Bell 407 helicopter Maintenance Manual
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REVISION ONLY BHT-407-MM
NOTICE
This Revision 25 updates the front matter, as well as Chapter 4, which incorporates new tailboom inspection requirements per ASB 407-08-84. In addition, Chapter 76 incorporates information on FADEC Software Version 5.358. General updates to
Chapters 7, 52, 53, 63, 65, 71, and 95 have also been included.
Please incorporate Revision 25 into the Maintenance Manual in accordance with the Log of Pages attached.
Bell Helicopter would also like to thank its Customers for providing us with
Customer Feedback information. This information is very much appreciated and allows us to improve the quality of our manuals with each revision.
22 FEBRUARY 1996
REVISION 25 — 29 SEPTEMBER 2008
BHT-407-MM-1
MAINTENANCE MANUAL
VOLUME 1
GENERAL INFORMATION
NOTICE
The instructions set forth in this manual, as supplemented or modified by Alert Service Bulletins (ASB) or other directions issued by Bell Helicopter Textron Inc. and Airworthiness Directives (AD) issued by the applicable regulatory agencies, shall be strictly followed.
COPYRIGHT NOTICE
COPYRIGHT 2008
BELL ® HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON
CANADA LTD.
ALL RIGHTS RESERVED
22 FEBRUARY 1996
REVISION 25 — 29 SEPTEMBER 2008
BHT-407-MM-1
PROPRIETARY RIGHTS NOTICE
These data are proprietary to Bell Helicopter Textron Inc. Disclosure, reproduction, or use of these data for any purpose other than helicopter operation or maintenance is forbidden without prior written authorization from
Bell Helicopter Textron Inc.
PN
Additional copies of this publication may be obtained by contacting:
Commercial Publication Distribution Center
Bell Helicopter Textron Inc.
P. O. Box 482
Fort Worth, Texas 76101-0482
Rev. 22 28 NOV 2005
WARNING
THIS MANUAL APPLIES ONLY TO HELICOPTERS AND COMPONENTS
MAINTAINED IN ACCORDANCE WITH BELL HELICOPTER TEXTRON
(BELL) APPROVED PROCEDURES USING BELL APPROVED PARTS.
ALL INSPECTION, REPAIR AND OVERHAUL PROCEDURES PUBLISHED
BY BELL, INCLUDING PART RETIREMENT LIFE, ARE BASED SOLELY ON
THE USE OF BELL PARTS THAT HAVE BEEN MAINTAINED USING BELL
APPROVED DATA. THE DATA PUBLISHED HEREIN OR OTHERWISE
SUPPLIED BY BELL IS NOT APPLICABLE TO NON-BELL PARTS OR
PARTS THAT HAVE BEEN REPAIRED USING DATA AND/OR PROCESSES
NOT APPROVED BY BELL.
BELL IS NOT RESPONSIBLE FOR ANY PART OTHER THAN THOSE THAT
IT HAS APPROVED.
BEFORE PERFORMING ANY PROCEDURE CONTAINED IN THIS MANUAL
YOU MUST INSPECT THE AFFECTED PARTS AND RECO RDS FOR
EVIDENCE OF ANY MANUFACTURE, REPAIR, REWORK OR USE OF A
PROCESS NOT APPROVED BY BELL.
IF YOU IDENTIFY OR SUSPECT THE USE OF PARTS NOT AUTHORIZED
BY BELL, EITHER REMOVE THE AFFECTED ITEM FROM THE AIRCRAFT
OR OBTAIN INSTRUCTIONS FOR CONTINUED AIRWORTHINESS FROM
THE MANUFACTURER OR THE ORGANIZATION THAT APPROVED THE
REPAIR.
BHT-407-MM-1
29 SEP 2008 Rev. 25 Warning
BHT-407-MM-1
CUSTOMER SUPPORT AND SERVICES
Flying smart means that no matter where you are, or what time it is, you can make a call and get additional information, clarification, or advice on a technical or operational issue concerning your helicopter or information contained in our Technical Publications.
Product Support Engineering (PSE) is just a phone call away and may be contacted as follows:
MODEL 47, 206, OR 407
Phone: 450-437-2862 or 800-243-6407 (U.S./Canada)
Fax:
E-mail:
450-433-0272 [email protected]
MODEL 222, 230, 427, 429, OR 430
Phone: 450-437-2077 or 800-463-3036 (U.S./Canada)
Fax:
E-mail:
450-433-0272 [email protected]
MODEL 204, 205, 212, OR 412
Phone:
Fax:
E-mail:
450-437-6201 or 800-363-8028 (U.S./Canada)
450-433-0272 [email protected]
MODEL 210, 214, HUEY II, SURPLUS MILITARY
Phone: 817-280-3548
Fax:
E-mail:
817-280-2635 [email protected]
For additional information on Customer Support and Services as well as Product Support
Engineering (PSE) and your local Customer Service Representative (CSR) network, please access http://www.bellhelicopter.com/en/support .
CSS Rev. 25 29 SEP 2008
BHT-407-MM-1
LOG OF REVISIONS
Insert latest revision pages and dispose of superseded ones.
On a revised page, the text and/or illustration affected by the latest revision is shown by a vertical line. A revised page with only a vertical line next to the page number indicates that text has shifted or that non-technical correction(s) were made on that page.
Original ........ 0 ............ 22 FEB 96
Revision....... 1 ............01 APR 96
Revision....... 2 ............ 01 JUN 96
Revision....... 3 ........... 28 OCT 96
Revision....... 4 ........... 16 DEC 96
Revision....... 5 .............18 JUL 97
Revision....... 6 ............14 APR 98
Revision....... 7 ........... 01 MAY 98
Revision....... 8 ............01 SEP 98
Revision ...... 9............30 NOV 98
Revision ...... 10.......... 16 FEB 01
Revision ...... 11 ......... 23 MAR 01
Revision ...... 12..........08 MAY 01
Revision ...... 13.......... 21 SEP 01
Revision ...... 14..........21 NOV 01
Revision ...... 15...........18 JAN 02
Revision ...... 16.......... 07 JUN 02
Revision ...... 17..........25 OCT 02
Revision .......18 .........29 AUG 03
Revision .......19 ......... 17 DEC 03
Revision .......20 .........12 NOV 04
Revision .......21 .........16 NOV 04
Revision .......22 .........28 NOV 05
Revision .......23 .........07 NOV 06
Revision .......24 ......... 02 OCT 07
Revision .......25 ......... 29 SEP 08
LOG OF PAGES
Page No.
Rev No.
VOLUME 1
Cover .......................................23
Title ..........................................25
PN............................................22
Warning ...................................25
CSS .........................................25
A – D........................................25
HELP .......................................24
Customer Feedback ................24
Helicopter Sale Notice .............24
Warranty ..................................25
BR............................................25
TR ............................................23
i – ii ..........................................23
Chapter 1
1/2............................................23
3 – 6.........................................23
7/8............................................23
Chapter 4
1 – 34.......................................25
35/36........................................25
Chapter 5
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5/6............................................23
7 – 8.........................................23
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91 – 92 .................................... 24
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95/96 ....................................... 23
VOLUME 2
Cover....................................... 23
Title.......................................... 25
PN ........................................... 22
Chapter 6
1/2 ........................................... 18
3/4 ........................................... 18
5/6 ........................................... 18
7/8 ........................................... 18
9/10 ......................................... 18
Chapter 7
1/2 ........................................... 22
3 – 4 ........................................ 22
5 .............................................. 25
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Page No.
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9/10..........................................22
Chapter 8
1 – 40.........................................2
41/42..........................................2
Chapter 9
1/2............................................18
3 – 4.........................................18
5/6............................................18
Chapter 10
1/2............................................18
3 – 12.......................................18
Chapter 11
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Chapter 12
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29 SEP 2008 Rev. 25 A
BHT-407-MM-1
Page No.
