follows:
CAUTION
Do not remove cadmium plate from yoke
extension except in the area where buffer pads
will be installed.
(a) Remove any buffer pad material that
remains bonded to the yoke extension with a plastic or
aluminum scraper.
TM 55-1520-234-23
cadmium plate from yoke extension. If necessary, touch
up cadmium plate. Refer to step (2).
(f) Check fit of yoke extension to yoke. If
the new buffer pads are too thick, sand the pads with fine
grit sandpaper (C112) to obtain a slip fit. The correct
dimension is 1.377 TO 1.379 inch over the yoke
extension and the buffer pad on each side of the
extension.
(5) Replace radius ring (29, figure 5-18) which
failed to pass inspection as follows:
(a)
aluminum drift.
Remove radius ring (29) with a soft
WARNING
WARNING
Provide adequate ventilation when using methylethyl-ketone. Avoid breathing solvent vapors
and avoid prolonged skin contact.
Provide adequate ventilation when using
methyl-ethyl-ketone. Avoid breathing solvent
vapors and avoid prolonged skin contact.
(b) Clean area where new buffer pad will
be installed with 400 grit sandpaper (C112). Remove
residue with clean cloths dampened with methyl-ethylketone (C87). Clean the side of the new buffer pad that
will be bonded in the same manner.
(b) Clean old adhesive from yoke
extension with a plastic scraper and clean cloths,
moistened with methyl-ethyl-ketone (C87).
WARNING
Use primer in a well ventilated area and away
from open flame.
Use primer in a well ventilated area and away
from open flame.
(c) Apply a coat of primer (C99) to the
cleaned surfaces of the yoke extension and allow to cure
for thirty minutes.
(d) Mix adhesive (C12 or C14a) in
accordance with instructions on the container. Apply a
thin coat of adhesive to the mating surfaces of the buffer
pad and the yoke extension. Position the buff- er pad on
the yoke extension and clamp in place. Use a C-clamp
and two flat smooth plates, or use a bolt and two plates
with holes in the plates for the bolt. Use cellophane (C33)
or polyurethane tape (C134.1) between the buffer pad and
the plates. Cure adhesive for 12 hours at 80 degrees F
(27 degrees C) or for one hour at 160 degrees F (71
degrees C).
(e) Remove clamp and use 180 grit
sandpaper (C112) to remove any excess adhesive that
was squeezed out during bonding. Avoid removing
WARNING
(c) Apply a light coat of primer (C99) to
mating surfaces of yoke extension and radius ring and
allow to cure for thirty minutes.
(d)
Mix adhesive (C12) in accordance
with instructions on the container. Apply a thin coat of
adhesive to the mating surfaces of the yoke extension
and the radius ring.
(e) Press the radius ring into position on
the yoke extension. Ensure that the radius ring is
completely seated. Wipe off excess adhesive. Install
grip (41) and housing (38) on yoke extension (28). Install
dome nut (45) and tighten until grip bearing dust seal (51)
evenly contacts radius ring (29) and then tighten an
additional 1/2 turn. Allow adhesive to cure for 24 hours at
room temperature or for thirty minutes at 150 degrees F
(66 degrees C).
(f) Remove dome nut (45), grip and
housing installed in preceding step and ensure that radius
Change 71 5-48
TM 55-1520-234-23
(d) Mix adhesive (C17) according to directions
on the container. Apply a thin coat of adhesive to the
mating surfaces of the chafing pad and the yoke.
Provide adequate ventilation when using
methyl-ethyl-ketone. Avoid breathing solvent
vapors and avoid prolonged skin contact.
Figure 5-30. Main rotor hub yoke chafing pad
installation dimensions
ring is completely seated on yoke extension.
(6)
Replace damaged or missing pads (figure
5-30).
NOTE
Chafing pads are located on the upper and the
lower surfaces of the yoke.
(a) Remove any chafing pad material that remains
on the yoke with a plastic scraper and clean cloths
dampened with naphtha (C88). Do not remove paint.
(e) When adhesive applied in preceding
step becomes tacky, install chafing pad at location
illustrated on figure 5-30. Work out any air pockets
and wipe up any excess adhesive with a cloth
dampened with methyl-ethyl-ketone (C87). Apply
weights or use a clamp to hold chafing strip firmly in
position.
Cure adhesive for 24 hours at room
temperature or one hour at 176 TO 190 degrees F (79
TO 88 degrees C).
(7) Replace seal (36, figure 5-18) and bearing (37) in
bearing housing (38) as follows:
(a) Grasp seal (36) with "duck bill" pliers and
tap pliers with mallet to remove the seal.
(b) Install bearing puller (T47) into housing
(38). Use the aluminum block as shown in figure 5-31 to
prevent damage to the housing. Apply moderate tension
with puller.
Do not use flame of any form on the housing
assembly.
(c)
Apply heat to housing with heat lamp for
approximately thirty minutes or until yielding of adhesive
is evident. Increase tension with puller and remove
bearing.
Provide adequate ventilation when using
methyl-ethyl-ketone.
Avoid breathing
solvent vapors and avoid prolonged contact
with skin.
(b) Clean new chafing pad with methyl- ethylketone (C87). Treat the side of the pad that is to be
bonded with tetra-etch (C62).
(c)
After etching,
methyl-ethyl-ketone (C87).
rinse the chafing pad with
Provide adequate ventilation when using
methyl-ethyl-ketone. Avoid breathing solvent
vapors and avoid prolonged contact with
skin.
(d) Clean all traces of adhesive from bearing
housing (38, figure 5-18) with a plastic scraper and cloths
moistened with methyl-ethyl- ketone (C87).
Change 22 5-49
TM 55-1520-234-23
Ensure that adhesive does not get on teflon
bearing fabric.
NOTE
Larger metal surface of seal which is marked with
part number will be in- stalled flush against
bearing, inboard of bearing cap.
(f)
Mix adhesive (C12) according to directions
on container: Apply a coat of adhesive to the mating
surfaces of bearing (37, figure 5-18) and housing (38).
Install bearing in housing and seat fully as shown in figure
(5-31). Wipe off excess adhesive with a cheese cloth
moistened with methyl-ethyl-ketone (C87). Cure for 24
hours at room temperature or for fifteen minutes at 1650
degrees F (66 degrees C).
(g) Apply a thin, even coat of adhesive (C9) to
mating surface of seal (36, figure 5-18). Allow cement to
dry about ten minutes or until it becomes tacky and press
seal into housing (38) with lip of seal facing outboard.
Allow to cure for four hours at room temperature.
(8) Replace seal (51, figure 5-18) and bearing (50) in
grip (41) as follows:
NOTE
Figure 5-31. Tool application-bearing removal from
housing
If bearing (50) is satisfactory for further service,
perform only the steps applicable to removal
and installation of seal (51).
(a) Grasp seal (51, figure 5-18) with "duck bill"
pliers and tap pliers with mallet to remove seal.
Use primer in a well ventilated area away from
open flame.
(e) Apply a light coat of primer (C99) and allow
to cure for thirty minutes.
(b) Install bearing remover (T43) in grip with
slot of-tool over tang of strap indexing ring (49, figure 518). See figure 5-32 for view of installed tool. Drive ring
and bearing from grip.
(c)
scrapers.
Clean all old adhesive from grip with plastic
(d)
Install strap index ring (49, figure 5-18) in
grip with lugs on ring engaged with slots in grip. Fill
keyway gap between strap index ring and grip with
sealant (C116).
Provide adequate ventilation when using
methyl-ethyl-ketone.
Avoid breathing solvent
vapors and avoid prolonged skin contact.
NOTE
No adhesive is required on bearing (50, figure 518) if it is P/N 540-011-110-7 or 540-011-110-13.
Change 38 5-50
TM 55-1520-234-23
(e) Press new bearing (50, figure 5-18)
into grip with arbor (T45). See figure 5-33 for view of
installed tool with ring of tool positioned to engage
bearing. Press bearing into grip as described in the
note on figure 5-33 until it fully engages the strap
indexing ring. Remove dome nut and special tools.
(f) Apply a thin even coat of adhesive
(C9) to mating surfaces of dust seal (51, figure 5-18)
and grip (41). Allow cement to dry about ten minutes
or until it becomes tacky. Position dust seal on tool
(T45) as shown on figure 5-33 and position in grip.
Press dust seal into grip as described in the note on
figure 5-33 until it fully engages the seat. Remove
dome nut and special tools. Spread any excess
cement that has squeezed out to form a fillet between
the dust seal and the grip.
(g) Repair sand deflector. Cracks of
two inches in length or less may be stop drilled.
Corrosion, erosion, or damage to doublers may be
polished out and painted with primer (C102).
(h)
Bonding of washer (AN970-4).
Replace if corroded, worn or damaged, separated or
missing.
1 Remove all remaining adhesive using
aluminum or plastic scraper.
2
After removing all traces of oil
adhesive, wipe area with cleaner (C87), allow to dry.
3 Lightly sand surface on one side of
new washer using 320 grit sandpaper (Cl12). Clean with
cleaner (C87), allow to dry.
Figure 5-32. Tool application-bearing
removal from grip
Change 65 5-50A/(5-50B blank)
TM 55-1520-234-23
4 Apply a thin coat of adhesive (C12) to
sanded and cleaned surface of washer and position on
deflector.
shims to go in the bearing on one side of the yoke must
be equal, within 0.002 inch, to the total thickness of the
two shims to go in the opposite bearing.
NOTE
The installation of standard dust deflector is
optional.
(d)
Place the new shims (7) selected in the
preceeding step in elastomeric bearings (6). Align holes
in shims and bearings and insert two bolts (1) to maintain
alignment.
(9) Pitch horn (P/N 209-010-109). Replace
worn or damaged bushings (20), and (22), figure (5-18)
in pitch horn (30). Replace bushings (20) and (21),
exceeding the limits of figure 5-26A.
(e) Carefully push bearings over end of
trunnion spindle and align bolts (1) installed in previous
step with holes in end of trunnion spindle. Thread bolts
into trunnion finger tight.
NOTE
(f)
Install two bolts (4) to secure each
elastomeric bearing (6) to the yoke (12). Install recessed
washers (5) on bolts (4) with recessed side toward bolt
heads. Install recessed washers (10) with recessed side
toward nuts (11). Tighten nuts (11) on one bearing (6)
sufficiently to hold flanges of bearing against yoke. Leave
nuts (11) on opposite bearing loose.
Bushings are the expandable type and should not
be a tight fit, except under torque load.
(a) Press out bushings using a plug
slightly smaller than O.D. of bushings.
(g)
Ensure that trunnion spindles are fully
seated in bearings (6) against shims (7).
WARNING
Before using methyl-ethyl-ketone, extinguish all
open flames and turn off electrical equipment.
Vapors are highly flammable, avoid prolonged
breathing of vapors and repeated skin contact.
Use in well ventilated area.
(b) Clean primer from bores using
cheesecloth (C36) dampened with methyl-ethyl-ketone
(C87).
(h)
Measure gap between flange of
elastomeric bearing (6) and yoke (12) on the side where
nuts (11) were left loose. Use a feeler gage to make
measurement and record the dimension. If no gap is
present,
shims (7) are not thick enough. Remove
elastomeric bearings and add an equal amount of shims
(7) to each bearing. Measure and record gap dimension
as described above.
(c) Inspect pitch horn bores for damage.
Damage not to exceed 0.002 inch for one-fourth
circumference.
(i)
Remove elastomeric bearings (6) from
trunnion and remove shims (7) from bearings. Keep
shims with the bearing from which. they were removed.
(d)
New bushings will be installed during
assembly.
f.
Assembly.
(1) Install trunnion (9, figure (5-19) on yoke
(12) as follows:
(a)
Position trunnion (9) in bosses on yoke (12).
(b) Install disk (8) in counterbore in end of each
trunnion spindle.
(c)
Select two new shims (7) to be installed in
each elastomeric bearing (6). The total thickness of two
(j)
Divide the dimension recorded in step (h)
by two. Record this dimension. Peel laminations equal to
this dimension plus 0.002 minus 0.000 from shim (7) for
each bearing (6). See the following example:
Dimension of original measured gap ..............0.022 inch
Original gap divided by two ............................0.011 inch
Thickness of laminations to be removed
from shim for each bearing ............................0.012 inch
(k)
Measure
thickness
of
shims
after
adjustment.
The thickness of shims (7) for each
elastomeric bearing must be equal within 0.002 inch.
Change 65 5-51
TM 55-1520-234-23
Figure 5-33. Tool application - bearing and seal installation in grip
5-52
TM 55-1520-234-23
Figure 5-34. Main rotor hub trunnion centering measurement
(l) Place shims (7) prepared in preceding steps in
elastomeric bearings (6). Align holes in shims and
bearings and install two bolts (1) in each bearing to
maintain alignment.
(m)
Apply unreduced primer (C102) to
mating surfaces of elastomeric bearings (6) and yoke
(12). Position trunnion (9) in bosses of yoke (12).
Install disk (8) in counterbore in each end of each
trunnion spindle. Carefully push bearings over end
of trunnion spindle while primer is still wet. Align
holes in elastomeric bearings and shims with holes in
end of spindle. Remove two bolts (1) installed for
alignment and install retainer (3) on each elastomeric
bearing with four bolts (1) and washers (2). Do not
torque bolts at this time.
(m.1)
(9) as follows:
Bond elastomeric bearings (6) to trunnion
1 Abrade mating areas of trunnion (9) and
elastomeric bearing (6) with 400 grit abrasive cloth (C44).
2 Clean abraded areas of trunnion and
bearings with alcohol (C22).
3 Apply a heavy coat of prepared adhesive
(7B) to spindles of trunnion (9).
4 Allow adhesive to cure for 24 hours before
using hub assembly.
(n) Install two bolts (4) to secure each
elastomeric bearing (6) to the yoke (12). Install recessed
washers (5) to bolts (4) with recessed side toward bolt
heads. Install recessed washers (1) with recessed side
toward nuts (11). Torque nuts (11) 160 TO 190 inchpounds. If more than five threads show at nut, add a
steel washer under the nut.
(o) Torque bolts (1) 120 TO 160 inch-pounds
and lockwire (C151) in pairs.
(p) Apply a fillet of sealant (C116) around
trunnion spindles at inboard end of elastomeric
bearings.
(q) Check to ensure that hub is centered by
method shown on figure 5-34. This method, using a
six inch scale, is considered to be less accurate than
the method used in the preceeding steps but can be
used as a check.
(2) Place an adapter plate (T34) on build-up
bench (T26). Install the yoke and trunnion on the
build-up bench. See figure 5-35.
(3) Inspect both yoke extensions (28, figure 518) to ensure that radius rings (29) are installed and
are in satisfactory condition. Position yoke extension
(28) on yoke with web on leading edge side and
install bolt (27), special washers (26 and 39) and nut
(40) at this time. Install opposite yoke extension in
the same manner.
Change 48 5-53
TM 55-1520-234-23
Figure 5-35. Tool application-grip spacing adjustment
install pin through yoke extension and strap. Install
retaining ring (34). Coat both ends of pin with sealant
(C116) described in step (4). Install opposite strap in
same manner.
Main rotor hub retention straps P/N 204-0121223 and-7 are spare part replacements for
straps P/N 20012- 112-7. Replace straps in
pair Do not intermix P/N 204-012122-3 straps,
P/N 204-01122-7
straps or 204-012-112-7
straps in the same main rotor hub.
(4) Position strap (54, figure 518) in fit ting
(52). Install retaining ring (59) on pin (53). Install pin
through fitting and strap. Install retaining ring (53).
Coat ends of pin with sealant (C116). Assemble
opposite strap and fitting in same manner.
(6) Hinge yoke extension (28, figure 5-18)
forward on bolt (27). Inspect housing (38) to ensure that a
serviceable bearing (37) and seal (36) are properly
installed. Position housing (38) on yoke extension: Hinge
the yoke extension back into position and install bolt (27),
special washers (26 and 39) and nut (40) in trailing edge
hole. torque nut (40) 450 TO 550 foot-pounds. Install
opposite housing in the same manner.
(7) Inspect grip (41, figure 5-18) to ensure that
strap indexing ring (49), bearing (50) and dust seal (51)
are properly installed. Apply a coating of sealant (Cl16) to
both slots in fitting (52) so that this area will be sealed
when the grip
(5) Insert assembled strap (54, figure 5-18)
and fitting into outboard end of yoke extension (28).
Install retaining ring (31) on pin (33) and
Change 48 5-54
TM 55-1520-234-23
is installed. Slide the grip (41) on the grip extension
with the side with provisions for mounting the pitch horn
on the trailing edge side. Engage the lugs on the strap
indexing ring in the grip with the slots in fitting (52).
Work grip onto extension far enough to expose three
threads on fitting (52). Use a fiber mallet to tap grip onto
extension. Install washer (46) and start dome nut (45).
Install two bolts (60) through grip and housing. Install a
maximum of four steel washers (47) on each bolt as
required and install nuts (48). Torque nuts 770 TO 950
inch-pounds. Install opposite grip in same manner.
Rotate grips gently through their full travel, ensuring
there is no interference between grip and extension.
NOTE
Check clearance of extension and grip
assembly. Refer to paragraph 5-6a (11).
CAUTION
Do not install washers under heads of bolts
(60).
(8) If drag brace (15, figure 5-18) was
disassembled, install nuts (14) and clevis ends on drag
brace. Adjust clevis until approximately 0.25 inch of
threads are exposed on each end and the dimension
between centers of clevis holes is 14.732 inches.
Tighten nuts (14) snug but do not torque. Position drag
brace on grip and install bolt (18), washers (16) and nut
(17). Do not torque nut (17) at this time. Install
opposite drag brace in the same manner.
CAUTION
Only bolts, nuts, washers, and bushings that
have been degreased will be used to attach
pitch horn to grip. During reassembly of
pitch horn to grip care should be taken to
ensure that the bolts, nuts, washers, and
bushings are clean and free of any lubricant
other than the dry film lubricant.