Rev No.
Chapter 18
1............................................... 16
2 – 7........................................... 2
8 – 9......................................... 19
10 – 17....................................... 2
18............................................. 19
19 – 25....................................... 2
26............................................. 16
26A/26B................................... 16
27............................................. 16
28 – 42....................................... 2
43/44.......................................... 2
VOLUME 3
Cover ....................................... 23
Title.......................................... 23
PN............................................ 23
Chapter 21
1 – 20......................................... 4
Chapter 25
1 – 4......................................... 23
5/6............................................ 23
7 – 44....................................... 23
45/46........................................ 23
47 – 54..................................... 23
55/56........................................ 23
57 – 84..................................... 23
85/86........................................ 23
Chapter 26
1/2.............................................. 0
3 – 4........................................... 0
VOLUME 4
Cover ....................................... 18
Title.......................................... 22
PN............................................ 22
Chapter 28
1 – 64......................................... 4
65/66.......................................... 4
Chapter 29
1 – 2......................................... 22
3/4............................................ 22
5 – 64....................................... 22
LOG OF PAGES (CONT)
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VOLUME 5
Cover ...................................... 18
Title ......................................... 25
PN ........................................... 24
Chapter 30
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Chapter 32
1 – 2 .......................................... 4
3/4 ............................................. 4
5 – 24 ........................................ 4
25/26 ......................................... 4
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31 – 32 ...................................... 4
Chapter 52
1 – 2 ........................................ 24
3/4 ........................................... 24
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Chapter 53
1 – 50 ...................................... 24
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93 ............................................ 24
94 ............................................ 25
95 – 112 .................................. 24
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115 – 120 ................................ 24
VOLUME 6
Cover ...................................... 23
Title ......................................... 25
PN ........................................... 24
Chapter 62
1 – 10 ...................................... 23
11 – 12 .................................... 24
13 – 18 .................................... 23
19 – 20 .................................... 24
21 – 31 .................................... 23
32 ............................................ 24
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94 – 101................................... 23
102........................................... 24
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131........................................... 23
132 – 133................................. 24
134 – 135................................. 23
136........................................... 24
137 – 138................................. 23
139........................................... 24
140 – 141................................. 23
142........................................... 24
Chapter 63
1............................................... 25
2 – 24....................................... 23
25 – 28..................................... 25
29 – 39..................................... 23
40............................................. 25
41 – 43..................................... 23
44............................................. 24
45 – 66..................................... 23
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200........................................... 24
201 – 212................................. 23
213/214.................................... 23
VOLUME 7
Cover ....................................... 23
Title .......................................... 25
PN............................................ 24
Chapter 64
1 – 2......................................... 23
3/4............................................ 23
5 – 22....................................... 23
23/24........................................ 23
B Rev. 25 29 SEP 2008
BHT-407-MM-1
Page No.
Rev No.
25 – 40 .....................................23
41/42 ........................................23
43 – 62 .....................................23
Chapter 65
1 – 4 .........................................23
5/6 ............................................23
7 – 25 .......................................23
26 .............................................25
27 – 31 .....................................23
32 .............................................24
33 – 39 .....................................23
40 .............................................25
41 – 76 .....................................23
77 .............................................25
78 .............................................23
79 .............................................25
80 .............................................23
81 .............................................25
82 – 166 ...................................23
VOLUME 8
Cover .......................................23
Title ..........................................24
PN ............................................24
Chapter 67
1 – 94 .......................................23
95/96 ........................................23
97 – 189 ...................................23
190 – 191 .................................24
192 – 195 .................................23
196 ...........................................24
197 – 258 .................................23
259/260 ....................................23
261 – 310 .................................23
VOLUME 9
Cover .......................................23
Title ..........................................25
PN ............................................24
Chapter 71
1 – 20 .........................................7
21 .............................................25
22 ...............................................7
23/24 ..........................................7
25 – 30 .......................................7
31/32 ..........................................7
LOG OF PAGES (CONT)
Page No.
Rev No.
33 – 34....................................... 7
35/36.......................................... 7
37 – 44....................................... 7
45/46.......................................... 7
47 – 48....................................... 7
49/50.......................................... 7
Chapter 75
1/2............................................ 22
3/4............................................ 22
Chapter 76
1 – 8......................................... 25
9/10.......................................... 25
11 – 46..................................... 25
47/48........................................ 25
49 – 52..................................... 25
53/54........................................ 25
55 – 68..................................... 25
69/70........................................ 25
71 – 86..................................... 25
87/88........................................ 25
Chapter 79
1 – 8........................................... 4
9/10............................................ 4
11 – 14....................................... 4
15/16.......................................... 4
VOLUME 10
Cover ....................................... 24
Title.......................................... 25
PN............................................ 24
Chapter 95
1 – 8......................................... 24
9 – 14......................................... 4
15/16.......................................... 4
17 – 18..................................... 16
19 – 72....................................... 4
73............................................. 25
74 – 80....................................... 4
81 – 82..................................... 24
83 – 92....................................... 4
93 – 97..................................... 24
98 – 101..................................... 4
102 – 108................................. 24
109 – 110................................... 4
111 ........................................... 25
112 – 123................................... 4
Page No.
Rev No.
124 ...........................................24
125 – 133 ...................................4
134 ...........................................24
135 .............................................4
136 ...........................................24
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147 ...........................................24
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156 ...........................................24
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175 – 177 .................................24
178 .............................................4
179/180 ......................................4
181 – 193 ...................................4
194 ...........................................24
195 – 206 ...................................4
207/208 ......................................4
Chapter 96
1 – 20 .......................................16
21 – 22 .....................................24
23 – 26 .....................................16
27 – 30 .....................................24
31 – 102 ...................................16
103/104 ....................................16
105 – 176 .................................16
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VOLUME 11
Cover .......................................18
Title ............................................6
PN ..............................................6
Chapter 97
1 – 6 ...........................................4
7/8 ..............................................4
9 – 72 .........................................4
73/74 ..........................................4
75 – 80 .......................................4
29 SEP 2008 Rev. 25 C
BHT-407-MM-1
Page No.
Rev No.
81/82.......................................... 4
83 – 88....................................... 4
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91 – 98....................................... 4
99/100........................................ 4
101 – 116................................... 4
117/118 ...................................... 4
VOLUME 12
Cover ....................................... 18
Title.......................................... 16
PN.............................................. 6
Chapter 98
1................................................. 7
2............................................... 16
3/4.............................................. 4
5 – 10......................................... 4
11 – 12....................................... 7
LOG OF PAGES (CONT)
Page No.
Rev No.
13 .............................................. 4
14 .............................................. 7
15 – 16 ...................................... 4
17 .............................................. 7
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29 .............................................. 7
30 .............................................. 4
31/32 ......................................... 4
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36 .............................................. 4
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49 .............................................. 4
50 .............................................. 7
Page No.
Rev No.
50A/50B................................... 16
51 – 53....................................... 4
54 – 60....................................... 7
VOLUME 13
Cover ....................................... 18
Title .......................................... 15
PN.............................................. 6
Chapter 99
1/2.............................................. 0
3............................................... 15
4................................................. 8
MMS – 1 .................................... 1
MMS – 5 .................................... 0
MMS – 6 .................................... 0
MMS – 7 .................................... 0
MMS – 21 .................................. 0
D Rev. 25 29 SEP 2008
BHT-407-MM-1
SPARE PARTS WARRANTY
ONE YEAR/1,000 HOURS PRORATED
WARRANTY AND REMEDY: Seller warrants each new helicopter part or helicopter part reconditioned by seller to be free from defect in material and workmanship under normal use and service and if installed on Bell model helicopters. Seller’s sole obligation under this warranty is limited to replacement or repair of parts which are determined to Seller’s reasonable satisfaction to have been defective with 1,000 hours of operation or one (1) year after installation, whichever occurs first and reimbursement of reasonable freight charges.
After 200 hours of use, there will be a prorated charge to the Purchaser for replacement parts (prorating the hours of total use against the then applicable part life or 2,000 hours, whichever is the lesser). Defective parts must be reported in writing to the Seller’s Warranty
Administration within 90 days of being found defective. Replacement of parts may be with either new or reconditioned parts, at Seller’s election. Warranty adjustment is contingent upon the Purchaser complying with the Warranty Remedies as described in the
Commercial Warranty Information brochure and the Seller’s Warranty Administration disposition instructions for defective parts. Failure to comply with all of the terms of this paragraph may, at Seller’s sole option, void this warranty.
NOTE: Parts, components and assemblies of all new helicopters may have been restored or reworked due to mars, blemishes, dents or other irregularities during the manufacturing process. Such restoration and/or rework is permitted under Seller’s approved manufacturing and engineering processes and guidelines. The restoration and/or rework so completed does not render such items defective in material or workmanship.
THIS WARRANTY IS GIVEN AND ACCEPTED IN PLACE OF (i) ALL OTHER WARRANTIES OR CONDITIONS, EXPRESS OR
IMPLIED, INCLUDING BUT NOT LIMITED TO THE IMPLIED WARRANTIES OR CONDITIONS OF MERCHANT ABILITY AND
FITNESS FOR A PARTICULAR PURPOSE AND (ii) ANY OBLIGATION, LIABILITY, RIGHT, CLAIM OR REMEDY IN CONTRACTOR
IN TORT (DELICT), INCLUDING PRODUCT LIABILITIES BASED UPON STRICT LIABILITY, NEGLIGENCE, OR IMPLIED
WARRANTY IN LAW.
This warranty is the only warranty made by Seller. The Purchaser’s sole remedy for a breach of this warranty or any defect in a part is the repair or replacement of helicopter parts and reimbursement of reasonable freight charges as provided herein. Seller excludes liability, whether as a result of a breach of contract or warranty, negligence or strict product liability, for incidental or consequential damages, including without limitation, damage to the helicopter or other property, costs and expenses resulting from required changes or modifications to helicopter components and assemblies, changes in retirement lives and overhaul periods, local customs fees and taxes, and costs or expenses for commercial losses or lost profits due to loss of use or grounding of helicopters or otherwise.
Seller makes no warranty and disclaims all liability in contract or in tort (delict), including, without limitation, negligence and strict tort
(delictual) liability, with respect to work performed by third parties at Purchaser’s request and with respect to engines, engine accessories, batteries, radios, and avionics, except Seller assigns each manufacturer’s warranty to Purchaser to the extent such manufacturer’s warranty exists and is assignable.
This warranty shall not apply to any helicopter part which has been repaired or altered outside Seller’s factory in any way so as, in
Seller’s judgment, to affect its stability, safety or reliability, or which has been subject to misuse, negligence or accident, or which has been installed in any aircraft which has been destroyed unless that helicopter has been rebuilt by Bell. A list of destroyed aircraft is obtainable from Bell Product Support. Repairs and alterations which use or incorporate parts and components other than genuine Bell parts or parts approved by Bell for direct acquisition from sources other than Bell itself are not warranted by Bell, and this warranty shall be void to the extent that such repairs and alterations, in Seller’s sole judgment, affect the stability, safety or reliability of the helicopter or any part thereof, or damage genuine Bell or Bell-approved parts. No person, corporation or organization, including Bell Customer
Service Facilities, is authorized by Seller to assume for it any other liability in connection with the sale of its helicopters and parts.
NO STATEMENT, WHETHER WRITTEN OR ORAL, MADE BY ANY PERSON, CORPORATION OR ORGANIZATION, INCLUDING
BELL CUSTOMER SERVICE FACILITIES MAY BE TAKEN AS A WARRANTY NOR WILL IT BIND SELLER.
Seller makes no warranty and disclaims all liability with respect to components or parts damaged by, or worn due to, corrosion. Seller makes no warranty and disclaims all liability for consumables (wear items) which are defined as items required for normal and routine maintenance or replaced at scheduled intervals shorter than the warranty period. “Consumables” include but are not limited to engine and hydraulic oil, oil filters, packings and o-rings, anti-corrosion and/or sealing compounds, brush plating material, nuts, bolts, washers, screws, fluids, compounds, and standard aircraft hardware that is readily available to aircraft operators from sources other than Seller.
All legal actions based upon claims or disputes pertaining to or involving this warranty including, but not limited to, Seller’s denial of any claim or portion thereof under this warranty, must be filed in the courts of general jurisdiction of Tarrant County, Texas or in the United
States District Court for the Northern District of Texas, Ft. Worth Division located in Ft. Worth, Tarrant County, Texas. In the event that
Purchaser files such an action in either of the court systems identified above, and a final judgment in Seller’s favor is rendered by such court, then Purchaser shall indemnify Seller for all costs, expenses and attorneys’ fees incurred by Seller in defense of such claims. In the event Purchaser files such a legal action in a court other than those specified, and Seller successfully obtains dismissal of that action or transfer thereof to the above described court systems, then Purchaser shall indemnify Seller for all costs, expenses and attorneys’ fees incurred by Seller in obtaining such dismissal or transfer.
January 2007
29 SEP 2008 Rev. 25 Warranty
BHT-407-MM-1
BULLETIN RECORD
All applicable Alert Service Bulletins and Technical Bulletins issued prior to and including the bulletins listed below have been incorporated in this manual. Subsequent bulletins will be incorporated in future revisions/reissues.
1
ASB NUMBER
407-08-84
ALERT SERVICE BULLETINS
SUBJECT
Tailboom Assembly 407-030-801-201, 407-030-801-203, and
407-030-801-205, New Inspection Requirements
DATE
18 AUG 2008
1
With the exception of ASB 407-98-16, 407-98-18, 407-99-30, 407-00-36, 407-00-38, 407-01-39, 407-01-41,
407-01-42, and 407-02-50 all applicable bulletins issued prior to, and including 407-07-81, have been incorporated.
1
TB NUMBER
407-07-77
TECHNICAL BULLETINS
SUBJECT
Improved Caution/Warning Panel 407-375-015-111,
Introduction of
DATE
04 JUN 2007
1
With the exception of 407-99-15, 407-99-19, 407-00-26, 407-00-28, 407-03-50, 407-03-52, 407-04-55,
407-04-56, 407-05-68 and 407-06-74 all applicable bulletins issued prior to, and including 407-07-77, have been incorporated.
29 SEP 2008 Rev. 25 BR
BHT-407-MM-1
TEMPORARY REVISION RECORD
This Temporary Revision Record provides a current listing of active Temporary Revisions against the manual.
Temporary Revisions, which have been canceled/incorporated, will only be maintained on the record until the next revision is issued. If there are no Temporary Revisions shown on the record, this is confirmation that there are no
Temporary Revisions issued against the manual.
TEMP. REV. NO.
TITLE DATE ISSUED DATE CANCELED
NOTE: For tracking purposes, Temporary Revisions are now being numbered (Example: TR-1).
TR Rev. 23 7 NOV 2006
Figure
Number
4-1
4-2
4-3
4-4
4-5
4-6
4-7
4-8
4-9
TC APPROVED
Paragraph
Number
4-1
CHAPTER 4 — AIRWORTHINESS LIMITATIONS SCHEDULE
CONTENTS — MAINTENANCE PROCEDURES
Title
Chapter/Section
Number
AIRWORTHINESS LIMITATIONS SCHEDULE
Airworthiness Limitations Schedule ...........................................
4-00-00
BHT-407-MM-1
Page
Number
5
FIGURES
Title
Disc Pack Coupling (P/N 406-040-340-101) — Inspection.............................
Main Rotor Yoke (P/N 407-010-101-101) — Inspection.................................