Change 72
NOTE
Bolt 209-010-112-1 may be used in place of bolt
209-010-112-3.
NOTE
If new bolts are not available, old bolts may be
reused after disassembly of pitch horn from
grips. Reused bolts must be degreased and
pass visual and magnetic particle inspection
prior to reinstallation. Bolts, nuts, washers,
and bushings will be vapor degreased using
PD680, Type 2. If vapor degreasing equipment
is not available, soak in PD680. Type 2 for one
hour and wipe off grease and PD680 using a
lint free cloth prior to installation. Use of
alternate cleaning agents is not recommended
as it may remove dry film lubricant protective
coating.
(9) Install bushings (20 and 22, figure 5-18) in pitch
horn (23). Position pitch horn on grip and install special
bolts (19 and 21). Install the longer bolts in the inboard
holes. Install a maximum of two washers (5) under each
nut (4). Torque nuts (4) to 700 TO 725 inch-pounds. Tap
the pitch horn and grip around the bolted area to set the
parts. Use a rawhide or non-metal hammer. Retorque to
700 TO 725 inch pounds. Fly main rotor head for one hour
and retorque to 700 TO 725 inch-pounds. Install opposite
pitch horn in the same manner.
NOTE
The pitch horn bolts are properly installed
when tapered shoulders on bolts are seated in
bushings in the pitch horn. The bolt heads will
not be in contact with the pitch horn or
bushings.
5-54A/(5-54B blank)
TM 55-1520-234-23
(10) Replace tape (C136) on inboard spacer (3,
figure 5-18) if required to obtain a snug fit. Position two
spacers (3) in web of yoke extension. If previously
removed, position sand deflector (1) on yoke extension
and install three bolts (2) through sand deflector and
spacers. Ensure that there is adequate clearance
between deflector and yoke (25) at both upper and lower
surfaces. Install opposite sand deflector in the same
manner.
(11) If grip spacing tool is available, proceed as
follows:
(a) Install bolt assembly (6, figure 5-18) in each
grip.
(d) Tighten dome nut (45, figure 5-18), which was
installed in step (7), until dimension between radius ring
(29) and dust seal (51) is 0.001 inch. See dimension "A" on
figure 5-36. Adjust opposite grip in the same manner.
(e) Install grip spacing tool (T56). See figure 5-35 for
view of installed grip spacing tool. Adjust tip, T101559-3,
on gage, T101559-5, to 2.0 inch dimension as illustrated.
Install plug, T101559-3, above rotor hub trunnion and
secure with knurled screw. Locate hole marked "540-011101" on gage of spacing tool (T56), and attach gage to plug
with bolt through this hole. Raise blade bolt and position tip
of spacing tool (T56) so that it rests on dowel pin in grip as
shown. Measure and record distance between blade bolt
and tip of grip spacing tool. Reverse grip spacing tool and
measure distance on opposite blade bolt.
(b) Install two flap stops, (T42), on trunnion with
540 side down as shown on figure 5-35. Use 3/8 inch
UNF threaded bolts of suitable length to secure flap
stops to trunnion.
CAUTION
Insure that during all grip spacing procedures,
the grips are seated against dome nuts. To
insure proper seating, back off dome nut one
full turn in excess of that required for
adjustment while tapping grip outboard with a
fiber mallet. Then, turn dome nut clockwise
while observing the dimension to insure grip is
properly seated against the dome nut.
If
dimension does not close, grip is not properly
seated.
(c) Remove one bolt (4, figure 5-19) from each
trunnion bearing and install two grip locks (T59) as
shown on figure 5-34.
CAUTION
Insure that during all grip spacing
procedures, the grips are seated against
dome nuts. To insure proper seating, back
off dome nut one full turn in excess of that
required for adjustment while tapping grip
outboard with a fiber mallet. Then, turn
dome nut clockwise while observing the
dimension to insure grip is properly seated
against the dome nut. If dimension does not
close, grip is not properly seated.
(f) Loosen dome nut (45, figure 5-18) on grip (41) that
was found to be most inboard in the preceding step. Adjust
this dome nut as required until the dimension is equal to
that of the most outboard grip within 0.002 inch.
Change 72
5-55
TM 55-1520-234-23
Figure 5-36. Main Rotor hub - grip dust seal to radius ring
5-56
TM 55-1520-234-23
(g) Check seal gap dimension "A" shown
on figure 5-36 on both grips. Dimension "A" must be
0.001 to 0.040 inch.
NOTE
The tolerance dimension "A" can be eased
up to 0.060, provided an assurance check
will be made anytime the grip spacing
exceeds 0.040. Correct procedures and seal
to radius ring will be verified.
(h) Install lock (44, figure 5-18), clamp (43),
bolt (42), thin steel washer (56), nut (55), and cotter pin
(57).
(i) Remove grip spacing tool.
(j) Install blade bolt (6, figure 5-18) keyway
washer
(7),
extended
washer
(8),
special
Change 29
5-56A/ (5-56B blank)
TM 55-1520-234-23
nut (9), screw (12), washer (11), and nut (10). Do not
torque special nut (9).
NOTE
(k) Remove two flap stops (T42) that were
installed in step (b).
(l) Remove two grip locks (T59) that were
installed in step (c). Install bolts (4, figure 5-19) with
recessed washers (5). Install washer with recessed side
toward bolt head. Install recessed washer (10) with
recessed side next to nut (11). Use more than one
special washer (10) to obtain proper engagement of nut
if necessary. Torque all four nuts (11) 160 TO 190 inchpounds.
Pot life of adhesive is 110 to 130 minutes.
(e) Fair out adhesive.
Remove excess
adhesive using cheesecloth (C36) dampened with methylethyl-ketone (C87).
(f) Maintain firm contact pressure and cure for
24 hours at 70°F (21°C) or 30 to 60 minutes at 200°F
(93°C). Edge voids are not allowed. Maximum strength
achieved in 6 to 7 days.
(13) Balance main rotor hub assembly (paragraph
CAUTION
After nuts (11) are tightened, no more
threads of bolts are permitted to be
beyond nuts (11), and a minimum
threads must be exposed to ensure
locking feature of the nuts is engaged.
5-6A).
than five
exposed
of three
the self-
5-6A.
Balancing - Main Rotor Hub Assembly.
NOTE
Refer to TM 55-4920-201-15 for additional
information on balancing tools if required.
(12) Installation of identification plate.
CAUTION
Stamping directly on the surface of any detail
hub part or installed data plate is prohibited.
(a) Stamp applicable data on replacement
plate using 1/16 inch characters.
(b) Lightly sand contact area on component
and mating surface with No. 180 grit abrasive cloth
(C44).
Before using naphtha or methyl-ethyl-ketone,
extinguish all open flames and turn off electrical
equipment.
Vapors are highly flammable.
Avoid prolonged breathing of vapors and
repeated skin contact. Use in well ventilated
area.
(c) Remove
sanding
residue
using
cheesecloth (C36) dampened with naphtha (C88).
(d) Mix EC2216 adhesive (C11) 100 parts
base to 140 parts hardener. Apply adhesive within 20
minutes to both mating surfaces and join parts.
(1) Set up the hub balancing stand and
accessories from balance kit (T80) as follows: See figures
5-37 and 5-38.
(a) Assemble hoist support structure with tube
assembly P/N 2769 instead of tube assembly P/N 2288
shown on figure 5-37 to provide additional hoist arm height.
(b) Center fixture (1, figure 5-38) from kit (T80)
on work stand.
(c) Install adapter (2), heavy end downward,
over top of fixture (1) and seat on upper shoulder of fixture
central projection.
Lock adapter in this position by
tightening adapter setscrew (3) using 1/8 inch hex wrench
(T79, T80 and T81 kits.)
(2) Balance main rotor hub assembly as follows:
See figure 5-38.
(a) Carefully lower rotor hub assembly (8) over
fixture (1); align inside diameter of splined trunnion with
piloting diameter of adapter (2), and ensure that cone
surface of splined trunnion seats firmly on cone surface of
adapter (2).
(b) Install yoke (4), legs downward, on arbor
(5) and position so that top surface of its locking collar
sensitivity setting reference (figure 5-38) aligns with 15-3/8
Change 56
5-57
TM 55-1520-234-23
Figure 5-37. Rotor balancing kit P/N 7A050
5-58
TM 55-1520-234-23
inch position on arbor scale (6). Lock yoke firmly in this
position on arbor with its collar screw, using 3/16 inch
hex wrench from kit (T81).
(c) Install arbor downward through rotor
trunnion and fixture assembly. Seat legs of yoke in
milled areas on top surfaces of hub yoke; center with
scribed lines.
(d) Position jacks (7) on top surface of the
rotor hub yoke so that their inboard ends bear against
the central boss of the hub yoke, centered below the
scribe lines mentioned in paragraph (4), and their
outboard ends bear centrally against the shoulders of
the inboard bearing housing of the blade grip
assemblies. Adjust jacks to provide uniform outward
pressure sufficient to ensure blade grips are seated in
their full outward positions.
(e) Install spacer (13) over lower end of
arbor; install handwheel (14) in lower end of arbor and
tighten to clamp both legs of yoke firmly against top
Change 56
surfaces of hub yoke.
(f) Using gage (10) as shown, adjust drag
braces (9) to symmetrical angular positions. Remove gage
from rotor hub during subsequent balance check.
(g) Install quick-disconnect assembly with 3/16
inch cable from kit (T79) on arbor suspension rod and hoist
balancing assembly approximately 1/4 inch off work stand
with hydraulic pump (15). Check to ensure that suspended
assembly is free from interference with work stand and
adjacent objects, and note balance condition indicated at
top end of arbor. (See figure 5-39.)
NOTE
In order to ensure that the handwheel P/N 2315 suspends
free of interference within the inside diameter of the stand
table, it may be necessary to adjust the level of the stand
assembly by installing suitable blocks under the two tubular
stand legs (6, figure 5-37).
5-58A/(5-58B blank)
TM 56-1520-234-23
Figure 5-38. Tool application - main rotor hub balancing
5-59
TM 55-1520-234-23
Figure 5-39. Interpretation of balancer indication
(h) After it is determined that the
handwheel, P/N 2315, suspends free of interference,
lower the hub to rest on the stand.
accessories.
(i) Use a protractor on the machined
surface next to the blade retaining bolt and set both
blade grips to zero degrees. Both grips must be equal
within zero degrees, five minutes.
The scissors and sleeve assembly is a component of
the mast controls. See figure 5-40. The scissors are
attached to the swashplate by the drive links for cyclic
control of the main rotor. The scissors are attached to the
collective control system through the collective sleeve (18)
and levers (10) for collective pitch control of the main rotor.
(j) Raise assembly approximately 1/4 inch
off work stand to obtain balance readings.
5-7.
(k) Balance the hub chordwise within 12
inch-pounds. Place weight on the light pitch horn at hub
station zero to obtain balance within limits.
Chordwise balance is used as an aid to spanwise
balance.
Remove weight from pitch horn after
completion of spanwise balance.
(l) Balance spanwise within one half inchpound about blade station 0.000, by inserting lead wire,
lead wool slugs (C82) or 0.44 inch diameter shot into
cavity of blade bolt assembly (16, figure 5-38).
Scissors and Sleeve Assembly
Premaintenance Requirements for Scissors and Sleeve
Assembly
Condition
Model
Part No. or Serial No.
Special Tools
(T35) (T48)
Test Equipment
(m) Install a plug (17) into blade bolt (16)
when balance has been accomplished.
(n) Color band hub assembly parts after
balance to ensure that parts of hub remain in same
respective position as they were when hub was
balanced.
(o) Remove
hub
balance
stand
and
Support Equipment
Minimum Personnel
Required
Consumable Materials
Special Environmental
Condition
5-60
Requirements
AH-1S
All
(T2) (T27) (T28) (T29)
Force Gauge (fish scale)
capable of measuring up
to 150 pounds
Dial Indicator
NA
2
(C44) (C45) (C70) (C87)
(C102) (C124) (C128)
None
TM 55-1520-234-23
Figure 5-40. Mast controls installation (Sheet 1 of 4)
5-61
TM 55-1520-234-23
Figure 5-40. Mast controls installation (Sheet 2 of 4)
5-62
TM 55-1520-234-23
Figure 5-40. Mast controls installation (Sheet 3 of 4)
Change 38
5-63
TM 55-1520-234-23
Figure 5-40. Mast controls installation (Sheet 4 of 4)
5-64
Change 71
TM 55-1520-234-23
a. Removal. See figure 5-40.
CAUTION
Remove scissors-and sleeve with caution to
avoid damage to mast.
(1) Remove main rotor. Refer to paragraph 5-4.
(2) Determine whether the extension installed
above the scissors and sleeve assembly is P/N 540-011487-1 as shown on detail view A or P/N 209-010-464-1
as shown on Detail view B and remove parts as
described in step (a) or step (b) as applicable.
(a) Helicopters with extension P/N 540-011487-1:
1 Cut lockwire and remove spacer (25)
and upper boot (26).
2 Remove bolts and remove
assembly (29).
clamp
3 Remove rubber ring (30).
4 Cut lockwire and remove spring pin
(37). Use spanner wrench (T2) to loosen threaded ring
(31). Remove the threaded ring and collet set (32).
Identify collect set for reinstallation as a set.
(a) Attach dial indicator on mast as shown on
figure 5-41 with indicator probe against flat of one of the
attachment bolts.
(b) Measure and record amount of radial play by
rotating scissors and sleeve assembly hub (6) forward and
then back to spline contact. Maximum allowable amount of
radial play, measured in this manner, is 0.040 inch.
(4) If scissors and sleeve assembly is to be
reinstalled without complete disassembly and inspection,
check wear on thrust washers (6, figure 5-42) prior to
removal. Maximum allowable play at thrust washers is
0.060 inch as shown on illustration.
(5) Remove bolts (36, figure 5-40) or bolts (56) as
applicable and remove extension and spline plate. Identify
spline plate as satisfactory or as worn beyond limits noted
in preceding step.
(6) Disconnect collective system control tube from
collective lever assemblies (7 and 10, figure 5-40).
Remove bolts (1, 11, and 12). Separate collective lever
halves (7 and 10) from collective sleeve (18) and link (21).
Keep spacer (9), thrust bearing washer (14), thrust washer
(17), be ring inner race (23) and similar parts on the
opposite side with the collective lever halves for
reassembly.
(b) Helicopters with extension P/N 209-010464-1:
1 Cut lockwire and remove spacer (40)
and upper boot (41).
2 Remove
assembly (44).
bolts
and
remove
clamp
3 Remove rubber ring (47).
4 Remove nuts (53), washer (54), and
bolts (48. Remove retainer (49) and collet set (50).
Identify collet set for reinstallation as a set.
(3) Check wear on spline plate prior to removal.
See figure 5-41.
NOTE
The procedure for checking wear on spline
plates is the same for installations with
extensions P/N 540-011-487-1 and extensions
P/N 209-010-464-1.
Figure 5-41. Tool application - drive plate spline wear
measurement
Change 38
5-65
TM 55-1520-234-23
CAUTION
Do not allow scissors lever to contact scissors hub
as damage to lever could result. Block scissors
lever with wood or other suitable material to
prevent damage.
(10) Lift scissors and sleeve assembly (57, figure 540) out of swashplate and off mast. Use caution to prevent
damage to friction sleeve and mast splines during removal.
(11) If swashplate is not to be removed, cover open
area around top of lower boot (60, figure 5-40) to prevent
entry of foreign materials.
(12) If scissors and sleeve assembly is to be
reinstalled without complete disassembly and inspection,
make the following inspections to ensure that parts are
suitable for reinstallation on helicopter.
(a) Check end play between scissors and
sleeve assembly (57, figure 5-40) and link (58) for
maximum axial looseness of 0.090 inch.
(b) Upper and lower boots (41 and 60, figure 540) for tears and deterioration.
(c) Rubber ring
deterioration and damage.
(47,
figure
5-40)
for
(d) Collet (50, figure 5-40) for missing fingers,
cracks, scoring, or other damage.
Figure 5-42. Collective lever thrust washer wear
limits
(e) Bearing assemblies (19, figure 5-40) for
binding, roughness and maximum radial play of 0.010 inch.
(7) Remove screws (13, figure 5-40) bearing
assembly (19) and spacer plate (20). Remove similar
parts from opposite side.
(8) Remove cotter pin (72, figure 5-40), nut (71),
and special washer (70). Remove drive link from
swashplate. Remove special washer (69). Remove
opposite drive link in the same manner.
(f) Inspect drive link (58, figure 5-40),
spherical bearing for roughness, binding, axial play,
slippage mark, and alignment. A maximum of 0.015 inch
axial play is permissible if excessive vibration does not
occur.
(9) Cut lockwire and detach lower boot (60,
figure 5-40) from collective sleeve.
The spherical bearing wear will be in both the radial
and axial direction. However, the only criteria
necessary for determining serviceability will be to
measure the axial play.
5-66
NOTE
Change 54
TM 55-1520-234-23
(g) Scissors levers for gouges
scratches especially on underside of pivot leg.
figure 5-44.
and
See
(h) Check clevis end of scissors, outboard
of bushing, if recess between bushing end and clevis
outside surface exceed 0.004 inch a shim is required,
refer to paragraph 5-7.e.
(6) (g) for installation
procedures.
(i) Swashplate horns for scoring.
(i) Spline plate wear in excess of 0.040
inch limit measured in step (3).
(k) Inspect clamp assembly (29 or 44)
segments for cracks, corrosion/mechanical damage
deformation and/or elongation of bolt holes. Replace if
any of the above conditions exist.
(l) Measure clearance between
bushing and head of boss bushing (82).
allowable clearance is 0.0615 inch, and
between end of boss bushing and link,
allowable
clearance
is
0.0595
link (86)
Minimum
clearance
minimum
inch.