Tail Rotor Hanger Bearing (P/N 406-040-339-ALL) — Inspection..................
Tail Rotor Blade (P/N 407-016-001-101) — Inspection ..................................
Tailboom Assembly (P/N 407-030-801-201, 407-030-801-203, and
407-030-801-205) — 300 Hour Inspection .....................................................
Tailboom Assembly (P/N 407-530-014-101, 407-530-014-103, or
407-030-801-107) — Daily Inspection............................................................
Tailboom Assembly (P/N 407-530-014-101, 407-530-014-103, and
407-030-801-107) — 100 Hour Inspection .....................................................
Inspection of Pylon Side Beams (P/N 407-010-201-105/407-010-203-105) ..
Horizontal Stabilizer (P/N 407-023-801-109) — Recurring 600 Hour/Annual
Inspection ......................................................................................................
TABLES
Airworthiness Limitations Schedule ................................................................
Page
Number
11
12
13
15
16
22
24
31
34
Page
Number
6
9
Chief Engineering
Aircraft Certification
Transport Canada
4-00-00
Page 1
BHT-407-MM-1
REVISION NO.
Revision 1 (01 Apr. 96)
Revision 2 (01 Jun. 96)
Revision 3 (26 Oct. 96)
Revision 4 (16 Dec. 96)
Revision 5 (18 Jul. 97)
Revision 6 (14 Apr. 98)
Revision 7 (01 May 98)
Revision 8 (01 Sept. 98)
Revision 9 (30 Nov. 98)
Revision 10 (16 Feb. 01)
Revision 11 (23 Mar. 01)
Revision 12 (08 May 01)
TC APPROVED
TC SIGNATURE DATE OF SIGNATURE
29 March 1996
N/A
26 October 1996
N/A
18 July 1997
14 April 1998
N/A
N/A
30 November 1998
N/A
N/A
8 May 2001
Chapter 4 not affected
Chapter 4 not affected
Chapter 4 not affected
Chapter 4 not affected
Chapter 4 not affected
Chapter 4 not affected
4-00-00
Page 2 Rev. 25 29 SEP 2008
TC APPROVED
REVISION NO.
Revision 13 (21 Sept. 01)
Revision 14 (21 Nov. 01)
Revision 15 (18 Jan. 02)
Revision 16 (07 Jun. 02)
Revision 17 (25 Oct. 02)
Revision 18 (29 Aug. 03)
BHT-407-MM-1
TC SIGNATURE DATE OF SIGNATURE
21 September 2001
21 November 2001
N/A
N/A
N/A
29 August 2003
Chapter 4 not affected
Chapter 4 not affected
Chapter 4 not affected
Revision 19 (17 Dec. 03)
Revision 20 (12 Nov. 04)
Revision 21 (16 Nov. 04)
Revision 22 (28 Nov. 05)
N/A
12 November 2004
16 November 2004
28 November 2005
Chapter 4 not affected
Revision 23 (7 Nov. 06)
Revision 24 (2 Oct. 07)
7 November 2006
2 October 2007
29 SEP 2008 Rev. 25
4-00-00
Page 3
BHT-407-MM-1
REVISION NO.
Revision 25 (29 Sept. 08)
DATE OF SIGNATURE
29 September 2008
TC APPROVED
TC SIGNATURE
4-00-00
Page 4 Rev. 25 29 SEP 2008
TC APPROVED BHT-407-MM-1
AIRWORTHINESS LIMITATIONS SCHEDULE
4-1.
AIRWORTHINESS LIMITATIONS
SCHEDULE
schedule applies to all the successive dash numbers for that component unless it is otherwise specified.
WARNING
WARNING
A L L R E P A I R A N D O V E R H A U L
PROCEDURES LIVES PUBLISHED BY
B E L L H E L I C O P T E R T E X T R O N ,
INCLUDING COMPONENT RETIREMENT
LIFE, ARE BASED SOLELY ON THE USE
O F B E L L A P P R O V E D P A R T S A N D
P R O C E S S E S . I F P A R T S O R
P R O C E S S E S D E V E L O P E D O R
APPROVED BY PARTIES OTHER THAN
BELL HELICOPTER ARE USED, THEN
THE DATA PUBLISHED OR OTHERWISE
SUPPLIED BY BELL HELICOPTER ARE
N O T A P P L I C A B L E . T H E U S E R I S
W A R N E D T O N O T R E L Y O N B E L L
HELICOPTER DATA FOR PARTS AND
PROCESSES NOT APPROVED BY BELL
H E L I C O P T E R. A L L A P P L I C A B L E
INSPECTIONS AND REPAIR METHODS
M U S T B E O B T A I N E D F R O M T H E
S U P P L I E R O F T H E P A R T S O R
PROCESSES NOT APPROVED BY BELL
HELICOPTER. BELL HELICOPTER IS
NOT RESPONSIBLE FOR PARTS OR
PROCESSES OTHER THAN THOSE
WHICH IT HAS ITSELF DEVELOPED OR
APPROVED.
The mandatory airworthiness limitations schedule
( Table 4-1 ) summarizes the mandatory maximum life, in hours, years or by Retirement Index Number (RIN) of components with a limited airworthiness life. Parts that are not on the schedule have an unlimited airworthiness life. The inspection limitations schedule
( Table 4-2 ) summarizes the mandatory inspection interval in hours.
Refer to the engine manufacturer's publications for the airworthiness limitations schedule of the engine and components.
NOTE
The airworthiness life or inspection interval for any part number contained in this
S O M E PA RTS A R E I NS TA L LE D A S
O R I G I N A L E Q U I P M E N T O N B O T H
M I L I T A R Y A N D C O M M E R C I A L
H E L I C O P T E R S A N D M A Y H A V E A
LOWER AIRWORTHINESS LIFE AND/OR
OVERHAUL SCHEDULE WHEN USED
O N A M I L I T A R Y H E L I C O P T E R. I N
A D D I T I O N , C I R C U M S T A N C E S
SURROUNDING THEIR USE MAY CALL
FOR OPERATION OF THE MILITARY
H E L I C O P T E R O U T S I D E O F T H E
AP P RO VE D C OM ME RCIA L F LIG HT
ENVELOPE. CONSEQUENTLY, PARTS
THAT HAVE BEEN USED ON MILITARY
HELICOPTERS SHOULD NOT BE USED
ON COMMERCIAL HELICOPTERS.
CAUTION
AIRWORTHINESS LIFE OF SOME KIT
COMPONENTS MAY NOT BE COVERED
I N T H I S S C H E D U L E . R E F E R T O
A P P L I C A B L E I N S TA L L A T I O N
INSTRUCTION (II) OR MAINTENANCE
MANUAL SUPPLEMENT (MMS) FOR KIT
COMPONENTS’ SCHEDULE.
NOTE
The airworthiness life given or the failure to give an airworthiness life to a component does not constitute a warranty of any kind.
The only warranty applicable to the helicopter or any component is the warranty included in the Purchase
Agreement for the helicopter or the component.
The airworthiness lives given to the components and assemblies are determined by experience, tests and the judgment of Bell Helicopter engineers. The
29 SEP 2008 Rev. 25
4-00-00
Page 5
BHT-407-MM-1 TC APPROVED
airworthiness lives and inspection intervals cannot be changed without the approval of the Minister of
Transport Canada.
that the parts and materials are disposed of in a manner that does not allow them to be returned for service. Refer to FAA Advisory Circular 21-38 for guidance on the disposal of unsalvageable helicopter parts and materials.