Change 71 5-66A/(5-66B blank)
TM 55-1520-234-23
b. Disassembly. See figure 5-43.
set (46) by pressing out sleeve assembly.
(12) Remove loose nut, two retaining rings and
boot support ring from sleeve assembly. Use adapter (T28)
to remove seal (43) from nut (44).
CAUTION
Inspect scissors and sleeve and scissors and sleeve
historical records for evidence that the assembly has
been subjected to an accident or incident such as an
overtorque. If scissors and sleeve assembly has been
subjected to an accident or incident outside the realm of
normal usage, perform conditional inspection of step d.
(1) prior to disassembly.
(13) Remove retaining rings (37 and 39) and
bearing sleeve (38) from collective sleeve (36).
c.
Cleaning. (AVIM)
(1) Remove bolt (30) and remove link (32).
Retain shim (17) for reassembly.
(2) Remove housing (18), washer (19) and
inner race (20) from scissors.
(3) Remove opposite scissors and link in the
same manner.
(4) (AVIM) Remove nut (29), bolt (9) and
washers (10, 25, 26 and 27) and remove scissors
assembly (16) from hub.
(5) Remove inner races (11 and 3) and spacer
(4).
(6) Install wrench (T29) on top of hub (2) with
two bolts. Invert assembly and secure wrench in a vise.
(7) Remove two screws (8) and lock plate (7).
Disengage spiral retaining rings (40 and 42) and move
spacer (boot support ring) (41) away from mounting
shoulder for access to bottom of hub.
(8) Use wrench (T48) to turn nut (44) out of
hub. Remove assembly from vise and remove tools.
(9) Place sleeve assembly on a press with
halves of support (T27) placed under hub. Insert small
end of ram adapter (T28) in top of sleeve. Press sleeve
assembly out of hub. Remove seal (1) from hub.
Remove spacer ring (49) from bearing stack on sleeve.
(10) Remove lockwire and pin (50). Install
wrench (T29) with pins engaged in holes of nut (48).
Insert bar (T35) in holes at lower end of sleeve and hold
against turning while removing left-hand threaded nut.
Remove tools.
Use solvent (C124) in a well ventilated area. Avoid
prolonged breathing of vapors and do not use in an area
with open flame or high temperature.
(1) Clean scissors and sleeve assembly parts,
drive links, mast friction parts, spline plate collective lever
halves, link and support with solvent (C124).
(2) Dry parts with dry, filtered, compressed air. Do
not allow bearings (45 or 47, figure 5-43) to spin while
drying.
(3) Protect
bearings
(45
and
47)
contamination. Keep bearings together in sets.
from
d. Inspection. (AVIM)
(1) Visually inspect visible parts of scissors hub
assembly and boat for signs of heat. Any heat discoloration
or distortion of components is cause for replacement.
(1a) If the assembly has been involved in an
accident or incident, perform conditional inspection of
scissors and sleeve as follows.
NOTE
The scissors and sleeve assembly should not have
been disassembled if a conditional inspection is
required. Refer to step b.
(a) Carefully
inspect
the
assembled
component visually for apparent damage and for abnormal
appearance. Obvious defects which are cause to scrap the
entire assembly are:
(11) Place sleeve assembly on a press with
support halves (T27) placed under inner race of lower
bearing. Remove bearing sets (47 and 45) and spacer
Change 50
1
5-67
Severe binding in any of the pivot joints.
TM 55-1520-234-23
Figure 5-43. Scissors and sleeve assembly (Sheet 1 of 2)
5-68
TM 55-1520-234-23
Figure 5-43. Scissors and sleeve assembly (Sheet 2 of 2)
2
Severe binding between the hub and
(6) Bearing assemblies (19, figure 5-40) for
binding, roughness and radial play in excess of 0.010 inch.
(b) Inspect components for surface damage
in accordance with normal inspection paragraph.
Surface damage in excess of established limits will
require scrapping only the damaged part.
(7) Inspect hub, sleeve, scissors and link for
corrosion and mechanical damage in excess of limits
shown on figure 5-44.
sleeve.
(c) Position an unworn bolt (30, figure 5-43)
in link (32). If the bolt does not fit freely through
bushings, scrap the link. Check the opposite link in the
same manner.
(8) Inspect drive link (32, figure 5-43), spherical
bearing for roughness, binding and axial play. A maximum
of 0.01 5 inch axial play is permissible if excessive
vibration does not occur.
NOTE
The spherical bearing wear will be in both the radial
and axial directions. However, the only criteria
necessary for determining serviceability will be to
measure the axial play.
(d) With a straight edge, check the
cylindrical portion of the collective sleeve for
deformation. If warpage is in excess of 0.005 inch in a
5.0 inch length, scrap the sleeve.
(e) Check all machined
flat surfaces
surrounding lugs, holes and bushings for deformation
with a straight edge. If deviations from flat in excess of
0.002 inch are found, scrap the part.
(f) Inspect clevis end of scissors, outboard
of bushing. If recess between bushing end and clevis
outside surface exceeds 0.004 inch a shim is required,
refer to paragraph 5-7.e.
(6) (g) for installation
procedures.
(2) Identify scissors and sleeve components
which will reach retirement time prior to next scheduled
inspection for replacement. Refer to overhaul and
retirement schedule.
(3) Inspect upper and lower boots (26, 41, and
60, figure 5-40) for damage and deterioration.
(9) Inspect bolts (9 and 30, figure 5-43), washers
(4, 19, 25, and 27) and inner races (3, 11 and 20) for
damage. If other than a smooth, unscored surface is found,
replace affected part. Maximum lateral chucking of the
scissors lever (16) will be 0.020. There will be no
longitudinal chucking permissible of bolt (9).
(10) Inspect bearing sets (45 and 47, figure 5-43)
as follows:
(a) Inspect for roughness and/or brinelling.
Reject bearings with brinnelling damage that is visible
under 5 power magnification.
(b) Inspect for galled or flaked areas on balls
and raceways. Use a strong light when making this
inspection.
(c) Inspect retainers (6 and 21, figure 5-4:3) for
(4) Inspect rubber ring (30 or 47, figure 5-40)
for damage and deterioration.
damage.
(5) Inspect collect (32 or 50, figure 5-40) for
missing fingers, cracks and scoring.
Change 65
5-69
TM 55-1520-234-23
Figure 5-44. Damage limits - hub, sleeve, scissors and link (Sheet 1 of 4)
5-70
Change 7
TM 55-1520-234-23
Figure 5-44. Damage limits - hub, sleeve, scissors and link (Sheet 2 of 4)
Change 7
5-71
TM 55-1520-234-23
Figure 5-44. Damage limits - hub, sleeve, scissors and link (Sheet 3 of 4)
5-72
Change 7
TM 55-1520-234-23
Figure 5-44. Damage limits - hub, sleeve, scissors and link (Sheet 4 of 4)
Change 7
5-73
TM 55-1520-234-23
(d) Inspect bearing races for damage.
(11) Inspect nuts (44 and 48, figure 5-43) for
damage with special attention to threads.
(12) Inspect the following parts by magnetic
particle method, code M, per MIL-16868 or fluorescent
penetrant method, code F, per MIL-1-6866. Items are
indexed to figure indicated.
FIGURE
ITEM NOMENCLATURE CODE
5-43
5-43
5-43
5-43
5-43
5-43
5-43
5-43
5-40
5-40
5-40
16
9
30
32
36
44
2
48
34 or 52
7 and 10
86
Scissors
Bolt
Bolt
Link
Sleeve, Collective
Nut
Hub
Nut
Spline plate
Collective Lever
Idler Link Assembly
F
M
M
F
M
M
M
M
F
M
F
(13) Inspect lockplate (7, figure 5-43) for broken
or deformed tangs.
(14) Inspect spacer sets (41 and 46, figure 5-43)
for corrosion, scoring and other mechanical damage.
No repair authorized.
(15) Inspect bolts (9 and 30, figure 5-43), inner
races (3, 11, and 20), and thrust washers (4, 19 and 25)
for scoring. No repair is authorized. If bearing inner
race outside diameter is scored, the race and the
matching bearing (5, 15, 22, or 24) must be replaced.
(16) Visually inspect sleeve (38, figure 5-43) for
indications of wear at contact points with retaining rings
(37 and 39).
(17) Visually inspect cap washer (26, figure 543), housing (18), retaining rings (37, 39, 40, and 42),
spacer (41), lock plate (7), pin (50), spacer ring (49) and
spacer (46) for cracks, corrosion and deformation.
(a) Inspect pin (50) for security and
presence of lockwire.
(b) Inspect two screws (8) and lockplate (7)
for security and presence of lockwire.
5-74
(c) If pin, screws or lockplate are loose, remove
(paragraph 5-7a), disassemble (paragraph 5-7b) and clean
(paragraph 5-7c).
(18) Inspect spline plate for damage in excess of
limits shown on figure 5-45.
(19) Inspect collective lever halves (7 and 10,
figure 5-40) for damage in excess of limits shown on figure
5-46.
(20) Inspect collective lever idler link for the
following:
(a) Damage in excess of limits shown on figure
5-47.
(b) Bearings for elongation looseness and
damage. Maximum allowable radial play is 0.010 inch.
(c) Rubber elements in lower end of link for
deterioration and separation.
(21) Dimensionally inspect the following parts if
visual inspection indicates wear. Items are indexed to
figure 5-43.
ITEM
NOMENCLATURE
36
2
2
Sleeve, Bearing Seat OD
Hub Bearing Seat ID
Hub Pivot Bolt Hole ID
REPLACE AT
4.248 Min.
5.2520 Max.
0.7520 Max.
(22) Refer to paragraph 5-8 for inspection and
repair procedures for swashplate anti-drive link (64, figure
5-40), bellcrank (75) and support (79).
(23) Inspect Collet sets for missing fingers, cracks,
and scoring. Inspect fingers for missing and badly worn
teflon.
e. Repair. (AVIM)
CAUTION
Repair by use of grinding wheel is not allowed.
(1) Polish out corrosion and mechanical damage in
parts inspected in preceding step. Use fine to medium
abrasive cloth (C44), crocus cloth (C45) or fine India stone
(C128). Blend repair smoothly into surrounding area.
Replace part if repair exceeds allowable area and/or depth
limits.
Change 65
TM 55-1520-234-23
(2) Replace defective scissors hearings (22 and
24, figure 5-43) as follows:
(a) Insert a punch through retainer (21) and
tap out bearings (22 and 24) and spacer (23). Inspect
retainer ID for scoring. Maximum allowable depth of
score marks after clean up is 0.002 inch. Clean retainer
and apply a light film of grease (C70) to retainer bearing
bore.
Change 65
5-74A/(5-74B blank)
TM 55-1520-234-23
Figure 5-45. Damage limits - spline plate
Change 7
5-75
TM 55-1520-234-23
Figure 5-46. Damage limits - collective lever
Use solvent in a well ventilated area. Avoid
prolonged breathing of vapors and do not
use in an area with open flame or high
temperature.
(b) Clean bearings with solvent (C124) and
hand pack bearings with grease (C70).
(c) Press bearing (22) into retainer (21) and
then press spacer (23) into retainer.
(d) Press bearing (24) into retainer with the
seal side of the bearing facing outboard.
(3) Replace defective bearing (5, figure 5-43.)
Replacement procedures for bearing (5) are the same
as bearings (22 and 24, step 2 above).
5-76
Use solvent in a well ventilated area. Avoid
prolonged breathing of vapors and do not use
in an area with open flame or high temperature.
(4) Replace defective bearing (15, figure 5-43).
Position scissors on a suitable support and apply pressure
to bearing race. Inspect sleeve (14) for scoring. Score
marks must not exceed 0.002 inch depth after cleanup.
Clean new bearing with solvent (C124) and hand pack
bearing with grease (C-70 ). Apply heat lamp to scissors
and press in new bearing (15). Check that bearing ends are
equally spaced on each side of scissors tang.
(5) Protect bearings from contamination until
reassembly.
Change 7
TM 55-1520-234-23
Figure 5-47. Damage limits - collective lever idler link
(6) Replace damaged or worn bushings (12 and
13, figure 5-43) as follows:
(a) Support the clevis end of scissors so as
to prevent a bending load on tang, and press defective
bushing from scissors.
(b) Inspect bushing bores in scissors for
scoring after bushings are pressed out. Polish out scoring
type damage to 0.002 inch maximum depth. Clean holes in
clevis and clevis inboard faces with solvent (C124).
(c) Coat bushing OD with wet, unreduced
primer (C102). Ensure that primer does not contact buffer
material on underside of bushing flange. Apply a heat lamp
to end of scissors and press bushings into clevis holes with
bushing flanges facing inboard. Check that bushings are
fully seated.
(d) Line ream bushings to 0.6250 TO 0.6255
inch diameter.
Use solvent in a well ventilated area. Avoid
prolonged breathing of vapors and do not use in
an area with open flame or high temperature.
Change 7
5-77
TM 55-1520-234-23
(e) Mill inboard bushing faces to establish
1.250 TO 1.252 inch dimension between faces. Keep
flange wall thickness equal within 0.005 inch.
(f) Chamfer the edge I.D. on inboard side
of bushings 0.005 TO 0.010 x 45 degrees.
(g) When recess exceeds 0.004 maximum
(figure 5-47A) between end of the bushing {items 12, 13,
figure 5-43} and outboard surface of the scissors, install
a 120-006C27E21 shim and peel as required to meet
0.002 to 0.004 maximum recess. Bond shim in place on
end of bushing using a cyanoacrylate adhesive (C14A).
Figure 5-47B. Collective Lever Idler Link Assembly
(a) Replace damaged lubrication fitting {7,
figure 5-47B) as follows:
1 Carefully remove old lubrication fitting
to avoid damage to link.
Figure 5-47A. Installation of Shims
(7) Replace components of collective lever idler
link (86, figure 5-40) which have damage in excess of
limits specified in inspection paragraph 5-7d(20) as
follows:
5-78
2 Press new lubrication fitting into link.
3 Attach a grease gun serviced with
clean grease (C70) to fitting and check to ensure that fitting
is properly installed
Change 38
TM 55-1520-234-23
(b) Replace damaged identification plate
(2) as follows:
NOTE
If data to be stamped on identification plate is
not available, send affected assembly to Depot
Maintenance for evaluation.
1 Stamp all data from the
identification plate on the new identification plate.
old
8 Clean adhesive squeeze-out from parts
with cheesecloth and MEK (C87) before adhesive cures.
9 Remove masking tape before adhesive
cures.
10 Allow adhesive to cure for 24 hours at
room temperature (approximately 75 degrees F) (24
degrees C). Full strength will be reached in six to seven
days.
(c) Replace damaged bushing (6) as follows:
CAUTION
Do not heat link assembly (1) to temperature above
limits noted in following step.
CAUTION
Do not heat link assembly (1) to temperature
above limit noted in following step.
2 Remove old identification plate from
link assembly. Heat link assembly to 200 ± 15 degrees
F (93 ± 9 degrees C) to loosen adhesive.
3 Mask off face side of new identification plate and the
area of the link assembly where identification plate will
be installed.
4 Clean masked-off area on link
assembly with 300 grit or finer sandpaper (C112A).
5 Form new identification plate to fit
closely on link assembly.
1 Support link assembly with suitable
sleeves and supports to avoid distortion and press out two
bushings (6). The bushings must be pressed outboard from
the link as illustrated. If bushings are a tight fit in the link,
to a maximum
of 200 ±15 degrees F (93 ± 9 degrees C). Then press
bushings out.
2 Clean bores where two bushings (6) will
be installed.
3 Select suitable sleeves and support
blocks to support legs of link assembly (1) during
installation of bushings (6).
CAUTION
Do not heat link assembly (1) to temperature above
limited noted in following step.
Cleaning solvent is flammable and toxic.
Provide adequate ventilation. Avoid prolonged
breathing of vapors and contact with skin or
eyes.
6 Scrub
mating
surfaces
of
identification plate and link assembly with cheesecloth
(C36) dampened with aliphatic naptha (C88). Wear
clean white cotton gloves (C66B) when handling parts
after cleaning and prior to bonding.
7 Mix two-part adhesive (C11) in
accordance with instructions on container. Apply a thin
coat of adhesive to each of the mating surfaces as soon
as possible after mixing. Place identification plate on
link and anchor in position with clamps or rubber bands.
Do not chill bushings (elastomeric-type bushings) (6) during
installation procedure.
4 Heat link assembly (1) to 200 ± 15
degrees F (93 ± 9 degrees C). Coat mating surface of
bushings (6) with primer (C100 or C102) and install with
flanges outboard as illustrated
while primer is wet.
(d) Replace damaged bearings (3) and (5) as
follows:
Cleaning solvent is flammable and toxic. Provide
adequate ventilation. Avoid prolonged breathing of
vapors and contact with skin or eyes.
Cleaning solvent is flammable and toxic.
Provide adequate ventilation. Avoid prolonged
breathing of vapors and contact with skin or
eyes.
Change 38
5-78A
TM 55-1520-234-23
1 Clean bore of link assembly (1)
where bearings will be installed. Clean new bearings (3
and 5) and sleeve (4) with dry cleaning solvent (C124).
Allow bearings to dry thoroughly and hand pack with
grease (C70).
2 Apply a thin coat of corrosion
preventive compound (C53) to mating surfaces of
sleeve (4) and bearings (3 and 5). Press spacer and
bearings into link.
3 Polish out mechanical and corrosion
damage that is within limits shown in figure 5-47.
f.
Assembly. (AVIM)
(1) Lubricate bearing sets (45 and 47, figure 543) as follows:
(4) Assemble bearing sets (45 and 47) and spacer
(46) according to etched numbers and V-mark on outer
races of bearings. Use adapter (T28) to press bearing
stack on upper end of sleeve, with V-mark pointing up.
(5) Start left-hand threaded nut (48) on sleeve.
Install wrench (T29.3) with pins engaged in holes of nut.
Hold sleeve with bar (T34.1) through holes at lower end.