Prior to disposing of unsalvageable helicopter parts and materials, caution should be exercised to ensure
COMPONENT
Table 4-1. Airworthiness Limitations Schedule
PART NUMBER
1
RETIREMENT
YEARS/HOURS/RIN
2
Yoke
Grip
Grip
Pitch Horn
Upper Plate
Lower Plate
Driving Ring Set
Blade Bolt, Expandable
Lower Cone Seat
Lower Cone Seat
MAIN ROTOR HUB AND BLADES
407-010-101-101/-109
406-010-108-119/-121
406-010-108-125/-131
407-010-103-101/-107/-113
406-010-115-119/-127
406-010-117-115/-125
406-010-126-107/-113
406-310-103-101
407-010-107-101
407-010-107-103/-105
MAIN ROTOR CONTROLS
Pitch Link Tube
Clevis
Bearing
Rod End Assembly
Drive Link
Drive Lever
Gimbal Ring
406-010-413-119/-133/-139
406-010-416-101
406-310-405-101/-103
406-310-404-101
406-010-426-101
406-010-425-107
406-010-427-109
Drive Hub Set
Swashplate Outer Ring
Swashplate Inner Ring
Anti-Drive Link
406-010-428-109
406-010-411-117
406-010-410-121
406-010-432-101
Anti-Drive Lever
Collective Lever
406-010-431-109
406-010-408-101
Collective Idler Link 406-010-407-101
Cyclic Longitudinal Bellcrank 407-001-526-101/-105
Collective Bellcrank
Cyclic Lateral Bellcrank
Cyclic Lateral Bellcrank
407-001-524-101/-105
407-001-527-101
407-001-528-101
8
3
1250 hrs
5000 hrs
5000 hrs
2500 hrs
2500 hrs
48000 RIN
5000 hrs
1250 hrs
10000 hrs
5000 hrs
5000 hrs
5000 hrs
5000 hrs
5000 hrs
5000 hrs
5000 hrs
5000 hrs
5000 hrs
5000 hrs
5000 hrs
5000 hrs
5000 hrs
5000 hrs
5000 hrs
5000 hrs
5000 hrs
5000 hrs
4-00-00
Page 6 Rev. 25 29 SEP 2008
TC APPROVED BHT-407-MM-1
COMPONENT
Servo Actuator Support
Support Assembly
Sleeve Assembly
Bellcrank Support
Transmission Top Case
Blade
Blade
Yoke
Main Rotor Mast
Input Drive Shaft
Input Drive Shaft
Input Drive Shaft
Input Drive Shaft
L/H Pylon Side Beam
L/H Pylon Side Beam
R/H Pylon Side Beam
R/H Pylon Side Beam
Corner Mount
Corner Mount
Pylon Restraint Spring
Tailboom Assembly
Fwd Crosstube Assembly
(Std. Gear)
Fwd Crosstube Assembly
(Std. Gear)
Aft Crosstube Assembly
(Std. Gear)
Aft Crosstube Assembly
(Std. Gear)
Table 4-1. Airworthiness Limitations Schedule (Cont)
PART NUMBER
1
RETIREMENT
YEARS/HOURS/RIN
2
MAIN ROTOR CONTROLS (CONT)
407-001-500-101/-105
406-010-406-119
406-010-409-107/-113
407-001-511-101
TRANSMISSION
406-040-052-105/-109
TAIL ROTOR
407-016-001-101
406-016-100-119
406-012-102-101/-109
8
DRIVE SYSTEM
407-040-038-101/-105
206-340-300-103
206-340-300-105
206-340-300-107
407-340-310-101
PYLON SUPPORT
407-010-201-101
407-010-201-105
407-010-203-101
407-010-203-105
406-010-217-107
407-310-203-101
407-010-206-103
TAILBOOM
407-030-801-107,
407-530-014-101/-103
LANDING GEAR
407-050-101-101/-103
3
7
10
9
15
3 6
3
11
3
6
3
11
8
12
4
10000 hrs
5000 hrs
5000 hrs
5000 hrs
900 hrs
5000 hrs
5000 hrs
5000 hrs or 18000 RIN
5000 hrs
5000 hrs
1000 hrs or 5500 RIN
5000 hrs or 17000 RIN
1000 hrs or 5500 RIN
5000 hrs or 17000 RIN
5000 RIN
5000 hrs
5000 hrs
5000 RIN
407-722-101
407-050-102-101/-103
407-723-104
4
14
4
4
14
5000 RIN
5000 RIN
5000 RIN
29 SEP 2008 Rev. 25
4-00-00
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BHT-407-MM-1 TC APPROVED
COMPONENT
Fwd Crosstube Assembly
(High Gear/Lightweight Float)
Fwd Crosstube Assembly
(High Gear/Lightweight Float)
Aft Crosstube Assembly
(High Gear/Lightweight Float)
Aft Crosstube Assembly
(High Gear/Lightweight Float)
Aft Crosstube Assembly
(High Gear/Lightweight Float)
Table 4-1. Airworthiness Limitations Schedule (Cont)
PART NUMBER
1
RETIREMENT
YEARS/HOURS/RIN
2
LANDING GEAR (CONT)
407-050-201-101/-103
4
5000 RIN
407-724-101
407-050-202-101/-103
407-725-104
407-704-007-119
4
14
4
4
14
5
4
5000 RIN
5000 RIN
5000 RIN
2500 Landings
5000 RIN
KITS
Reservoir 1271762
13
15 Years
NOTES:
1
Airworthiness limitation for the part number listed applies to all successive dash numbers for that component unless otherwise specified.
2
RIN: Retirement Index Number. Components sensitive to operational events, such as torque events or run-on landings, are assigned a maximum RIN number. This number is based on the fatigue damage that results from normal helicopter lifts and take-offs (torque events) or from run-on landings, whichever applies.
New components will begin with an accumulated RIN of zero, which will increase as the helicopter is subjected to torque events or run-on landings. The operator must record the number of torque events or run-on landings, and increase the accumulated RIN as directed. When a component reaches the maximum RIN or retirement flight hours, whichever occurs first, the component must be retired from service.
3
For every one (1) torque event, you add one (1) RIN to the previous total. A torque event occurs for every takeoff (one takeoff plus the subsequent landing equals one RIN) and every load lift.
A load lift (internal or external) may be defined as a sling load, a rescue hoist load or any load that is added to the helicopter while airborne. For example: if an operator performs one (1) takeoff, picks up and drops ten (10) sling loads, and then lands, he must record eleven (11) torque events.
4
For every one (1) run-on landing, you add one (1) RIN to the previous total. A run-on landing is defined as one where there is forward ground travel of the helicopter greater than three (3) feet with the weight on the skids.
Refer to ASB 407-03-59.
5
This crosstube is modified from crosstube assembly P/N 407-050-202-101 per ASB 407-02-50. Depending on status of part determined per Table 1 of ASB 407-02-50, crosstube may be subject to retirement upon reaching 2500 landings.
6
When the helicopter is used for pilot training operations that include repeated autorotation landings, record
0.4 flight hour for each practice autorotation landing (including hover throttle chops). Autorotation approaches with power recovery to hover do not have to be counted as autorotation landings.
7
The input drive shaft 206-340-300-103 must be removed from service not later than September 30, 1998.
Refer to ASB 407-98-19 and Transport Canada Aviation Airworthiness Directive CF-98-25 Bell.
4-00-00
Page 8 Rev. 25 29 SEP 2008
TC APPROVED BHT-407-MM-1
COMPONENT
Table 4-1. Airworthiness Limitations Schedule (Cont)
PART NUMBER
1
RETIREMENT
YEARS/HOURS/RIN
2
NOTES (CONT):
8
These components were identified as life-limited items in Revision 6 of the airworthiness limitation schedule.
They are now “on condition” items. They will be removed from Table 4-1 at a subsequent revision if their status stays unchanged.
9
Input drive shaft 206-340-300-107 must be overhauled every 1,250 hours of operation. The overhaul is to be done by Kamatics Corp. only.
10
The input drive shaft 206-340-300-105 must be removed from service upon reaching 1250 hours in service.
Refer to ASB 407-01-45 for details.
11
Pylon side beams 407-010-201-105 and 407-010-203-105 are to be subjected to an airworthiness inspection.
See Table 4-2 for details.
12
Tailboom assemblies 407-030-801-107, 407-530-014-101 and 407-530-014-103 are also subjected to an airworthiness inspection. Refer to Table 4-2 and ASB 407-07-80 for details. Tailboom assemblies
407-030-801-201, -203 and -205 are only subjected to an airworthiness inspection. See Table 4-2 for details.
13
Reservoir per DOT-3HT and DOT exemption letter DOT-E-8162. This reservoir is part of reservoir assembly
P/N 407-073-848-101.
14
Aeronautical Accessories Incorporated (AAI) crosstube assemblies listed are BHT-approved production and spare alternates.
15
Input drive shaft 407-340-310-101 must be overhauled every 2,500 hours of operation. The overhaul is to be done by Kamatics Corp. only.
COMPONENT
Pylon Side Beam
Tail Rotor Blade
Yoke Assembly
Tailboom Assembly
Tailboom Assembly
Horizontal Stabilizer
Table 4-2. Inspection Limitations Schedule
PART NUMBER
406-040-340-101
1
407-010-201-105, 407-010-203-105
Bearing, Tail Rotor Hanger 406-040-339-ALL
Bearing, Tail Rotor Hanger 407-340-339-101/-103
Swashplate Bearing 406-310-402-101
407-016-001-101
407-010-101-101
407-030-801-201/-203/-205
407-030-801-107, 407-530-014-101/
-103
407-023-801-109
8 11
9 11
7
4
3
6
12
5
INSPECTION INTERVAL
(HRS OPERATING TIME)
2
25 hours
300 hours
25 hours
50/150 hours
300 hours
100 hours
300 hours
Daily and 100 hours
13 600 hours or annual
29 SEP 2008 Rev. 25
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Page 9
BHT-407-MM-1 TC APPROVED
COMPONENT
Table 4-2. Inspection Limitations Schedule (Cont)
PART NUMBER
1
INSPECTION INTERVAL
(HRS OPERATING TIME)
2
NOTES:
1
Inspection limitation for the part number listed applies to all successive dash numbers for that component unless otherwise specified
.
2
Refer to Chapter 5 for inspection requirements.
3
Inspect couplings 406-040-340-101 every 25 hours of operation. Refer to ASB 407-97-13, Transport
Canada Aviation Airworthiness Directive CF-97-20R1 and Federal Aviation Authority Airworthiness Directive
97-24-17. Refer to Figure 4-1 for inspection details.
4
Refer to Figure 4-2 for inspection details.
5 The swashplate bearing 406-310-402-101 must be removed from service not later than December 31,
1998 and replaced by bearing 406-310-402-103. Refer to ASB 407-97-11 and Transport Canada Aviation
Airworthiness Directive CF-97-22 Bell.
6
Refer to Figure 4-3 for inspection details.
7
Refer to Figure 4-4 for inspection details.
8 You must do an inspection of tailbooms 407-030-801-201/-203, and -205 every 300 hours of operation.
Additional inspection requirements are applicable for tailbooms that have accumulated 6900 hours in service and
8600 hours in service. Refer to Figure 4-5 for inspection details.
9
You must do an inspection of tailbooms 407-030-801-107 and 407-530-014-101/-103 daily and every 100 hours of operation. Refer to Figure 4-6 and Figure 4-7 for inspection details.
10
You must do a 10X inspection of pylon side beams 407-010-201-105 and 407-010-203-105 that have
upon reaching 2500 hours in service. Refer to
Figure 4-8 for inspection details.
11 Contact Bell Helicopter Textron Product Support Engineering for changes or modifications to the structure in area where a mandatory airworthiness inspection is specified.
12
The oil cooler blower and tail rotor segmented driveshaft bearings 407-340-339-101 and -103 must be
13 Selected serial numbers of horizontal stabilizer 407-023-801-109 must be removed from service no later than 30 September 2008. Refer to BHT ASB 407-06-72 for listing of serial numbers affected. Applicable
horizontal stabilizers must be inspected every 600 hours or annually until replaced. Refer to
Figure 4-9 for inspection details.
4-00-00
Page 10 Rev. 25 29 SEP 2008
TC APPROVED BHT-407-MM-1
Figure 4-1. Disc Pack Coupling (P/N 406-040-340-101) — Inspection
29 SEP 2008 Rev. 25
4-00-00
Page 11
BHT-407-MM-1 TC APPROVED
Figure 4-2. Main Rotor Yoke (P/N 407-010-101-101) — Inspection
4-00-00
Page 12 Rev. 25 29 SEP 2008
TC APPROVED BHT-407-MM-1
Figure 4-3. Tail Rotor Hanger Bearing (P/N 406-040-339-ALL) — Inspection (Sheet 1 of 2)
29 SEP 2008 Rev. 25
4-00-00
Page 13
BHT-407-MM-1 TC APPROVED
4-00-00
Page 14
Figure 4-3. Tail Rotor Hanger Bearing (P/N 406-040-339-ALL) — Inspection (Sheet 2 of 2)
Rev. 25 29 SEP 2008
TC APPROVED BHT-407-MM-1
4
STA
14.50
12.0 IN.
(304.8 mm)
NO REPAIR
TO SKIN PERMITTED
NOTES
1.
This inspection applies to the tail rotor blade 407-016-001 only.
2.
Do this inspection every 300 hours of component operation.
3.
Remove the tail rotor blade.
4 Inspect the blade roots inboard of blade station 14.50 for signs of damage.
Examine the inside of the blade root pocket (where the tail rotor yoke is inserted) for signs of crazing on the surface of the blade skins. If crazing is detected, remove the blade from service.
Crazing will look like a series of minute cracks on the surface of the skin and will make the surface of the skin look cloudy.
5.
Using a coin or heavy washer, lightly tap the outer surface of the blade in the blade root area to detect delamination in the skin. An area of delamination will sound hollow when gently tapped. If delamination is detected, remove the blade from service.
6.
If no crazing or delamination is detected, return the blade to service.
7.
Install the tail rotor blade.
8.
Balance the tail rotor (Chapter 18).
407MM_04_0017
Figure 4-4. Tail Rotor Blade (P/N 407-016-001-101) — Inspection
29 SEP 2008 Rev. 25
4-00-00
Page 15
BHT-407-MM-1 TC APPROVED
BS
39.00
TAILBOOM IDENTIFICATION PLATE
AREA
D
BS
140.428
BS
157.00
BS
107.397
BS
72.370
AREA
H
AREA
I
AREA
A
AREA
C
AREA
B
300-HOUR VISUAL INSPECTION
CRITICAL INSPECTION AREAS (SEE DETAILS)
407_MM_04_0033a
Figure 4-5. Tailboom Assembly (P/N 407-030-801-201, 407-030-801-203, and 407-030-801-205) — 300 Hour
Inspection (Sheet 1 of 6)
4-00-00
Page 16 Rev. 25 29 SEP 2008
TC APPROVED BHT-407-MM-1
Figure 4-5. Tailboom Assembly (P/N 407-030-801-201, 407-030-801-203, and 407-030-801-205) — 300 Hour
Inspection (Sheet 2 of 6)
29 SEP 2008 Rev. 25
4-00-00
Page 17
BHT-407-MM-1 TC APPROVED
Figure 4-5. Tailboom Assembly (P/N 407-030-801-201, 407-030-801-203, and 407-030-801-205) — 300 Hour
Inspection (Sheet 3 of 6)
4-00-00
Page 18 Rev. 25 29 SEP 2008
TC APPROVED BHT-407-MM-1
4.25 IN.
(108.0 mm)
1
2.0 IN.
(50.8 mm)
AREA
C
3 6 7
2 4
3
7 6 3
5.5 IN.
(139.7 mm)
AREA
D
2.0 IN.
(50.8 mm)
BS
120.75
BS
98.89
AREA
C
AND AREA
D
(HORIZONTAL STABILIZER NOT SHOWN FOR CLARITY)
GENERAL VISUAL INSPECTION (300 HOURS)
CRITICAL INSPECTION AREA
407_MM_04_0033d
Figure 4-5. Tailboom Assembly (P/N 407-030-801-201, 407-030-801-203, and 407-030-801-205) — 300 Hour
Inspection (Sheet 4 of 6)
29 SEP 2008 Rev. 25
4-00-00
Page 19
BHT-407-MM-1 TC APPROVED
Figure 4-5. Tailboom Assembly (P/N 407-030-801-201, 407-030-801-203, and 407-030-801-205) — 300 Hour
Inspection (Sheet 5 of 6)
4-00-00
Page 20 Rev. 25 29 SEP 2008
TC APPROVED BHT-407-MM-1
1.