Apply maximum torque, 200 foot-pounds, to nut (48). Allow
the stack-up to set for 10 minutes, release the torque, and
then torque to 150 foot-pounds.
Increase torque as
needed, maximum 200 foot-pounds, to align a hole in the
nut with a hole in the collective sleeve. Do not loosen the
nut, reduce torque, to align the holes. Install pin (50).
Remove tools. Secure pin with lockwire; insert lockwire
through drilled head and twist in space between sleeve and
nut.
(6) Use adapter (T28) to press seal (1), with lip
upward, into lop of hub) until bottom of seal is flush with or
slightly below lower edge of hub seal bore.
Use solvent in a well ventilated area. Avoid
prolonged breathing of vapors and do not
use in an area with open flame or high
temperature.
(a) Wash bearing sets (45 and 47), in
solvent (C124) and air dry.
(b) land pack bearing sets (45 and 47) with
grease (C70).
(c) Protect bearing sets (45 and 47) from
contamination until installation.
(2) Position bearing sleeve (38, figure 5-43) on
collective sleeve (36), and install retaining rings (37 and
39) in grooves near lower end of sleeve.
(3) Position spacer (41) and retaining rings (40
and 42), loosely on collective sleeve. Use adapter (T28)
to press seal (43) into nut (44), with seal lip toward
notched side of nut. Place nut, notched side down,
loosely on sleeve below shoulder.
Take suitable
precautions to avoid marring loose parts in handling.
(7) Place sleeve assembly on a press with support
halves (T27) under bottom bearing of stack. Place spacer
ring (49) on top bearing stack. Press hub assembly from
press.
(8) Install wrench (T29) on top of hub with two
bolts. Secure wrench in vise. Start lower nut (44) into hub
and torque 400 TO 500 foot-pounds. Use wrench (T4).
(9) Position lockplate (7)
(10) Install ring (41) ,on collective sleeve with
retaining rings (40 and 42) in mounting grooves of
collective sleeve below hub.
(11) (AVUM) Assemble inner race (20) washer (19)
and housing (18) to scissors.
(12) Attach link (32) to scissors (16) with inner race
{20), washer (19), and housing (18) in position, but (do not
install shim (17) at this time. Install nut (35) on bolt (30)
finger tight.
NOTE
The mating surface of the bearings in the
bearing set should be wiped free of excess
grease before they are installed.
5-78B
Change 65
TM 55-1520-234-23
(13) Measure gap between housing (18) and
bushing lace of drive link (32) with feeler gauge. Record
this figure.
Prepare a shim (17) by peeling off
laminations to obtain a shim thickness 0.000 TO 0.002
inch less than measured gap. Remove bolt (30) and
reinstall with inner race (20) shim (17), washers (19, 31
and 33) and housing (18) in position. Torque nut (35)
85 TO 104 foot-pounds and install cotter pin (34).
Repeat for opposite scissors and link.
NOTE
End play between scissors and drive link is
necessary after establishing torque. Maximum
end play shall not exceed 0.116 inch.
(14) (AVIM) Install scissors (16) on hub with
inner races (3 and 14), spacer (4), washer (25), cap
washer (26), and special washers (10 and 27) in
position. One special washer (10) must be under bolt
head and one special washer (27) must be under nut.
The bolt head must face in direction of rotation.
TORQUE NUT (29) TO 150 - 175 FOOT-POUNDS and
install cotter pin.
Repeat for opposite scissors.
Change 65 5-78C/(5-78D blank)
TM 55-1520-234-23
(15) Lubricate all bearings in scissors and hub
as specified in Chapter 1.
nut finger tight
(16) Collet finger teflon can be rebonded with
versilock 204, (C156) and accelerator No. 5, (C157),
using work aid (figure 5-47C). Teflon pads from one
collet set may bc used on other collet sets.
POUNDS.
(17) Collets may be used in mixed sets
providing minimum gap of 0.040 inch each side is
maintained.
(g) TORQUE NUT (15) TO 1250 - 1550 INCH
POUNDS and install cotter pin.
(e) TORQUE NUT (5) TO 50 - 70 INCH-
(18) No missing fingers, badly worn teflon pads,
cracks or scoring allowed.
g. Installation. See Figure 5-40.
(f) TORQUE NUT (3) TO 160 - 190 INCHPOUNDS.
(h) Check for a maximum of 0.060 inch
clearance between thrust washers and bearing housings. A
minimum zero clearance is acceptable as long as no
binding is evident. The clearance is to be measured
collectively from both levers.
(1) Install swashplate and lower boot if not
previously accomplished. Refer to paragraph 5-8.
CAUTION
Special washers (63 and 70, figure 5-40) are not
interchangeable and must be installed in correct
location to perform fail-safe function.
(2) Coat mating splines on mast and in scissors
and sleeve assembly spline plate (34 or 52) with grease
(C70).
(3) Carefully lower scissors and sleeve
assembly over mast. Insert lower end of collective
sleeve down through lower boot (60) and top of
swashplate support. Use care to avoid damage to
teflon-lined bearing inside support.
(4) Turn collective sleeve (18) so that the two
bearing mounting bosses at lower end are aligned with
openings in swashplate support as illustrated. Position
spacer plate (20) and bearing assembly on boss with the
"TOP" marking up so that curve inner surface of bearing
housing is aligned to mast surface. Install screws (13)
and lockwire (C151) in pairs. Install opposite bearing
assembly in same manner.
(6) Place a special washer (69, figure 5-40), with
chamfer facing outboard, on swashplate outer ring as
illustrated. Position drive link (58) on swashplate then
install special washer (70) with collar inboard and the letters
"AFT" facing out-board. Install nut (71) and TORQUE TO
770-950 INCH-POUNDS and install cotter pin. Bend cotter
pins ends closely around nut to avoid contact with
swashplate during operation. Install opposite drive link in
the same manner.
Measure vertical clearance from the bottom of both
drive links, P/N 209-010-408-7, to all three horns of
stationary swashplate. The minimum clearance
must not be less than .035 (thirty five thousandths)
inch. Replace swashplate if clearance is below
minimum.
(5) Install collective lever halves (7 and 10) as
follows:
(a) Place a thrust washer (14) over the
bearing boss of each lever half.
(b) Mount lever halves on bearing
assemblies (19) and install bolt (11), washer (6) and nut
(5). Install nut finger tight.
(c) Position lever halves on link (21) with
inner race (23) and thrust washers (17 and 24) in place.
Install bolt (1) washers (2 and 16) and nut (15). Install
nut finger tight.
(d) Position spacer (9) between levers and
install bolt (12). Install washer (4) and nut (3). Install
(7) Slip lower boot (60, figure 5-40) on grooved
ring on collective sleeve hub and on grooved ring on
collective sleeve below hub. Secure both ends of boot with
lockwire.
(8) Determine whether the extension to be installed
above the scissors and sleeve assembly is P/N 540-011487-1 as shown in detail view A, figure 5-40 or is P/N 209010-464-1 as shown on detail view B.
Change 65
5-79
TM 55-1520-234-23
Install parts as described in step (a) or step (b) as
applicable.
(a) Install extension (33, figure 5-40) and
associated parts shown on detail view A as follows:
1 Coat mating splines on mast and
spline plate (34) with grease (C70). Position spline plate
and extension (33) on mast and install bolts (36) and
washers (35). Lockwire bolt heads in sets of three.
Provide adequate ventilation when using
methyl-ethyl-ketone. Avoid breathing vapors
and avoid prolonged skin contact.
2 Clean friction sleeve on mast,
clamp assembly (29), nut (31) collet set (32) and
extension (33) with methyl-ethyl-ketone (C87).
3 Seat collet set (32) in top of
extension (33) and install nut (31). TORQUE NUT TO
140 - 180 FOOT-POUNDS. Check that no gap exists
between collet set and friction sleeve on mast. Install
spring pin (37) through aligned holes in nut and
extension and secure with lockwire looped through
spring pin and over edge of nut and extension.
4 Position rubber ring (30) around
collet set (32) and top of nut (31). Place clamp
assembly (29) around rubber ring and install bolts (38),
washers (28 and 39) and nuts (27). Tighten nuts evenly
so that gaps between clamp sections are equal within
1/16 inch.
5
Adjust collective friction.
Refer to
step (9).
(b) Install extension (51, figure 5-40) and
associated parts shown on detail view B as follows:
1 Coat mating splines on mast and
spline plate (52) with grease (C70). Position spline and
extension (51) on mast and install bolts (56) and
washers (55). TORQUE BOLTS EVENLY TO 80 - 100
INCH-POUNDS and lockwire in sets of three.
Provide adequate ventilation when using
methyl-ethyl-ketone (C87) Avoid breathing
solvent vapors and avoid prolonged skin
contact.
5-80
2 Clean friction sleeve on mast,
clamp assembly (44), retainer (49) collet set (50) and
extension (51) with methyl-ethyl-ketone (C87).
3 Seat collet set (50) in top of
extension (51) and install retainer (49) with bolts (48)
washers (54) and nuts (53). TORQUE NUTS EVENLY
TO 80 - 100 INCH-POUNDS. Check that no gap exists
between collet set and friction sleeve on mast.
4 Position rubber ring (47) around
collet set (50) and on top of retainer (49). Place clamp
assembly (44) around rubber ring and install bolts (46),
washers (43 and 45) and nuts (42). Tighten nuts evenly
so that gaps between clamp sections are equal within
1/16 inch.
5
Adjust collective friction.
Refer to
step (9).
(9) Adjustment - collective mast friction collet.
See figure 5-40.
(a) Disconnect collective controls
collective lever halves (7 and 10) if connected.
from
(b) Attach a force gauge (fish scale) to
collective lever halves (7 and 10) at point where
collective controls are normally attached. Place lever
halves in full down position and measure amount of
force in pounds required to raise the lever halves.
Adjust bolts (38 or 46) as applicable until a load of 125
to 135 pounds on the spring scale is required to required
to raise the lever halves.
NOTE
Between five and ten hours of operation
following installation, recheck friction as outlined
in preceding step If friction setting is not within
limits, readjust. Do not exceed 130 inch-pounds
torque on clamp bolts. If correct friction cannot
be obtained within this limit, check for grease on
mast friction sleeve and clean with methyl-ethylketone (C87).
(c) Attach collective controls to lever
halves (7 and 10) with bolt, washers, nut, and cotterpin.
(10)
Install upper boot (26, figure 5-40) with
spacer (25) or upper boot (41) with spacer (40) as
applicable. Secure upper boot with lockwire. Apply a
small bead of sealant (C116) around top of upper boot.
(11)
Install main rotor hub and blade
assembly and perform maintenance test flight. Refer to
paragraph 5-4.
Change 64
TM 55-1520-234-23
Figure 5-47C. Collet Work Aid
Change 56
5-80A/(5-80B blank)
TM 55-1520-234-23
(a) Disconnect anti-drive link (64,
figure 540), lateral control tube and fore and aft control
tube from swashplate inner ring assembly.
5-8. Swashplate and Support
a.
Description
The swashplate and support assembly is a component
of the main rotor controls (figure 548). The swashplate
support is an open cylinder with a flange for mounting to
the transmission at the lower end and spherical surface
or uniball at the upper end for mounting the swashplate.
Side openings are provided to accommodate the
collective lever halves which move the collective
sleeve. The swashplate inner ring is clamped on the
pivot ball of the support by upper and lower sets of
contoured, teflon-lined bearings. This design allows the
swashplate to tilt in any direction when actuated by the
cyclic control rods. The anti-drive link prevents the inner
ring from rotating. The swashplate outer ring tilts with
the inner ring, but rotates with the scissors and mast. It
is mounted to the inner ring through a duplex ball thrust
bearing, and is connected to the scissors with two drive
links.
Premaintenance Requirements for Swashplate
and Support
Condition
Support Equipment
Minimum Personnel
Required
Consumable Materials
Special Environmental
Condition
b.
(c) Apply a force gage (fish scale)
to bolts inserted through both the lateral and fore and aft
devises on control horn inner ring.
(d) Check for 15.5 TO 20 pounds of
force required to actuate swashplate about the uniball at
each clevis. If friction is within limits, reconnect the
items that were disconnected in steps (a) and (b). If
friction is not within limits, adjust and recheck thickness
of shim (3, figure 5-48).
Ensure that wood wedges remain in position to
support inner ring during shim adjustment
procedure or uniball damage may result.
1
Insert two wood wedges under
swashplate inner ring to support the ring during shim
adjustment procedure.
Requirements
Model
Part No. or Serial No.
Special Tools
Test Equipment
(b) Disconnect scissors and sleeve
drive links (58), from swashplate outer ring assembly.
AH-1S
All
NA
Force gauge (fish scale)
capable of measuring up
to 30 pounds
Two
(C9) (C32) (C37) (C44)
(C45) (C52) (C76) (C87)
(C102) (C116) (C124)
None
NOTE
A one piece stainless steel shim (3) may be
used in place of the four piece shim in the steps
below.
2
Remove shield (1), upper
bearing (2) and shims (3). Measure thickness of each
section of shim with a micrometer. All four shim
sections must be the same thickness.
3
Remove or add one shim
laminate to each of the four shim sections (pieces).
Adjustment
(1)
Adjust friction on swashplate installed
on helicopter as follows:
NOTE
Do not apply more than 70 inch-pounds torque
to nuts (14) that secure upper bearing and shield
(1) for any reason.
Prior to disconnecting anti-drive link, check for
excessive wear in drive link bearings. Rotation of the
swashplate inner ring, measured at the swashplate aft
horn pin, in excess of 0.110 inch indicates worn antidrive link bearing and bushings.
4
Recheck shim (3) to be sure all
sections (pieces) are the same thickness and install the
shim. Ensure that the inner diameter of shim (3)
Change 48
5-81
TM 55-1520-234-23
Figure 5-48. Swashplate and support assembly
PAGE 5-82A/5-82B DELETED
5-82 Change 29
TM 55-1520-234-23
does not extend over the edge of inner ring (4). Fill
gaps between ends of sections of shim (3) with corrosion preventive compound (C52).
Install upper
bearing (2) and shield (1). Install aluminum washers
(13) and nuts (14). Torque nuts evenly 50 TO 70 inchpounds while rocking ring assembly to ensure seating of
bearings.
5
Repeat friction check and if
friction is not within limits, disassemble and make
additional adjustment of thickness of shims (3).
6
Deleted.
7
Remove wood
were placed under inner ring in step (a).
wedges
that
8
Lubricate thrust bearing through
fittings on outer ring (10) with grease (C70).
swashplate.
(5)
Disconnect cyclic control cylinder and
spring from right control horn.
CAUTION
Do not rotate inner ring unnecessarily while
swashplate linkage is disconnected.
(6)
Remove bolts (90) and washers (91).
Lift swashplate and support off mast. Use caution to
avoid damage to mast friction sleeve and mast splines.
(7)
Remove cotter pin (68), nut (67), flat
washer (92), bolt (65), special washer (66), and remove
anti-drive link (64). Remove bolt (78) and remove
bellcrank (75). Remove support (79).
e.
(e)
Install anti-drive link (64, figure
5-40) and drive links (58).
(2)
Install hydraulic control cylinders for
lateral and fore and aft controls.
c.
Lubrication.
Lubricate swashplate and support
lubrication chart in Chapter 1.
d.
as
shown
on
Inspection.
NOTE
If allowable inspection limits are exceeded,
forward swashplate and support to depot
maintenance.
(1)
Inspect swashplate inner ring horns
for wear caused by improperly installed cotter pins in
drive link to swashplate attachment bolts. Maximum
permissible wear is 0.060 inch.
Removal. See figure 5-40.
(2)
Rotate outer ring and check for
binding and roughness of bearings. No binding or
roughness is acceptable.
CAUTION
Remove swashplate and support carefully to
avoid damage to mast.
NOTE
(1)
Remove main rotor hub and blade
assembly. Refer to paragraph 5-4.
Do not disassemble swashplate and support for
inspection.
(2)
Remove
scissors
assembly. Refer to paragraph 5-7.
and
sleeve
(3)
Remove nut (62), special washer (63)
and disconnect anti-drive link (64) from rear horn of
swashplate.
(3)
Check visible portions of assembly for
nicks, dents and corrosion in accordance with limits
shown in figure 5-49 and 5-50 for limits (4).
(4)
Check security of pin in aft horn of
swashplate inner ring. No noticeable looseness is
acceptable.
(4)
Disconnect cyclic control cylinder tube
and elevator control tube from forward horn of
Change 71
5-83
TM 55-1520-234-23
Figure 5-49. Damage limits - swashplate and support assembly (Sheet 1 of 3 )
5-84 Change 29
TM 55-1520-234-23
Figure 5-49. Damage limits - swashplate and support assembly (Sheet 2 of 3 )
5-85
Change 29
TM 55-1520-234-23
Figure 5-49. Damage limits - swashplate and support assembly (Sheet 3 of 3 )
5-86
Change 29
TM 55-1520-234-23
Figure 5-50. Damage limits - swashplate anti-drive link, bellcrank, and support (Sheet 1 of 3 )
5-87
Change 56
TM 55-1520-234-23
Figure 5-50. Damage limits - swashplate anti-drive link, bellcrank, and support (Sheet 2 of 3 )
5-88
Change 7
TM 55-1520-234-23
Figure 5-50. Damage limits - swashplate anti-drive link, bellcrank, and support (Sheet 3 of 3 )
5-89
Change 29
(2)
Replace
identification plates (AVIM).
NOTE
If friction check in step 5 below is made with
swashplate
and
support
installed
on
transmission, disconnect drive links, antidrive
links, control tubes and spring on inner ring.
(5)
Check
uniball. (Refer to step b.)
friction
of
swashplate
TM 55-1520-234-23
damaged
or
missing
NOTE
to
(6)
Inspect bushings in inner ring at
attachment points, and control tubes for looseness, wear
and mechanical damage. Maximum allowable wear on
bushing inner faces contacted by control tube bearings
is 0.060 inch.
If data to be stamped on identification plate and
is not available, send affected assembly to
Depot Maintenance.
(a) Stamp all data from
identification plate on the new identification plate.
old
(b) Remove screws and remove old
identification plate from support assembly.
(7)
Visually inspect swashplate support,
inner ring and outer ring for damaged grease fitting,
missing or damaged identification plates or other
damage in excess of limits shown in figure 5-49.
Cleaning solvent is flammable and toxic.
Provide adequate ventilation. Avoid pro- longed
breathing of vapors and contact with skin or
eyes.
Cleaning solvent is flammable and toxic.
Provide adequate ventilation. Avoid prolonged
breathing of vapors and contact with skin or
eyes.
(8)
Inspect bracket (23, figure 5-48).
Replace if loose or missing. Clean mating surface of
inner ring (19) with MEK (C87) and wipe dry with a clean
cloth. Apply primer (C99) to surface and allow to air dry.
Remove protective peel ply from film adhesive on
bracket (23). Coat mating surface of bracket (23) with
adhesive (C17). Install bracket (23) on inner ring (19)
with radius of bracket parallel to edge of inner ring and
with holes align- ed. Use caution to prevent adhesive
squeeze-out from obstructing bolt hole. Maximum
allowable wear on bushing inner faces contacted by
control tube bearings is 0.060 inch.
f.
(3)
Clean area for identification plate on
support assembly with clean cloth saturated with solvent
(C124).
Do not overtorque screws.
(4)
Position new identification plate on
sup- port (20). Place drive screw through identification
plate and in hole in support. Drive screw in until
identification plate is tight against support.
(5)
Remove and replace retaining ring
(22 and bearing bushing (21) if damaged.
g.
Repair.
NOTE
NOTE
Replace swashplate and support if allowable
inspection
limits
are
exceeded.
Send
unserviceable swashplate and sup- port to next
higher level of maintenance.
(1)
Installation.
If swashplate and support assembly is new,
lubricate in accordance with Figure 1-2, item 7,
prior to installation.
Figure 5-51 deleted.
Pages 5-91 and 5-92 deleted.
Replace damaged or missing grease
fitting.
5-90
Change 48
TM 55-1520-234-23
(1)
Install swashplate and support as
follows:
(a) Install support assembly (79,
figure 5-40) on transmission.
(b) Install bellcrank (75) on support
(79) with bolt (78) and nut (80). Torque nut 190 TO 210
inch-pounds and install cotter pin.
(5)
Install scissors and sleeve assembly
(paragraph 5-7).
5-9. Pitch Link Assembly.
The pitch links (14, figure 5-1) are attached to the
main rotor hub pitch horns (17) and to the scissors and
sleeve assembly (16).
a. Removal.
(c) Install anti-drive link (64) on
bellcrank with bolt (65), washer (66), (92), and nut (67).
Torque nut 190 TO 210 inch-pounds and install cotter
pin.
(1)
Prior to accomplishment of MWO 551520-244-50-9, remove lockwire from bolt (20). Remove
bolt (20) and washer (19). Disconnect pitch links (14)
from rotor hub pitch horns.
NOTE
The raised letters "AFT" identify the rear side of the antidrive link, and must be positioned toward rear of
helicopter.
(d) Lower swashplate and support
assembly over mast on to top of transmission. Avoid
damage to mast splines.
(e) Align holes in swashplate
support with holes in transmission cap. Install bolts (90)
with washers (91) and torque 200 TO 250 inch-pounds.
Lockwire bolt heads in pairs.
(1A) After accomplishment of MWO 551520- 244-50-9, remove lockwire from bolt (20).
Remove bolt (20) and bushing assembly (19A, view A,
figure 5-1). Disconnect pitch link (14) from pitch horn
(17).
(2)
Remove cotter pin (27), nut (26),
washer (25) and bolts (24) at lower end of pitch links.
Remove pitch links (14) from scissors and sleeve
assembly (16).
(3)
Repeat steps a(1) or a(la) through
a(3) to remove opposite pitch link (14).
(f)
Turn swashplate inner ring to
align stud with anti-drive link (64). Position link on stud
and install special washer (63) with marked surface
facing aft. Install nut (62) and torque 480 TO 690 inchpounds. Install cotter pin.
(4)
When maintenance requires removal
of universal bearing (23) proceed as follows. Remove
cotter pin (32), nut (31), recessed washers (29 and 30)
and bolt (28). Remove universal bearing (23) from
scissor and sleeve assembly (16).
(2)
Connect lateral control tube to left
horn of swashplate inner ring.
b.
Inspection. Inspect pitch links (14, figure 51) for damage such as metal to metal contact on upper
bearing housing, normal bearing wear, surface wear and
straightness of tube assembly. All damage limits are
given in figures 5-52 and 5-52A.
(3)
Connect
fore-and-aft
hydraulic
cylinder control tube, elevator control tube, and spring to
right horn of swashplate.
(4)
Position lower boot (60, figure 5-40)
loosely on swashplate.
Change 71
(1)
When inspecting ends of pitch tube
remove the first three inches of paint and primer
5-93
TM 55-1520-234-23
material from contacting the metal set material. Mask
the center of the tube length so that only the top and
bottom three inches of the tube are exposed.
NOTE
The success and reliability of penetrant
inspection depends upon the thoroughness with
which inspector prepares the part from the
process all the way through to the final
interpretation of the indications. All inspections
should be with the fluorescent penetrant (Type I,
Method Cl in strict accordance with TM 43-0103.
remove the paint.
NOTE
The gold colored finish on the metal is a very
thin chemcoat metal primer. This material is
not, repeat not to be removed.
(3)
Clean the prepared surfaces with a
soft cloth.
(4)
Apply a fluorescent dye penetrant to
the prepared surfaces from either a spray can or with a
soft hair brush and in strict conformance to the
procedure specified in TM 43-0103, Chapter 6.
(5)
Allow penetrant
minimum of 30 minutes.
Prolonged or repeated inhalation of vapors or
powders may result in irritation of the mucous
membrane areas of the body.
Provide
adequate ventilation.
to
dwell
for
a
(6)
Clean off all excess penetrant in
accordance with TM 43-0103 standard procedures.
(Check for complete excess penetrant removal from
surface by using a blacklight.)
(7)
Apply applicable developer consistent
with Type I, Method C penetrant method in TM
43-0103.
Continual exposure to penetrant inspection
materials may cause skin irritation. Avoid
prolonged breathing of solvent vapors and
contact with skin or eyes.
(8)
Inspect suspected area with blacklight
source in subdued white light.
Injury to eyes and skin may occur when
blacklight is not used in accordance with
manufacturer's instructions.
Unfiltered light
sources (if filter is required) may possibly
damage the eyes.
(9) If any apparent cracks appear (or suspect
surface defects), the suspect area must be reevaluated
utilizing certified NDI personnel with the Eddy Current
method per TM 43-0103, Chapter 3. (Tube material is
2024 aluminum.)
Temperatures in excess of 120 degree F may
cause bursting of pressurized cans and injury to
personnel.
NOTE
Normal manufacturing machining marks may be
observed on the tube surfaces. These will not
be cause of part rejection.
NOTE
If a physical or penetrant crack is observed and
confirmed, report failures or QDR Form 368 in
accordance with TM38-750 and hold tube as an
exhibit.
(10) Clean tube with solvent and wipe dry.
Volatile fumes may occur, creating both a fire
and health hazard.
NOTE
Paint will not be removed by any mechanical
means under any circumstances because it may
mask over any potential surface cracks.
(2)
With a soft hair brush, apply MEK
(Metyl-Ethyl-Ketone (C87)) or paint remover and
5-94
(11) Inspecting for straightness is done by
doing a Total Inline Runout (TIR).
(12)
Deleted.
(13) Set up bearing blocks for pitch tube
so that the edges of tube are resting between roller
bearings.
Change 29
TM 55-1520-23423
(14) Set up indicator within a 1/4-inch from
the edge of the larger diameter of the pitch tube. Dial
indicator must not contact necked down area.
(15) Rotate pitch tube to find highest and
lowest reading and zero dial at that position.
(16) Rotate pitch tube and record the total
bend. Maximum allowed TIR is 0.020 inch; tubes in
excess of 0.020 inch shall be reported and held as an
exhibit.
(17)
Repaint areas to original color (C80)
of tube.
Change 56
5-94A
TM 55-1520-234-23
Figure 5-52. Damage limits-pitch link assembly
(prior to accomplishment of MWO 55-11520-244-50-9) (Sheet 1 of 2)
5-94B
Change 71
TM 55-1520-234-23
NOTES:
1.
All edges may be radiused or chamfered 0.030 inch to remove nicks or dents.
2.
Repair of nicks and dents on threads must not exceed one-third of the thread depth. Length of repair shall not
exceed 0.250 inch. Each threaded segment may have two repair areas.
3.
Coat repair areas on steel parts with brush cadmium or zinc chromate and aluminum parts with zinc chromate. Do
not use zinc chromate on threads.
4.
Corrosion must be cleaned up to twice the depth of damage on aluminum. Corrosion must be cleaned up to
remove all traces of damage on steel.
5.
Minimum radius of repair on adapter is 0.100 inch. The repair must be polished to match the surrounding surfaces.
6.
Maximum play in bearing Part No. 209-010-443-1 is 0.020 inch axial or radial.
7.
Maximum play in bearings in universal Part No. 214-010-434-1 is 0.010 inch axial or radial.
8.
Do not remove adapter or clevis from tube.
9.
Visual irregularities caused by swaging operation at each end considered acceptable provided there are no sharp
ridges or grooves.
10.
Visually inspect tubes for any indications of bending. No bending allowed.
11.
Do not change color of tube when repainting or touching-up.
12.
Every 150 hours at phase inspection, inspect control tubes of TIR. No more than 0.020 inch is allowed at one inch
from each end of control tube.
13.
Repaired areas may not overlap.
14.
No cracks allowed.
15.
Width of repairs to the 209-010-460-3 tube cannot exceed 1/3 of the tube circumference.
209010-105-2B
Figure 5-52. Damage limits-pitch link assembly
(Prior to accomplishment of MWO 55-1520-244-50-9) (Sheet 2 of 2)
Change 71
5-94C
TM 55-1520-234-23
NOTES:
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
All edges may be radiused or chamfered 0.030 inch to remove nicks or dents.
Repair of nicks and dents on threads must not exceed one-third of the thread depth. Length of repair shall not
exceed 0.250 inch. Each threaded segment may have two repair areas.
Coat repair areas on steel parts with brush cadmium or zinc chromate and aluminum parts with zinc chromate. Do
not use zinc chromate on threads.
Corrosion must be cleaned up to remove all traces of damage on steel.
Minimum radius of repair on adapter is 0.100 inch. The repair must be polished to match the surrounding surfaces.
Maximum play in bearing in universal Part No. 214-310-401-101 is 0.020 inch axial or radial.
Maximum play in bearings in universal Part No. 214-010-434-1 is 0.010 inch axial or radial.
Do not change color of tube when repainting or touching-up.
Width of repairs to 209-010-518-101 tube cannot exceed 1/3 or tube circumference. Four repairs are allowed on
209-010-518-101 tube in this area.
Repairs may not overlap.
Figure 5-52A. Damage limits-pitch link assembly
(After accomplishment of MWO 55-1520-224-50-9)
Change 71
5-94D
TM 55-1520-234-23
c.
d.
Repair.
Adjustment.
Adjust both pitch links to length between centers
of bearings as shown on figure 5-6.
(1)
Polish out corrosion and mechanical
damage that is within limits and touch up repair areas in
accordance with figures 5-52 and 5-52A.
e.
(2)
Replace bearings if worn beyond
limits shown on figures 5-52 and 5-52A.
Installation.
Refer to paragraph 54.
Section II. TAIL ROTOR SYSTEM
5-10. Tail Rotor System.
Special Tools
The tail rotor system consists of the tail rotor hub
and blade assembly and the tail rotor controls installed
on the tail rotor gear box with the tail rotor.
Condition
Requirements
Test Equipment
Support Equipment
Minimum Personnel
Required
Consumable Materials
Special Environmental
Condition
Tracking stick
NA
Two men
5-11. Troubleshooting and Tracking CheckTail Rotor System.
Premaintenance Requirements for Troubleshooting and
Tracking Check of Tail Rotor System
Condition
Requirements
Model
Part No. or Serial No.
AH-1S
All
a.
NA
(C103)
None
Troubleshooting.
Troubleshoot the tail rotor system in accordance with
table 5-2. Refer to step b for tracking instructions.
Table 5-2. Troubleshooting-Tail Rotor System
NOTE
Before you use this table, be sure you have performed all normal operational checks.
CONDITION
TEST OR INSPECTION
CORRECTIVE ACTION
1.
High Frequency Vibration.
STEP 1. Check tail rotor track.
Adjust pitch link to bring tail rotor in track. (Refer to paragraph 5-11.b.)
STEP 2. Tail rotor out of balance.
Remove tail rotor and balance. (Refer to paragraph 5-12a. and b.)
Change 71
5-95
TM 55-1520-234-23
Table 5-2. Troubleshooting - Tail Rotor System (Cont)
CONDITION
TEST OR INSPECTION
CORRECTIVE ACTION
STEP 3. Check for worn or loose trunnion bearings. (Refer to paragraph 5-14c.)
Replace trunnion bearings. (Refer to paragraph 5-14d.)
STEP 4. Check for loose or worn blade retention bearings. (Refer to paragraph 5-14c.)
Replace blade retention bearings. (Refer to paragraph 5-14d.)
STEP 5. Cracked or loose tail rotor hub retaining nut.
Inspect and retorque nut. (Refer to paragraph 5-12c.)
STEP 6. Check for unserviceable pitch change links.
(Refer to paragraph 5-15d.)
STEP 7. Check for worn or loose pitch change link bearings. (Refer to paragraph 5-15d.)
Replace pitch change link. (Refer to paragraph 5-15f.)
STEP 8. Check for worn or loose pitch change crosshead bearing.
Replace pitch change bearing. (Refer to paragraph 5-15d.)
STEP 9. Check for loose or improperly torqued bipod and tripod engine mounts. Refer to paragraph 2-44a.)
Torque bipod and tripod engine mount bolts. (Refer to paragraph 2-44d.)
STEP 10. Check for loose mounting bolts on intermediate and tail rotor gearboxes.
Torque mounting bolts. (Refer to paragraphs 6-25e and 6-27f.)
STEP II. Check for elongated mounting bolt holes for intermediate and tail rotor gearboxes. (Refer to paragraph
6-25c and 6-27c for limits for tail rotor gearbox mounting bolt holes in support fitting.) (Refer to paragraph 6-25c
and 6-27c for limits for intermediate gearbox mounting bolt holes.)
2.
Inability to make normal right or left turn in flight.
STEP 1. Not applicable.
Check tail rotor rigging. Refer to paragraph 11-13a for rigging check. Return to paragraph 5-11b for
tracking instructions.
5-96
Change 2
TM 55-1520-234-23
b.
Tracking Check.
Following replacement or installation of the tail
rotor hub, blades or pitch change systems, check the tail
rotor system rigging and track the tail rotor blades.
(1)
Make a tracking stick by attaching a
small piece of sponge rubber 1/8 to 1/4 inch thick to end
of a 1/2 x 1/2 inch pine stick or any other flexible device.
Coat sponge rubber with Prussion blue (C103) or similar
type of coloring thinned with oil.
(a) Start engine and run at 6600
rpm, with pedals in neutral position.
(b) Turn force trim-OFF.
Place
HYD control switch to SYS 2-ON, so tail rotor hydraulic
cylinder is not powered. Observe for pedals remaining
at neutral when foot pressure is removed.
(c) If left pedal creeps forward less
weight is needed on tail rotor counterweight.
NOTE
The runup shall be performed by personnel
authorized in accordance with AR95-1.
(e) Move HYDR switch to ON
position and shut down engine.
WARNING
Do not approach the tail rotor area for the
purpose of tracking until 6600 rpm has been
established and it is certain that the aircraft is
not going to yaw left or right due to rigging error
or slippery parking surface. Severe injury or
death could result from being struck by the tail
rotor blades.
(2)
Start engine.
Run at 100% with
pedals in neutral position. Rest tracking stick on
underside of tail fin assembly as shown on figure 5-53.
Slowly move tracking stick toward disc of tail rotor just
far enough to lightly contact a blade approximately one
inch from tip. Pull stick back immediately.
(3)
After contact is made, stop engine
and allow rotor to stop. Shorten pitch link of unmarked
blade 1/2 turn of rod end. Reinstall pitch link bolt,
washers and cotter pin.
(4)
Recheck track of blades. Proceed
with adjustments if required by adjusting pitch links
equally in opposite directions.
(5)
Make operational test flight and check
that normal right and left turn can be made in
autorotational and powered flight.
(6)
(d) If right pedal creeps forward
more weight is needed on tail rotor counterweight.
CAUTION
Identical weight must be maintained at all four
positions on bellcrank (P/N 212-010-709-1).
Use a maximum of two weights (P/N 212-010710-1) and one washer (P/NAN960-416) or a
minimum of one weight and five washers.
(7)
figure 5-67):
(a) Remove nut (P/N MS21042L4)
and washer (P/N AN960-416) from bolt (P/N AN4-10A).
(b) Add or remove weight as
determined in steps (6)(c) or (6)(d). Remove only one
weight at each position. Replace with AN960 or AN970
washers, or a combination of both types.
(c) Install washer and nut on bolt.
Torque nut 50 to 70 inch-pounds.
(d) Repeat steps (6) and (7) until
forces are satisfactory.
(e) Track tail rotor blades. Tail
rotors with pitch link measurements as stated in
paragraph 5-15.f.(4) do not require tracking.
Check tail rotor forces as follows:
Change 71
Adjust tail rotor forces as follows (see
5-96A/(5-96B blank)
TM 55-1520-234-23
Figure 5-53. Tracking tall rotor
5-12. Tall Rotor Hub and Blade Assembly.