2.
3.
Upper tailboom skin
Horizontal stabilizer
Left upper stabilizer attachment support
1.
2.
3
4
NOTES
AT EVERY 300-HOUR INTERVAL:
Inspect the complete tailboom assembly for general condition.
On the left side of the tailboom only, do a detailed inspection of all areas shown for cracks in the tailboom skin and loose rivets. Do not remove the paint or primer. If a loose rivet is found, remove rivet and inspect hole for cracks using a 10X magnifying glass. If no cracks are found, install correct diameter rivet. Do not exceed maximum diameter prescribed.
Use a 10X magnifying glass to inspect for cracks in tailboom skin and around fastener heads as indicated in area shown in Details E, F, G, Areas C, D, H and I.
Do not remove horizontal stabilizer.
5
FOR TAILBOOMS IN SERVICE FOR 6900 FLIGHT-HOURS OR MORE, OR WITH TOTAL TIME UNKNOWN:
Inspect Area H and Area I at every 150-hour interval using 10X magnifying glass inspection method, or at every 500-hour interval using Eddy current inspection method.
6
7
FOR TAILBOOMS IN SERVICE FOR 8600 FLIGHT-HOURS OR MORE, OR WITH TOTAL TIME UNKNOWN:
Visually check Area C and Area D daily or at every 50-hour interval using 10X magnifying glass inspection method. Do not exceed a distance of 12 inches (30.48 cm) from the tailboom surface when conducting the daily visual inspection.
Pay close attention to skin just above edge of upper support.
8.
9.
Some tailboom components not shown for clarity.
If a crack is found on the tailboom skin, replace the tailboom before the next flight and contact Product
Support Engineering using the following information:
Bell Helicopter Textron
Product Support Engineering
Helicopters
Tel: 1-800-243-6407 (Continental USA and Canada)
Tel: 1-800-363-8023 (Continental USA)
Tel: 1-800-361-9305 (Within Canada)
Tel: 1-450-437-2682 (All other areas - call collect)
450-433-0272 [email protected]
10.
Contact Bell Helicopter Textron Product Support Engineering for changes or modifications to the structure in areas where a mandatory airworthiness inspection is specified.
407_MM_04_0033f
Figure 4-5. Tailboom Assembly (P/N 407-030-801-201, 407-030-801-203, and 407-030-801-205) — 300 Hour
Inspection (Sheet 6 of 6)
29 SEP 2008 Rev. 25
4-00-00
Page 21
BHT-407-MM-1 TC APPROVED
407MM_04_0024
Figure 4-6. Tailboom Assembly (P/N 407-530-014-101, 407-530-014-103, or 407-030-801-107) — Daily
Inspection (Sheet 1 of 2)
4-00-00
Page 22 Rev. 25 29 SEP 2008
TC APPROVED BHT-407-MM-1
DAILY INSPECTION
NOTES
1. Before the first flight of the day, do a check of the left side of the tailboom assembly in the areas where
paint was removed for cracks, as follows: a) If required, use a clean cloth moistened with cleaning compound (C-318) prepared in accordance with
manufacturer’s recommendations to remove any exhaust residues or dirt, from both areas that require
checking, as shown.
b) Make sure that the tailboom skin is clean and that adequate lighting exists. Visually inspect the two
areas where original paint and primer was removed and protected with clear paint coating. Look for
cracks from a distance not exceeding 12 inches (30.48 cm).
c) Pay close attention near the edge of the stabilizer upper support.
d) If a crack is found on the tailboom skin, replace the tailboom before the next flight and contact
Product Support Engineering at the following numbers:
Bell Helicopter Textron
Product Support Engineering
Tel: 1-800-243-6407 (Continental USA and Canada)
Tel: 1-800-363-8023 (Continental USA)
Tel: 1-800-361-9305 (Within Canada)
Tel: 1-450-437-2682 (All other areas - call collect)
Fax: 450-433-0272 Internet: [email protected]
e) If no crack is found, the intent of the daily inspection is complete.
2. Make an entry in the helicopter records to show that the daily inspection is completed.
3. Comply with subsequent 100 hour inspection requirements described in Figure 4-7 when time is reached.
407MM_04_0025
Figure 4-6. Tailboom Assembly (P/N 407-530-014-101, 407-530-014-103, or 407-030-801-107) — Daily
Inspection (Sheet 2 of 2)
29 SEP 2008 Rev. 25
4-00-00
Page 23
BHT-407-MM-1 TC APPROVED
BS
39.00
BS
72.370
BS
107.397
BS
140.428
AREA
A
AREA
D
AREA
C
AREA
B
BS
90.00
100 HOUR VISUAL INSPECTION
CRITICAL INSPECTION AREAS (SEE DETAILS)
407_MM_04_0026
Figure 4-7. Tailboom Assembly (P/N 407-530-014-101, 407-530-014-103, and 407-030-801-107) —
100 Hour Inspection (Sheet 1 of 7)
4-00-00
Page 24 Rev. 25 29 SEP 2008
TC APPROVED BHT-407-MM-1
BS
31.920
SEE
DETAIL
I
BS
74.370
BS
90.00
BL
0.00
SEE
DETAIL
E
BS
39.000
BS
45.515
SEE
DETAIL
F
BS
52.600
BS
58.995
AREA
B
AREA
B
SEE
DETAIL
G
SEE
DETAIL
G
BS
72.475
BS
77.000
BS
85.955
SEE
DETAIL
H
LBL
7.50
FWD
BL
0.00
0.50 IN.
(12.7 mm)
0.50 IN.
(12.7 mm)
LBL
7.50
9 FASTENERS
(LEFT SIDE ONLY)
3
BS
45.515
DETAIL
E
GENERAL VISUAL INSPECTION (100 HOURS)
100 HOUR INSPECTION (10X MAGNIFYING GLASS)
407_MM_04_0027
Figure 4-7. Tailboom Assembly (P/N 407-530-014-101, 407-530-014-103, and 407-030-801-107) —
100 Hour Inspection (Sheet 2 of 7)
29 SEP 2008 Rev. 25
4-00-00
Page 25
BHT-407-MM-1 TC APPROVED
Figure 4-7. Tailboom Assembly (P/N 407-530-014-101, 407-530-014-103, and 407-030-801-107) —
100 Hour Inspection (Sheet 3 of 7)
4-00-00
Page 26 Rev. 25 29 SEP 2008
TC APPROVED BHT-407-MM-1
Figure 4-7. Tailboom Assembly (P/N 407-530-014-101, 407-530-014-103, and 407-030-801-107) —
100 Hour Inspection (Sheet 4 of 7)
29 SEP 2008 Rev. 25
4-00-00
Page 27
BHT-407-MM-1 TC APPROVED
4.0 IN.
(101.6 mm)
HORIZONTAL
STABILIZER
(REF)
4
2.0 IN.
(50.8 mm)
FWD
BS
97.89
BS
98.89
8 RIVETS
(LEFT SIDE ONLY)