The tail rotor hub and blade assembly counteracts
torque of the main rotor and provides directional control.
It consists of the hub and two blades. The hub
assembly has a preconed, flex- beamed-type yoke and
a two-piece, delta-hinged trunnion to rotor flapping. The
trunnion is splined to the tail rotor gear box shaft and
drives the rotor. The yoke has two self-lubricated;
spherical bearings for attaching points for each- rotor
blade. Rotor pitch change is accomplished at these
bearings. The blades are all metal bonded assemblies
with a stainless steel spar and honeycomb core. A
system of counterweights is attached to the pitch control
system to balance control forces and assist in controlling
blade pitch. Premaintenance Requirements for Tail
Rotor Hub and Blade Assembly
Condition
Requirements
Model
Part No. or Serial No.
Special Tools
Test Equipment
Support Equipment
Minimum Personnel
Required
Consumable Materials
(C151)
Special Environmental
Condition
AH-IS
212-010-701-11
(T1) (T79) (T82) (T83)
NA
NA
1
a.
(C51)
(C70)
(C124)
None
Removal. See figure 5-54.
(1)
If the rotor and controls are to be
reinstalled, check color code dots and, if missing,
Change 29
5-97
TM 55-1520-234-23
Figure 5-54. Tail rotor installation
Change 54
5-98
TM 55-1520-234-23
re apply color code so that parts can be reinstalled in the
same relative position.
(2)
Remove bolts and separate
counterweight links (12) from counterweight
support (17).
both
(3)
Remove lockwire, screws (1), washer
(2), lock (3) and retainer (4) from crosshead (10).
NOTE
The special tools required to balance the tail
rotor hub and blade assembly are contained in
the special tool kits listed in the
premaintenance requirements list preceding
step a.
(4)
Push right pedal forward against stop
and remove cotter pin (6), nut (5) and washer (7) from
end of pitch change control tube (21).
The specific tools from these kits that are
required for the balancing procedure are
illustrated in figures 5-37 and 5-55. The tools
are identified in the legend by part number
and/or by the number of the kit which contains
that particular tool.
(5)
Grip both tail rotor blades firmly with
hands and twist blades to disengage bearing (8) from
control tube (22). Remove nylatron washer (9) from
crosshead.
NOTE
The area used for balancing must be a room
which can be closed off to provide a draft free
environment.
NOTE
If outer race of bearing (8) separates from inner
race, reprove inner race as outlined in step (6)
and scrap bearing.
(1)
Assemble parts of rotor balancing kit
(T79) that are shown on the right side of figure 5-37
except do not install balancing arbor (23). Use arbor (6,
figure 5-55) to balance this tail rotor.
(6)
Push left pedal forward against stop
and place 11/16 wrench between inner race of bearing
(8) and crosshead (10). Push right pedal to disengage
inner race from control tube (22).
(2)
Install fixture (2, figure 5-55) on lower
end 'of arbor (6) and tighten the two lower set screws
(10). There are a total of four set screws in fixture (2).
Do not tighten the two upper set screws (10).
(7)
Disconnect pitch links (11) from each
tail rotor blade pitch horn (19) by removing lockwire and
bolt (26). If same tail rotor is to be reinstalled, secure
bushing (27) and barrel nut and retainer (18) in place
with bolt (26). If tail rotor is being replaced, remove
bushing (27) and barrel nut and retainer (18).
NOTE
Prior to installing post assemblies (15), adjust
movable index pin (8) of the positioning post to
a dimension of 1.765 inch. (Length "L" in view
"B".) Tighten the locking set screw in index pin
(8), using 3/32 inch hex wrench to maintain
proper setting. (1.765 inch will achieve a zero
degree pitch in the tail rotor blades.)
(8)
Remove crosshead assembly (10)
from gearbox output shaft (23).
(9)
Cut lockwire and remove shield (13)
and retaining nut (14) as an assembly using spanner
wrench (T1). Remove counterweight support (17).
(10) Move hub and blade assembly (20)
outboard on splines and remove split cone set (24) as it
is released. Remove hub and blade assembly from
gearbox and place on a rack to prevent damage to
blades.
NOTE
Sleeve (21) normally remains on control tube
(22) unless the sleeve or control is to be
replaced. If sleeve is to be replaced, pull sleeve
outboard to engage threads and turn until
disengaged.
b.
Balancing.
Change 38
(3)
Install two post assemblies (15, figure
5-55) in the outboard holes in fixture (2). These holes
are designated "A" on detail view A. Thread the posts
into the fitting to full thread depth and tighten finger
tight.
(4)
Place the arbor and fixture on a work
bench with the arbor vertical. Install pilot bushing (4,
figure 5-55) on arbor with larger diameter end down as
illustrated.
(5)
Place the tail rotor hub and blade
assembly on the arbor with the data plate side of the
rotor yoke (3, figure 5-55) facing up.
(6)
Install a floating bushing (9) in each
pitch horn (16) if not previously accomplished. Rotate
the rotor on the arbor until indexing pins (8) are fully
seated with the set screw side of index pin in the floating
bushings and the floating bushing are fully seated in the
pitch horn.
5-99
TM 55-1520-234-23
Figure 5-55. Tool Application - Tail Rotor Hub and Plate Assembly Balancing
Change 38
5-100
TM 56-1520-234-23
(7)
Install positioning yoke (5, figure 5-55)
on arbor in same relative position to rotor yoke as
illustrated in "top view". Locate the 6-3/8 inch mark on
the scale marked on arbor (6). Align the upper surface
of the positioning yoke, which is identified on figure 5-55
the "Sensitivity Setting Reference Surface", with the 63/8 inch mark on the scale. Tighten two set screws (7)
to secure the positioning yoke to the arbor.
(8)
Move the tail rotor assembly and
balancing tools to the stand that was assembled in step
(1). Attach arbor (6, figure 5-55) to the stand with cable
P/N 2264, (13, figure 5-37) and quick disconnect
coupling P/N 2266 (10). Operate hydraulic pump (15) to
take up slack in cable.
(9)
Loosen two lower set screws (10,
figure 5-55). The two upper set screws should already
be loose as directed in step (2). This will allow fixture
(2) to slide down and contact workstand (1). Open the
hydraulic pump valve to lower the arbor and rotor
assembly until these parts are resting on fixture (2).
Ensure that all the following conditions are met, and
then tighten two lower set screws (10).
(a) Fixture (2) must be seated
firmly on workstand (1).
(b)
firmly on fixture (2).
(b) Use a bolt (16) of proper length
to accommodate washers. Minimum length bolt is P/N
AN4H-4A. Maximum length bolt is P/N AN4H-10A.
(c) At least one washer (17) or (18)
must be used if a bolt (16) is installed.
(d) After chordwise balance is
attained, lower the assembly until it rests on stand.
Torque bolts (16) 50 TO 70 inch-pounds. Do not
lockwire at this time.
(12) Correct spanwise imbalance by
adjusting special washers (2, 3, 11, and 12, figure 5-56)
and steel washers (4 and 11). Operate hydraulic pump
on stand to raise the rotor assembly approximately 1/4
inch above stand table. Comply with the following
limitations:
(a) When adding washers to
balance the rotor assembly spanwise, leave special
washers (2, 26, 19, and 13) P/N 140-007-33-32C4 next
to blade. Assemble the washers listed in step (b) with
the heaviest washers next to washers (2, 26, 19 and 13).
Pilot bushing (4) must be seated
(c) The rotor yoke (3) must be
seated firmly on pilot bushing (4).
(d)
The pitch horns (16), floating
bushings (9) and indexing pins (8) must be fully
engaged.
(e) The positioning yoke (5) must
be oriented with the rotor yoke as shown on the "top
view". The legs of positioning yoke (5) must contact a
flat surface of the rotor yoke.
(10) Operate hydraulic pump (15, figure 537) to raise the assembly approximately 1/4 inch above
stand table, close doors and windows, stop fans, etc., to
make the area draft free. Allow the rotor to stabilize
and observe the balance indication on the black disc.
See figure 5-39. Record the indication and correct
imbalance as outlined in steps (11) and (12).
(11) Correct chordwise inbalance by
adding special washers (18, figure 5-56) and steel
washers (17). Comply with the following limitations.
(a)
washer (18) P/N AN970-4 and steel washers (17) P/N
AN960-416, but only a maximum of ten washers can be
used on one bolt.
(b) Use combinations of special
washers (3 and 12) P/N AN970-8, steel washers (4 and
11) P/N AN960-816 and thin steel washers P/N AN960816L (not illustrated) as required to balance the
assembly.
(c) Use bolts (20) and (25) of the
proper length to accommodate washers. Minimum
length bolt is P/N NAS 1308-34. Maximum length bolt is
P/N NAS 1308-36.
(d) After spanwise balance is
attained, recheck chordwise balance and then lower the
rotor assembly until it rests on the stand. Torque nuts (5
and 10) and corresponding nuts on opposite blade to
500 inch-pounds.
(e)
Lockwtre two bolts (S)(16) to
hole in pitch horn with lockwire (C 151 ).
(13) Remove
balancing tools as follows:
(a)
55) from stand
.
Use any combination of special
Change 65
5-101
tail rotor assembly
from
Disconnect arbor (6, figure 5-
TM 55-1520-234-23
(b)
Remove the tail rotor assembly
and arbor from the stand and place on a work bench.
(c) Loosen two set screws (7, figure
5-55) and remove positioning yoke (5) from arbor (6).
(d) Rotate tail rotor assembly to
disengage indexing pins (8, figure 5-55) from pitch horns
and remove tail rotor assembly from arbor (6). Secure
floating bushings (9) to pitch horns.
Use solvent in a well ventilated area. Avoid
prolonged breathing of vapors and do not use in
an area with open flame or high temperature.
(6)
Wash bearing (8) and cavity of
retainer (4) with solvent (C124) and allow to air dry.
(7)
Handpack bearing (8), fill cavity of
retainer (4) and lubricate spline surfaces of crosshead
(10) with grease (C70).
(e) Disassemble pilot bushing (4),
arbor (6), post assembles (15) and fixture (2).
c. Installation. See figure 5-54.
(1)
Position hub and blade assembly (20)
on gearbox output shaft (23) with data plate side of hub
outboard and trunnion flap stops inboard. Align master
tooth of trunnion with master spline of gearbox output
shaft and slide hub and blade assembly on shaft until
trunnion is just started on second set of splines.
(2)
Place split cone set (24), with bevel
outboard, in groove between splines and shoulder on
gearbox output shaft. Ensure that split cone end gaps
are equal and slide hub and blade assembly inboard to
seat trunnion on split cone set.
(3)
Install counterweight support (17) on
gearbox output shaft and seat against hub. Install
retaining nut (14) and shield (13) as an assembly. Hold
rotor at hub, rotate counterweight support (17) as far
clockwise as possible and hold in position. Torque
retaining nut (14) to 900 inch- pounds with spanner
wrench (T1). Make final check to ensure that split cone
set (24) is properly seated and end gaps equally
spaced.
Lockwire retaining nut (14) to
counterweight support with lockwire (CI51). Prior to
installation, inspect split cones for any nicks, scratches,
indentations or any type deformities in the cones. If
damaged, replace.
NOTE
Install split cones as matched set only.
Spacing between split cone sets may vary after
operation. This variation of spacing does not adversely
affect the assembly.
(4)
Install control tube (22) and sleeve
(21) if not previously accomplished. Refer to paragraph
5-14.
Ensure that cotterpin (6) is properly installed
during the following step. After installation of
retainer (4) it will not be possible to inspect
cotter pin.
(8)
Place nylatron washer (9) and bearing
(8) in outboard end of crosshead (10). Align master
splines and position crosshead assembly on gear box
output shaft. Install steel washer (7) and new nut (5) on
end of control tube (22). Ensure that nylatron washer
(9) is properly seated. Torque nut 70 TO 100 inchpounds and install cotter pin.
NOTE
Do not intermix counterweight links P/N 212010-711-1 and -3.
(9)
Connect counterweight link (12) to
counterweight support (17) with bolt (28) two washers
(29) and nut (16). Torque nut 60 TO 110 inch-pounds
and install cotter pin (15). Install opposite link in same
manner.
(10) Coat mating surface of both bushings
(27) and pitch horns (19) with compound (C51).
(11) Position barrel nut and retainer (18) in
hole in pitch horn (19). Position pitch link (11) in pitch
horn (19) and install bolt (26) and bushing (27) with
flange next to bolt head. Torque bolt to 136 inchpounds and lockwire bolt to pitch horn with (C151)
lockwire. Install opposite pitch link in same manner.
Check both bushings (27) to ensure that bushing flanges
do not seat against pitch horn (19).
Ensure that cotter pin (6) is correctly installed prior to
installation of retainer (4).
(5)
Assemble crosshead and controls if
not previous accomplished. Refer to paragraph 5-14.
and
5-102 Change 54
(12) Install retainer (4) on crosshead (10)
torque
retainer
to
300
inch-pounds.
TM 55-1520-234-23
assembly from helicopter (paragraph 5-12). Place the
tail rotor assembly on a padded bench or similar work
area to prevent damage.
Ensure that lock (3) is properly installed to
secure retainer (4) to crosshead (10). Failure to
comply can result in loss of directional control.
(13) Install lock (3) on crosshead with two
washers (2) and two screws (1). Lockwire (C151)
screws and deform lock (3) into notches of retainer (4) in
two places near screws (1).
(14) Lubricate grease fitting in end of
retainer (4) with two shots of grease (C70).
(15) Perform rigging check.
Refer to
Chapter 11.
(16) Perform tracking checks.
paragraph 5-11.
(3)
Remove nuts (7 and 8) and washers
(6 and 9). Re- move bolts (23 and 22) and washers (24
and 21). Remove pitch horn. Remove opposite pitch
horn in same manner.
c.
Description.
The tail rotor blade is of all-metal, bonded
construction. Upper and lower aluminum alloy skins are
bonded to an aluminum honeycomb core. Externally
attached balance weights and balance screws inside the
blade tip facilitate blade balancing.
Premaintenance requirements for tail rotor blade assembly.
Condition
Requirements
Model
AH-1S
Part No. or Serial No
212-010-750-13
Special Tools
None
Test Equipment
None
Support Equipment
Paint spray equipment
Minimum Personnel
One
Required
Consumable
(C3) (C7) (C 1) (C12) (C17) (C23)
Material
(C24) (C37) (C39) (C41)
(C53) (C76) (C77) (C87)
(C88) (C100) (Cl12) (C123)
(C124) (C128)
Special Environmental
NA
Condition
b.
(2)
Cut lockwire and remove bolt (16) and
washers (17 and 18) from blade.
Remove
corresponding parts from opposite blade.
Refer to
5-13. Tall Rotor Blade Assembly.
a.
NOTE
The tail rotor hub and blade assembly must be
rebalanced if any parts are replaced or repaired.
It is good practice to index special balance
washers and bolts at time of disassembly so that
these parts can be reassembled in the same
location. This will make rebalancing easier.
NOTE
If during a movement a tinkle or a sandy sound
emits from inside of blades, the presence of this
condition is not cause for rejection of the blades.
This sound is debris (particles of aluminum
honeycomb) that were not thoroughly removed
during blade manufacture.
NOTE
Any damage caused by external forces to
installed blades that requires blade replacement
will cause mandatory tail rotor yoke
replacement. Evidence or documentation of
any abnormal side load applied to the blade as
a result of ground mishap sufficient to cause
yoke to impact severely on flapping stop is
cause for removal.
(1)
Inspect tail rotor historical records and
tail rotor blades for evidence that the blades have been
subjected to an accident or incident outside realm of
normal usage.
If such evidence exists, perform
applicable special inspections for overspeed, sudden
stoppage, hard landing and overtorque outlined in
Chapter 1 and the following:
NOTE
If there is no evidence of accident or incident,
proceed to step (2).
Removal. See figure 556.
(1)
Remove tail rotor, hub and blade
Change 49
Inspection.
5-103
TM 55-1520-234-23
Figure 5-56. Tail rotor and blade assembly
5-104
TM 55-1520-234-23
(a)
Overspeed inspection:
1
Check
for
bond
separation anywhere on the blade. If any separation
exists, dispose of the blade locally.
Change 29
2
Check balance screws
(9, figure 5-57) and external balance weights (3) for
movement. If any of these parts have moved outboard
due to centrifugal force, dispose of blade locally.
5-104A/(5-104B blank)
TM 55-1520-234-23
3
Check the blade grip bolt hole
bushings (2) for evidence of looseness. If any of the
four bushings are loose, forward the blade to next higher
maintenance level for repair.
4
Inspect four external buffer
pads (figure 5-60) for looseness or damage. If any of
the four buffer pads are loose or damaged, forward the
blade to next higher maintenance level for repair.
5
If blade inspection is noted in
steps 1, 2, 3 and 4, and there is no other visible
damage, the blade is serviceable.
5
If one of the blades, of a pair,
has been damaged slightly by denting, scrap the blade;
but the other blade may be reused after inspection by
depot level maintenance for water tightness and
spanwise balance.
(2)
Accomplish
normal
inspection of tail rotor blades and pitch horns by
inspecting blades visually and with standard inspection
equipment. Classify any damage present as negligible
at AVIM level, repairable at depot, or nonrepairable.
NOTE
Blades with only negligible damage may be returned to
service without repair.
(b) Sudden stoppage inspection:
1
Inspect the blade visually for
evidence that the blade has come in contact with the
ground, tailboom or other foreign object.
If such
evidence is found replace both blades.
1
Non-sharp
dents
located
inboard of station 30.0 that are not in excess of 0.015
inch deep are negligible (figure 5-58).
2
Inspect the blade skin visually
for wrinkles and deformations.
If this damage is
present, replace damaged blade. Inspect opposite
blade for damage and replace as necessary.
2
Non-sharp
dents
located
outboard of station 30.0 that are not in excess of 0.030
inch are negligible (figure 5-58).
3
If blade passes inspections
noted in steps 1 and 2, and there is no other visible
damage due to sudden stoppage, the blade is
serviceable.
3
Voids between doublers (4,
figure 5-57) and skin (6) or spar (10) which do not
exceed 0.50 inch chordcwise by 2.0 inches spanwise
and are not within 0.50 inch of the edge of doubler are
negligible.
(c)
Hard landing inspection:
1
Inspect the blade visually for
evidence that the blade has come in contact with the
ground, tailboom or other foreign object.
If such
evidence is found replace both blades.
2
Check for bond separation
anywhere on the blade. If any separations exist,
dispose of blade locally.
3
Check root end weights for
evidence of being moved. If such evidence is found,
dispose of blade locally.
4
If one of the blades, of a pair,
has been damaged badly enough that the metal has
been torn or any bond lines have been separated,
dispose of blade locally.
Change 22
4
Voids between the skin (6) and
spar (10) which do not exceed 0.50 inch chordwise by
2.0 inchs spanwise and are not within 0.250 inch of the
edge of the skin and are not in the outboard area where
skin overlaps the spar are negligible. Voids between the
skin and spar in the outboard area where the skin
overlaps the spar which do not exceed 0.250 inch
chordwise and 2.0 inches spanwise are negligible.
5
Voids between the skin (6) and
honeycomb core (7) which do not exceed 0.50 inch
chordwise by 2.0 inches spanwise are negligible.
(b) Nick, dent and scratch
reparable at AVIM level maintenance:
5-105
type
damage
TM 55-1520-234-23
Figure 5-57. Tail rotor blade assembly
Figure 5-58. Tail rotor blade station diagram and scratch-type damage area locations
5-106
TM 55-1520-234-23
NOTE
Blades with the type damage defined in this
paragraph require that the damage be polished
out to the depth required to remove the damage
including any nicks or scratches which may be
present in dents.
Use 300 grit or finer
sandpaper (C112) and scotchbrite (C113). Do
not fair-in or fill sharp or non-sharp dents with
adhesive as this would interfere with subsequent
inspections for cracks. Touch-up paint in areas
where mechanical damage is polished out.
1
Nicks and scratches inboard of
Station 30.0 which run within 0 TO 15 degrees of the
span line and are not in excess of 0.005 inch depth
(figure 5-58).
2
Nicks and scratches inboard of
Station 30.0 which run within 0 TO 75 degrees of the
chordline and are not in excess of 0.003 inch in depth
(figure 5-58).
NOTE
Blades which have voids within the limits
defined in this paragraph require that the voids
be repaired to return the blades to servicable
condition. Refer to paragraph 5-13d. for repair
instructions.
A void is defined as an unbonded area that is
supposed to be bonded. Many sub definitions of
voids have been made such as lack of
adhesive, gas pocket, misfits, etc. However,
the general term "void" as used herein makes
no distinction between those definitions.
1
Determination of limits when
two or more separate voids are involved:
a
When separate voids are
closer together than one inch, consider them as one
void.
3
Sharp dents inboard of Station
30.0 which are not in excess of 0.010 inch in depth.
b
If the voids are in two
areas, such as one void between the core and the skin
that is located within one inch of a void between the skin
and the doubler, consider them as one void and use the
limits for the area that are most strict.
4
Non-sharp dents inboard of
Station 30.0 which are not in excess of 0.030 inch in
depth.
2
Edge voids between butt block
(13, figure 5-57) and spar (10) or inner grip plates (14)
within the following limits:
5
Nicks and scratches outboard of
Station 30.0 which are not in excess of 0.01 0 inch in
depth.
a
Butt block and spar: any
length and 0.250 inch in depth.
6
Sharp dents outboard of Station
30.0 which are not in excess of 0.015 inch in depth.
7
Non-sharp dents outboard of
Station 30.0 and also outside of patchable area which
are not in excess of 0.040 inch in depth.
b
Butt block and inner grip
plate: 1.50 inches in length and 0.250 inch in depth.
3
Edge voids which are a
maximum of 0.060 inch in depth or 2.0 inches in length
and are located between the following components.
8
Non-sharp dents outboard of
Station 30.0 and also within patchable area as shown in
figure 5-59 which are not in excess of 0.125 inch in
depth.
doublers (4).
9
Nicks and scratches in the
trailing edge up to 0.030 inch in depth chordwise are
repairable, but the damage must be polished out over a
distance of three inches on each side of the defect.
(10).
(c)
Voids reparable at AVIM level
maintenance:
Change 22
a
Grip
plates
(11)
and
b
Doublers (4) and skin (6).
c
Doublers (4) and spar
d
Inner grip plates (14) and
e
Inner grip plates (14) and
(10).
skin (6).
spar
5-107
TM 55-1520-234-23
f
Butt block (13) and skin (6).
g
Skin (6) and spar (10).
4
Edge voids which are a
maximum of 0. 20 inch in depth (chordwise) and 3.0
inches in length (spanwise) and are located between
skin (6) and trailing edge strip (5).
5
Edge voids which are a
maximum of 0.50 inch in width (chordwise) between the
spar (10) and tip block (8).
(d) Damage to
reparable at AVIM level maintenance.
skins
that
is
1
Damage caused by a foreign
object that results in a crack or hole in the skin and is
located in the authorized area for repair by patching as
shown in figure 5-59.
2
Nick and scratch type damage
that exceeds the limits defined in step b.(2), and is
located in the authorized area for repair by patching as
shown in figure 5-59.
3
The maximum size of hole that
can be repaired is restricted by the requirement that all
of the defect must be cut out and the maximum size of
cut is 1-1/2 inch diameter.
(e) Cracks in adhesive at bond line
that are reparable at AVIM level maintenance:
1
Cracks in adhesive at bond line
between the phenolic blocks and skin, spar inside the
drain hole, inner grip plates or joint between phenolic
blocks are reparable by sealing. Inspect adhesive at
bond lines in blade butt area for cracks.
Place
inspection emphasis on the areas shown in figure 5-60.
Do not saturate the bond lines with MEK as it
will soften the adhesive.
2
If cracks in adhesive are
suspected, remove paint from area with clean cloths
dampened with MEK (C87) and reinspect.
(f)
Blade
damage
that
is
nonrepairable:
1
Fatigue cracks at any location
on the blade require local disposition of the cracked
blade.
2
A
blade
with
water
in
honeycomb core.
3
A blade with one or more cracks
developed in a previously repaired area.
4
A blade with nicks or cracks that
are located in dents and the total depth is in excess of
the limits specified in steps c(2)(b)3 through 8 if damage
is not within area where patches are allowed.
5
A blade with any void within
0.50 inch of the edges of the drain hole doubler (12,
figure 5- 57).
6
A blade with any void between
the drain hole doubler (12) and spar (10) within .50 inch
of the edge of the drain hole.
7
A blade with one or more holes
that do not fall within the area authorized for patches as
shown in figure 5-59 and/or a blade with a hole that
exceeds the 1.5 inch diameter restriction noted in step
c(2)(d)3.
8
A blade with any corrosion that
penetrates entirely through the skin.
Cleaning solvent is flammable and toxic.
Provide adequate ventilation. Avoid prolonged
breathing of vapors and contact with skin or
eyes.
9
A blade that is worn completely
through the spar at the tip.
10
A blade with edge voids deeper
than 0.1 inch at the tip end of any of the root end
doublers or grip plates.
5-108
Change 22
TM 55-1520-234-23
Figure 5-59. Tail Rotor Blade - Area Authorized for Patch-Type Repair
11
A blade with edge voids in the
leading edge or trailing edge of the doublers that are
0.25 inch or more in depth and show indications of
corrosion in the void.
12
A blade that failed to pass the
special inspections for overspeed, overtorque and
sudden stoppage.
(g) Inspect tail rotor pitch horns as
follows:
1
Inspect pitch horns for damage
in excess of the limits shown in figure 5-61.
2
damage to 1 threads.
Inspect
threaded
insert
Change 22
for
3
Slide a new pitch link bolt
through the bushing and matching hole used for
attaching the pitch link. If the bolt does not fit freely
through the holes, the pitch horn is not suitable for
further service.
4
Inspect for distortion of the pitch
horns with a straight edge placed against the machined
surfaces. Any distinct deviation from flat indicates that
the pitch horn is distorted and not suitable for further
service.
BY
fluorescent
5-108A/(5-108B blank)
5
Inspect pitch horns for cracks
penetrant method (TM 43-0103).
TM 55-1520-234-23
Figure 5-60. Tail rotor blade butt area repair
Change 22
5-109
TM 55-1520-234-23
Figure 5-61. Damage limits - tail rotor blade pitch horn
Change 22
5-110
TM-55-1520-234-23
d. Repair (A VIM).
(1) Replace any blade which has incurred
nonreparable damage (paragraph 5-13c.)
(2) Repair blades with voids that are within the limits
specified in paragraph 5-13c.
(d) Heat the cut out size of skin to 200 degrees F
(maximum) and remove the disc while it is heated.
Avoid damage to the honeycomb core.
(e) Deburr the edges of the hole and polish out any
scratches and nicks. Use 350 grit or finer sandpaper
(C112) and scotchbrite (C113).
(f) Cut a patch from aluminum alloy sheet 0.020 inch
thick. The patch must be large enough to overlap the
hole at least 0.750 inch all around the perimeter. Deburr
the edges of the patch.
Cleaning solvent is flammable and toxic. Provide
adequate ventilation. Avoid prolonged breathing of
vapors and contact with skin or eyes.
Do not allow MEK to enter rotor blade when
removing paint to inspect for cracks or when
cleaning prior to sealing edge voids. MEK will
soften the adhesive used in manufacture of the
rotor blades.
(a) Clean area around reparable edge voids with a
clean cloth moistened with MEK (C87) and dry with a
clean cloth.
(b) Prepare a small quantity of adhesive (C12) in
accordance with the manufacturers instructions. Apply
the adhesive to the edge void with the flat side of an
applicator such as wooden tongue depressor. Fill the
void with adhesive as deeply as possible.
(3) Repair blades with crack and hole type damage in
the skin by patching if the damage is within limits
specified in paragraph 5-13c.
(a) Ensure that the damage to be patched is in the
authorized area for patches (figure 5-59).
(b) Remove paint in area to be patched with 120
grit sandpaper (C112). After paint is removed, smooth
the area with 250 grit sandpaper (C12).
(c) Cut out the damaged skin with a hole saw or
use a sharp instrument to cut through the skin. Do not
exceed the 1.50 inch-diameter maximum cut out.
Change 65
(g) Clean the side of the patch that will be bonded by
sanding with 250 grit sandpaper (C112).
Cleaning solvent is flammable and toxic. Provide
adequate ventilation. Avoid prolonged breathing of
vapors and contact with skin or eyes.
(h) Wipe the mating surfaces of the patch and blade
with a clean cloth dampened with MEK (C87).
(i) Apply a thin coat of adhesive (C12 or C17) to the
mating surfaces of the patch and blade. Place the patch
on the blade, press down on the patch and move it back
and forth slightly to expel all air pockets in the adhesive.
Blend the excess adhesive around the edge of the
patch.
(j) Maintain pressure on the patch. Use weights,
clamps or rubber bands cut from inner tubes.
(k) Cure (C17) adhesive at 75°F for five days or at
180°F for sixty minutes. Cure adhesive (C12) at 80 to
90°F for 24 hours or at 140°to 155°F for thirty minutes.
(I) Touch up paint In area of patch (paragraph 5-13c
and TM 55-1500-345-23).
(4) Repair blades with nick, scratch, dent and notch
damage that is within the limits specified in paragraph
5-13.
(a) Polish out all nicks and scratches. Use aluminum
wool (C24) on aluminum parts. Use sandpaper (C112)
on stainless steel spar.
5-111
TM-55-1520-234-23
(b) Polish out damage in trailing edge over
distance of three inches on each side of the defect. Use
a steel hand file to remove most of the damage then
smooth out the area with aluminum wool (C24).
(c) Touch up paint in area of repair (paragraph
5-13a and TM 55-1500-345-23).
Do not saturate the bond lines with MEK as it will soften
the adhesive.
(a) Clean area around cracks in adhesive at bond line
in area illustrated in figure 5-60 with MEK (C87). Dry
area with clean cloths.
(5) Replace loose or damaged buffer pads.
NOTE
(b) Apply a thin film of adhesive (C17) to the bond
lines shown in figure 5-60 areas A and B.
Exercise extreme care to ensure grip
plates are not gouged or otherwise
damaged.
(c) Apply a small bead of adhesive (C17) to area
shown in figure 5-60.
(a) Remove old pad, using a knife or similar tool.
(b) Remove any remaining adhesive by sanding in
a spanwise direction to bare metal. Surface finish shall
be 32 RMS or better.
NOTE
Use cellophane or similar material
between washers and buffer pads to
prevent adhesive squeeze-out from
contacting washers.
(d) Swab the inside diameter of the leading edge drain
hole with a thin film of adhesive (Cl7) to ensure that
adequate sealing exists (figure 5-60).
(e) Touch up paint in repair area (paragraph 5-13c and
TM 55-1500-345-23).
(f) Balance the tail rotor hub and blade assembly prior
to installation on helicopter (paragraph 5-12).
(c) Bond new buffer pads to blade, using adhesive
(C17). Apply pressure to pad by installing a 0.5 inch
diameter bolt through a blade bolt hole with an AN970-8
washer under both the bolt head and the nut. Tighten
nut to apply pressure on buffer pad.
(d) Refinish blade as required.
(6) Repair blades with cracks in adhesive at bond line
detected in inspection described in paragraph 5-13c, as
follows:
(7) Repair pitch horn as follows:
(a) Polish out mechanical and corrosion
damage on pitch horns. Use 300 grit or finer sandpaper
(C112) and scotchbrite (C113). Inspect repaired areas to
ensure that limits shown in figure 5-61 have not been
exceeded.
(b) Touch up repaired areas with chemical film (C37).
(c) Replace pitch horn (5, figure 5-56) if hole for
floating bushing exceeds 0.5005 inch or is corroded.
(d) Replace bushing (27, fig. 5-54) if bolt is loose in
bushing or if bushing is loose in pitch horn.
e. Painting (A VIM).
Cleaning solvent is flammable and
toxic. Provide adequate ventilation.
Avoid prolonged breathing of vapors
and contact with skin or eyes.
(1) Paint touch-up is required when paint is
deteriorated and/or the paint is removed to repair
scratches, nicks, or dents. Refer to TM 55-1500-345-23.
5-112
Change 65
TM-55-1520-234-23
(2) Prepare blade for painting as follows:
(a) Polish out nick, scratch, and dent damage
(Paragraph 5-13d).
Cleaning solvent is flammable and toxic. Provide
adequate ventilation. Avoid prolonged breathing of
vapors and contact with skin or eyes.
Cleaning solvent is flammable and toxic. Provide
adequate ventilation. Avoid prolonged breathing of
vapors and contact with skin or eyes.
(b) Clean area where paint is to be applied with
aliphatic naphtha (C88).
(c) Mask off or plug retention bolt holes and mask
off holes for attaching the pitch horns to revent entry of
refinishing materials.
(d) Remove all surface oxides and aged paint from
aluminum surfaces.
(e) Wash blade with cleaning and polishing
compound (C39). Thoroughly rinse soap :from blade
and check for a water break free surface, which is, a
continuous unbroken film of water on the surface.
Repeat washing as required until the water break free
surface is attained.
Do not touch blades with bare hands during
remaining procedures or quality of paint will
be adversely affected.
(f) Apply coat of chemical film (C37) to bare
aluminum surfaces.
NOTE
If chemical film is not available, substitute
commercial "Metal-Prep", alcoholic phosphoric solution
(C23) or solution of chromic acid (C3).
(3) Apply finish to blade as follows:
(b) Mix a small quantity of adhesive (C7) according
to directions on container. Mix 13 to 15 percent by
weight of primer (C100) into the adhesive (C7). Mix
thoroughly and thin to sprayable consistency by adding
MEK (C87). Do not exceed 50 percent by volume; 35
percent should produce a sprayable consistency. The
pot life of the epoxy primer mixture is approximately
three hours.
(c) Ensure that masking tape applied in step b. (3) is
still in place. Apply three wet spray coats of the
adhesive prepared in the preceding step to all surfaces
at the root of the blade for a distance of 0.750 inch to 2
inches outboard of the perimeter of the doublers. Allow
each coat to dry 45 to 60 minutes. Make each coat 1.5
to 2.0 mils thick. Apply one wet spray coat of the same
adhesive material to the entire length of the blade on
both sides. Use the leading edge of the skin as the
centerline of the spray.
(d) After the final coat is applied in preceding step,
allow the blade to air dry for 16 TO 24 hours.
(e) Apply one thin mist coat of primer (C100) to all
touch up areas and allow to dry for a minimum of 45
minutes and a maximum of 8 hours prior to applying
next coat.
NOTE
It is necessary to cover all touch up areas with
primer (C100) and comply with the time limit noted in
step (e) or the finish coat of lacquer will not adhere to
the blade.
(f) Apply the final coats of lacquer to the touch up
areas only. Use lacquer (C76) to touch up areas except
on the blade tip. Use lacquer (C77) on
(a) Apply one coat of primer (C100) to the touch up
area.
Change 22
5-113
TM 55-1520-234-23
touch up areas of the six inch wide band on the blade
tip. Paint thickness to be approximately 1.2 TO 1.5 mils.
(4) Install the opposite pitch horn in the same manner.
(g) Air dry the blade for 3 hours prior to handling and
for 48 hours prior to flying. If a faster cure time is
required, air dry the blade 1 hour. Remove the masking
tape and oven dry the blade at 180 TO 190 degrees
F(82 TO 88 degrees C) for 1 hour.
(h) Apply a coating of corrosion preventive compound
(C53) to the inside of the retention bolt bushings.
f. Installation
Check for correct washers with
chamfered internal diameter under bolt
head.
(1) Position hub assembly (1, figure 5-56) on bench
with data plate side up. Slide blade (14) on hub yoke
with the data plate side up. Install bolts (20 and 25) with
special washers (19 and 26) under bolt heads. Install
special washers (2 and 13) next to blade. If special
balance washers were indexed at time of disassembly,
reinstall them in the same position. If they were not
indexed, do not install them until the assembly is
balanced. Install nuts (5 and 10) but do not torque until
after the assembly has been balanced.
(2) Install opposite blade in the same manner. The
four blade retention bolts (20 and 25) may be installed
from either side but all four bolts must be installed from
the same side.
(3) Position pitch horn (15, figure 5-56) on blade and
install bolts (22 and 23) with steel washers (21 and 24)
under heads. Install bolts with heads facing same
direction as blade retention bolts (20 and 25, figure 556). Install steel washers (6 and 9, figure 5-56) and nuts
(7 and 8). Torque nuts 50 TO 70 inch-pounds. If special
washer (18) was indexed at disassembly, install it at this
time with steel washer (17) and bolt (16). If special
washer (18) was not indexed, install steel washer (17)
and bolt (16). Do not torque until assembly has been
balanced.
5-114 Change 56
TM 55-1520-234-23
5-14. Tail Rotor Hub Assembly.
Premaintenance requirements for tail rotor hub.
Condition
Model
Part No. or Serial No.
Special Tools
Test Equipment
Support equipment
Minimum Personnel
Required
Consumable Materials
Special Environmental
Condition
Requirements
AH-1S
212-010701-9
(T57)
NA
Drill press
One
(C32) (C45) (C102)
(C112) (C119) (C124)
NA
a. Disassembly. (AVIM)
(1) Remove nut (9, figure 5-62), washer (8), bolt (1),
and washer (2). Remove similar parts on the opposite
side of the trunnion.
(2) Remove trunnion halves (3 and 7). The trunnion
halves are a matched set; keep those parts together for
reinstallation on the same hub.
Use solvent in a well ventilated area. Avoid
prolonged breathing of vapors and do not use in an area
with open flame or high temperature.
Use solvent in a well ventilated area.
Avoid prolonged breathing of vapors
and do not use in an area with open
flame or high temperature.
b. Cleaning. (AVIM) Clean the disassembled tail
rotor hub assembly parts with solvent (C124).
c. Inspection. (AVIM)
(1) Inspect the tail rotor historical records, and the
tail rotor hub for evidence that the tail rotor has been
subjected to an accident or incident outside the realm of
normal usage. If such evidence exists, perform
applicable special inspections for overspeed, sudden
stoppage, engine compressor stall and overtorque
outlined in chapter 1 and the following.
(a) Inspect yoke (5, figure 5-62) and trunnion halves
for obvious damage such as deformation. Reject hub if
any part is deformed.
TM 55-1520-234-23
Figure 5-62. Tail rotor hub assembly
5-115
TM 55-1520-234-23
(b) Inspect parts for surface damage in excess of
limits shown on figures 5-63 and 5-64.
(c) Inspect trunnion retention bolts (1, figure 5-62)
and blade retention bolts for shear offset. Reject hub if
any bolts show shear offset.
(d) Inspect bearings (4 and 6) for looseness in yoke.
Replace yoke if bearings are loose.
(e) Inspect yoke (5, figure 5-62) and trunnion halves
(3 and 7) by fluorescent penetrant method, MIL-I-6866.
(1) Replace bearings (4 and 6, figure 5-62) which did
not pass inspection.
(a) Press faulty bearings out of yoke. Use a sleeve of
slightly larger diameter than the bearing outside
diameter to support the yoke. Use a sleeve of slightly
smaller diameter than the bearing sleeve to press the
bearing and bearing sleeve out of the yoke.
(b) Clean aged primer and dirt from the yoke.
NOTE
NOTE
When finish is removed from yoke, repaint, using
primer (C100) and overspray with lacquer (C75).
Bearing removal is not necessary for yoke penetrant
inspection. Bearings should be covered to prevent entry
of penetrant.
(2) Normal Inspection of tail rotor hub:
(a) Move bearings (4 and 6, figure 5-62) through lull
throw and check for corrosion. If any corrosion is
detected, remove the products of corrosion with
Scotchbrite (C 113) and clean cloths.
(b) Check bearings (4 and 6) for indication of teflon
deterioration, bond separation, and teflon protruding
from bearing. Replace bearings if any of these
conditions exist. Check bearings (4 and 6) for axial
looseness. Maximum allowable axial looseness is 0.015
inch. Replace bearings if axial limits are exceeded.
NOTE
Due to the spherical shape of the bearing, any
looseness in the radial direction will also result in axial
looseness. For this reason, measuring axial looseness
only is adequate for determining bearing serviceability.
(c) Deleted.
(d) Inspect yoke for mechanical and corrosion
damage. If damage exceeds limits shown on figure 563, scrap the yoke.
(e) Inspect trunnion set for mechanical and corrosion
damage and also wear damage on the splines. If
damage and/or wear exceeds the limits shown on figure
5-64 on either trunnion half, scrap both halves of the
set.
d. Repair. (AVIM)
Remove preservative oil from bearings (6) using
solvent (C124 or C142) prior to installing new bearings.
(c) Apply primer.(C102) to mating surfaces of yoke
and of new bearing.
(d) Position bearing in yoke while primer (C102) is
still wet. Use a backstop so that bearing can be set with
one operation. Clean off excess primer and prevent it
from entering bearing.
(e) Select proper anvil (backstop) and staking tool
from staking tool set (T57) for the bearing being
installed. Use staking tool, T101577-11, and anvil
(backstop), T101557-13, to stake bearings (6, figure 562). Use staking tool, T101577-17, and anvil (backstop),
T101577-19, to stake bearings (4, figure 5-62). See
figure 5-65, sheet 1 and sheet 2 for views of assembled
staking tools.
(f) Install staking tool, selected in preceding step,
in chuck of a hand-feed-type drill press. Lubricate
staking tool rollers with lubricating oil (C92).
(g) Place anvil (backstop) selected in step (e) on drill
press table with flanged side down. See figure 5-66,
detail A. Position yoke on drill press with data plate side
up and with bearing to be staked in contact with anvil.
Lower the drill press chuck and the staking tool and
check to ensure that the anvil bearing and staking tool
are aligned and that the staking tool rollers are in
contact with the groove in the bearing.
(h) Set drill press speed at 250 to 350 rpm. Start the
drill press and apply steady hand pressure. to the feed
lever for a minimum of ten seconds. Raise drill press
chuck and staking tool. Check to determine the amount
that the bearing has been staked in comparison with
figure 5-66, view B.
5-116 Change 65
TM 55-1520-234-23
NOTE
Rotating the yoke 90 degrees while maintaining
applied pressure will help to ensure uniform
staking.
(i) Reposition the anvil (backstop) with the flanged
side up. Invert the yoke and stake the opposite side of
the bearing in the same manner. Repeat staking the
bearing in small increments to attain the 0.008 inch
maximum gap shown on figure 5-66, detail B.
(j) Check the bearing axial position in the yoke to
ensure that it is within the limits shown on figure 5-66,
detail B.
(k) Check the bearing by feel for smooth operation.
Check the bearing visually to ensure that the ball was
not
scored
by
the
staking
tool.
Change 65
5-116A/(5-116B blank)
TM 55-1520-234-23
Figure 5-63. Damage limits - tall rotor hub yoke
Change 29
5-117
TM 55-1520-234-23
Figure 5-64. Damage limits - tail rotor trunnion set
5-118 Change 7
TM 55-1520-234-23
Figure 5-65. Bearing staking tool P/N T101577 (Sheet 1 of 2)
Change 7
5-119
TM 55-1520-234-23
Figure 5-65. Bearing staking tool P/N T101577 (Sheet 2 of 2)
5-120 Change 7
TM 55-1520-234-23
Figure 5-66. Tool application - bearing installation (staking) in tail rotor yoke
5-121
TM 55-1520-234-23
(2) Polish out any mechanical damage and corrosion
damage on the yoke and on the trunnion that was not
accomplished during inspection. Polish out damage on
trunnion splines as well as the outer surfaces. Use
crocus cloth (C45), sandpaper (C 112), and india stone
(C128) of correct grades to polish out the damage and
to leave a smooth, scratch-free surface. If damage
exceeds the limits shown on figures 5-63 and 5-4, scrap
the affected part.
(3) Touch up repair areas with brush cadmium plate
(C32) or primer (C102). Do not apply coating to trunnion
splines.
side of gearbox.
(6) Remove nuts and washers (38, figure 5-67).
Remove housing (40). Remove retaining ring (42),
housing (41), excluder (43) and bearing (52).
b. Disassembly. See figure 5-67.
(1) Remove tail rotor retaining nut (36) and shield
(37) from crosshead (131 if not previously
,accomplished.
(2) Remove cotter pin (10). Remove nut (11), bolt
(15) and pitch link (31). Remove opposite pitch link in
the same manner.
e. Assembly. See figure 5-62.
(1) Inspect trunnion halves (3 and 7) to ensure
that they are a matched set.
(2) Position trunnion half (3) on the data plate
side of yoke (5). Position trunnion half (7), with the static
stop ear, on the opposite side of yoke with the master
spline of trunnion halves aligned.
Check for correct washers with chamfered internal
diameter under bolt head.
(3) Remove nut (19), bolt (16), washer (18) and
weights (17). The quantity of weights and washers may
vary from those illustrated. Record the quantity of
weights and washers removed. Remove remaining
weights in the same manner.
(4) Remove cotter pin (21). Remove nut (20),
washer (22), bolt (39), washer (29) and counterweight
link (26). Remove opposite counterweight link in the
same manner.
(5) Remove cotter pin (23). Remove nut (24),
washer (25), bellcrank (27), and washer (29). Remove
opposite bellcrank in the same manner.
(3) Install two bolts (1) with washers (2) under
bolt heads and washers (8) under nuts (9). Torque nuts
evenly to 500 inch-pounds.
Use solvent in a well ventilated area. Avoid
prolonged breathing of vapors and do not use in an.
area with open flame or high temperature.
5-15. Tail Rotor Controls - Crosshead, Weights,
Links and Control Tube.
The tail rotor controls crosshead, weights, links and
control tube and associated linkage control tail rotor
pitch. The control tube is attached to the anti-torque
pedals through control linkage. The crosshead, weights
and links balance control forces and assist in controlling
blade pitch.
c. Cleaning. Clean the parts disassembled in
paragraphs a. and b. with solvent (C124) and dry
with filtered compressed air. Use only solvent
(C124) on bearing (8, figure 5-67).
d. Inspection.
a. Removal. See figure 5-67.
(1) Remove crosshead, weights and links as an
assembly. Refer to paragraph 5-12, step a.
(2) Remove counterweight support. Refer to
paragraph 5-12, step a.
(3) Slide sleeve (35, figure 5-67) outboard on
control tube (34) slightly and rotate counterclockwise to
engage internal threads in sleeve. Continue to rotate
counterclockwise until threads are disengaged and
remove sleeve.
(4) Remove attaching bolts and remove lever (48,
figure 5-67), idler (47) and link (51).
(5) Remove control tube (34, figure 5-7) from left
(1) Inspect cross head for damage in excess of
limits shown on figure 5-68.
(2) Inspect counterweight bellcrank for wear and
damage in excess of limits shown on figure 569.
Maximum allowable radial play between bellcrank (item
27, figure 5-67) and crosshead (13) is 0.010 inch.
Maximum axial play not to exceed .015 inch.
(3) Inspect pitch links for damage in excess of
limits shown on figure 5-70.
(4) Inspect counterweight links for damage in
excess of limits shown on figure 5-71.
5-122 Change 54
TM 55-1520-234-23
Figure 5-67. Tail Rotor Control - Crosshead, Weights, Links and Control Tube (Sheet 1 of 3)
Change 65
5-123
TM 55-1520-234-23
Figure 5-67. Tail rotor control - crosshead, weights, links and control tube (Sheet 2 of 3)
5-124 Change 65
TM 55-1520-234-23
Figure 5-67. Tail Rotor Control - crosshead, weights, links and control tube (Sheet 3 of 3)
5-125
TM 55-1520-234-23
Figure 5-68. Damage limits - tail rotor control crosshead
5-126 Change 7
TM 55-1520-234-23
Figure 5-69. Damage limits - tail rotor control counterweight bellcrank
Change 7
5-127
TM 55-1520-234-23
Figure 5-70. Damage limits - tail rotor control pitch link
5-128 Change 65
TM 55-1520-234-23
Figure 5-71. Damage limits - tail rotor control counterweighrt link
Change 7
5-129
TM 55-1520-234-23
Figure 5-72. Damage limits - tail rotor active counterweighrt support
5-130 Change 7
TM 55-1520-234-23
Figure 5-73. Damage limits - tail rotor control tube
(5) Inspect counterweight support for damage in
excess of limits shown on figure 5-72.
NOTE
The control tube support can be any fixture, locally
manufactured which can be placed on a flat surface and
the tube rotated along the center axis in order To check
for straightness using a dial indicator.
(6) Inspect tail rotor control tube for damage in
excess of limits shown in figure 5-73. Mount the control
Change 65
tube on centers and check run out. Maximum allowable
run out in areas designated Area A is 0.010 inch.
Maximum allowable run out in other areas is 0.020 inch.
Inspect two corks for secure installation in tube and for
damage. Inspect threads for damage.
(7) Inspect link assembly for damage in excess of
limits shown on figure 5-74.
(8) Inspect idler for damage in excess of limits
shown on figure 5-75.
5-131
TM 55-1520-234-23
(9) Inspect lever for damage in excess of limits
shown on figure 5-76. Inspect bearings in lever for wear.
(10) Inspect housing (40, figure 5-67) for corrosion
and mechanical damage.
(11) Inspect retaining ring (42, figure 5-67) for
damage that would affect function.
(12) Inspect housing (41, figure 5-67) and excluder
(43) for damage that would affect function.
(13) Inspect bearing (52, figure 5-67) for damage that
would affect function.
(14) Inspect shield (37, figure 5-67) for cuts and
deterioration.
(15) Inspect tail rotor retaining nut (36, figure 5-67)
for damaged threads, corrosion, nicks and dents.
(16) Inspect race (44, figure 5-67) for corrosion
and wear that would affect function.
(17) Inspect sleeve (35, figure 5-67) for corrosion,
wear, and damaged threads.
(18) Inspect bearing (8, figure 5-67) for evidence of
separation of races. Rotate bearing manually. If
roughness is noted or if axial looseness exceeds 0.005
inch, replace bearing.
5-132
(19) Inspect following parts by magnetic particle
method, MIL-I-6868 Code M, or fluorescent penetrant
method, MIL-I-6866 Code F. If damage occurs inspect
following parts. See figure 5-67.
ITEM
NOMENCLATURE
CODE
13
Tail Rotor Controls Crosshead
F
26
Counterweight Link
F
27
Counterweight Bellcrank
F
34
Tail Rotor Control Tube
M
33
Counterweight Support
F
36
Tail Rotor Retaining Nut
M
47
Idler
F
48
Lever Assembly
F
51
Link Assembly
M
e. Repair.
(1) Replace any part which failed to pass inspection
in paragraph d.
(2) Polish out mechanical and corrosion damage that
is within limits specified in paragraph d. Touch up repair
areas with primer (C102).
Change 65
TM 55-1520-234-23
WARNING
Use solvent (C124) in a well ventilated area.
Avoid prolonged breathing of vapors and do
not use in an area with open flame or high
temperature.
(3) Replace loose or damaged corks in control tube.
Clean area where cork will be installed with solvent
(C124) and allow to dry. Coat new cork with shellac
(C120) and press into position illustrated on figure 5-73.
(4) Remove worn bushing in bellcrank (27, figure 567). Install new bushing using wet zinc chromate primer
(C102) between mating surfaces.
f Assembly. See figure 5-67.
(1) Position nylatron washer (28), bellcrank (27) and
special washer (25) on crosshead (13). In- stall nut (24)
and torque 70 TO 125 inch-pounds. Install cotter pin
(23). Install opposite bellcrank in the same manner. If
cotter pins do not engage castellations in nuts, add one
thin steel washer under nuts (24).
opposite counterweight link in the same manner. If
cotter pins do not engage castellations in nuts, add one
thin steel washer under nuts (20).
(4) Adjust both pitch links (31) to 6.105 to 6.125 inch
dimension between centers of bearings. This will yield a
distance of 2.72-2.92 inches between inside face of jam
nuts.
(5) Position pitch link (31), in crosshead (13) with the
end with rivet (32) installed away from the crosshead as
illustrated. Install bolt (15) with steel washer (14) under
head and steel washer (12) under nut (11). Torque nut
110 TO 165 inch-pounds and install cotter pin (10).
Install opposite pitch link in same manner. If cotter pins
do not engage castellations in nuts, add one thin steel
washer under nuts (11).
g. Installation. See figure 5-67.
5-15.1. Disassembly-Tail Rotor Pitch
Change Link (AVIM).
a. Remove rivet (32, figure 5-67) from rod end.
b. Loosen jam nuts (55) and remove rod end
assemblies (54).
CAUTION
Weights (17, figure 5-67) must be installed on
outboard side of bellcranks (27) as illustrated or
interference may occur.
(2) Install weights on each end of the two bellcranks
(27). Install the same weights that were removed if they
were indexed; if not, install two weights (17) and one
steel washer (18) at each of the four locations. Install
bolt (16) and nut (19). Torque nut 50 TO 70 inchpounds.
(3) Position counterweight link (26) in bellcrank (27)
and install bolt (30) with steel washer (29) under head
and steel washer (22) under nut (20). Torque nut 60 TO
110 inch-pounds and install cotter pin (21). Install
c. Remove jam nuts (55).
5-15.2. Assembly-Tail Rotor Pitch Change
Link (AVIM).
a. Thread ja