BS
107.397
SEE
VIEW
J
3
BS
120.75
BS
121.75
4.0 IN.
(101.6 mm)
AREA
BS
107.397
C
2.0 IN.
(50.8 mm)
BL
0.00
8 RIVETS
(LEFT SIDE ONLY)
3
LBL
7.00
3.00 IN.
(76.2 mm)
VIEW
J
GENERAL VISUAL INSPECTION (100 HOURS)
100 HOUR INSPECTION (10X MAGNIFYING GLASS)
407_MM_04_0030
Figure 4-7. Tailboom Assembly (P/N 407-530-014-101, 407-530-014-103, and 407-030-801-107) —
100 Hour Inspection (Sheet 5 of 7)
4-00-00
Page 28 Rev. 25 29 SEP 2008
TC APPROVED
BS
140.428
SEE
VIEW
D
LEFT SIDE ONLY
BHT-407-MM-1
FWD
AREA
D
BS
140.428
BL
0.00
3
6 RIVETS
(LEFT SIDE ONLY)
LBL
5.00
6.00 IN.
(152.4 mm)
VIEW
D
GENERAL VISUAL INSPECTION (100 HOURS)
100 HOUR INSPECTION (10X MAGNIFYING GLASS)
407_MM_04_0031
Figure 4-7. Tailboom Assembly (P/N 407-530-014-101, 407-530-014-103, and 407-030-801-107) —
100 Hour Inspection (Sheet 6 of 7)
29 SEP 2008 Rev. 25
4-00-00
Page 29
BHT-407-MM-1 TC APPROVED
100 HOURS OF OPERATION
1.
NOTES
Inspect the complete tailboom assembly for general condition.
2.
On the left side of the tailboom only, do a detailed inspection of all areas shown, for cracks in the tailboom skin and loose rivets. Do not remove the paint or primer. If a loose rivet is found, remove rivet and inspect hole for cracks using a 10X magnifying glass. If no cracks are found, install correct diameter rivet. Do not exceed maximum diameter prescribed.
3 Use a 10X magnifying glass to inspect for cracks in tailboom skin and around fastener heads as indicated in area shown in Detail E, F, G, H, I, View J, View D and Area C .
4 Do not remove horizontal stabilizer.
5.
Some tailboom components not shown for clarity.
6.
If a crack is found on the tailboom skin, replace the tailboom before the next flight and contact
Product Support Engineering at the following numbers:
Bell Helicopter Textron
Product Support Engineering
Light Helicopters
Tel: 1-800-243-6407 (Continental USA and Canada)
Tel: 1-800-363-8023 (Continental USA)
Tel: 1-800-361-9305 (Within Canada)
Tel: 1-450-437-2682 (All other areas - call collect)
Fax: 450-433-0272 Internet: [email protected]
7.
Contact Bell Helicopter Textron Product Support Engineering for changes or modifications to the structure in areas where a mandatory airworthiness inspection is specified.
407_MM_04_0032
Figure 4-7. Tailboom Assembly (P/N 407-530-014-101, 407-530-014-103, and 407-030-801-107) —
100 Hour Inspection (Sheet 7 of 7)
4-00-00
Page 30 Rev. 25 29 SEP 2008
TC APPROVED BHT-407-MM-1
Figure 4-8. Inspection of Pylon Side Beams (P/N 407-010-201-105/407-010-203-105)
(Sheet 1 of 3)
29 SEP 2008 Rev. 25
4-00-00
Page 31
BHT-407-MM-1 TC APPROVED
Figure 4-8. Inspection of Pylon Side Beams (P/N 407-010-201-105/407-010-203-105)
(Sheet 2 of 3)
4-00-00
Page 32 Rev. 25 29 SEP 2008
TC APPROVED BHT-407-MM-1
1 2
3.
NOTES
Inspect the upper outboard flange and flange/web area for cracks. Inspect the area between 1.25 inches (31.7 mm) and 2.0 inches (50.8 mm) above the mounting pad.
If a crack is found in a pylon side beam, replace before the next flight and contact Product Support
Engineering at the following numbers:
Bell Helicopter Textron
Product Support Engineering
Light Helicopters
Tel: 1-800-243-6407 (Continental USA and Canada)
Tel: 1-800-363-8023 (Continental USA)
Tel: 1-800-361-9305 (Within Canada)
Tel: 1-450-437-2862 (All other areas - call collect)
Fax: 450-433-0272 Internet: [email protected]
Figure 4-8. Inspection of Pylon Side Beams (P/N 407-010-201-105/407-010-203-105)
(Sheet 3 of 3)
407MM 04 0006
29 SEP 2008 Rev. 25
4-00-00
Page 33
BHT-407-MM-1 TC APPROVED
1
12
9
4
8
6
19 2
20
21
22
17
3
SEE DETAIL
A
14
15
18
1
16
DETAIL
A
SEE DETAIL
B
7
17
24
FWD
13
C
11
10
2
C
SECTION
C-C
23
5
12
DETAIL
B
INSPECTION AREA
Inspect stabilizer on upper and lower surfaces
1.5 IN.
(38.1 mm)
9
6.
7.
8.
9.
10.
11.
12.
1.
2.
3.
4.
5.
1
Finlet
Slat assembly
Rivet
Upper attachment support
Lower attachment support
Tailboom
Horizontal stabilizer
Rivet
Screw
Support bracket
Screw
Washer
17.
18.
19.
20.
21.
22.
23.
24.
13.
14.
15.
16.
Washer
Washer
Screw
Upper and lower doubler
Casting supports
Flap
Spacer (left side only)
Screw
Washer
Nut
Inserts
Identification tag
2
NOTES
EFFECTIVITY
Aircraft Serial Numbers 53057 and subsequent.
Aircraft Serial Numbers 53000 to 53056 Post TB 407-96-2.
SEALANT MIL-PRF-81733 (C-392)
20 TO 25 IN-LBS
(2.26 TO 2.83 Nm)
2 Tailboom spacer effectivity: P/N 407-030-801-201, 407-030-801-203, 407-030-801-205.
407MM_04_0022
Figure 4-9. Horizontal Stabilizer (P/N 407-023-801-109) — Recurring 600 Hour/Annual Inspection
(Sheet 1 of 2)
4-00-00
Page 34 Rev. 25 29 SEP 2008
TC APPROVED BHT-407-MM-1
PROCEDURE
1.
Inspect the horizontal stabilizer (7) as follows:
NOTE
Horizontal stabilizer does not require to be removed from tailboom.
CAUTION
PROTECTION UNDERNEATH STABILIZER IS REQUIRED TO NOT DAMAGE
STABILIZER OR TAILBOOM WHEN STABILIZER IS DETACHED FROM UPPER
SUPPORTS. SLAT MAY REMAIN INSTALLED HOWEVER SIDE MOVEMENT WILL BE
LIMITED. ENSURE NO DAMAGE TO SLATS AND/OR TAILBOOM SURFACES.
NOTE
Do not damage liquid shim compound between upper and lower supports (4 and 5) and stabilizer (7). If liquid shim is damaged during removal of the parts, the shimming procedure shall be repeated as described in BHT-407-MM, Chapter 53.
2.
Remove the screws (9) and the washers (12) from the L/H and R/H lower supports (5). Remove the nuts (22), washers (21) and screws (20).
3.
Remove L/H and R/H lower supports (5).
4.
Protect tailboom (6) underneath stabilizer (7) with padding or equivalent.
5.
Hold the stabilizer (7). Remove the screws (9) and the washers (12) that attach the L/H and R/H upper supports (4) to stabilizer (7).
6.
Carefully lower stabilizer (7) and move to left side.
7.
Refer to Figure Detail B. Clean and prepare surface of the stabilizer for inspection using cheesecloth (C-486) moistened with acetone (C-316).
8.
Inspect stabilizer in area of the inserts (23), on upper and lower surfaces for cracks and deformation using a
10X power magnifying glass. Inspect the area around each insert and between inserts for debonding using tap test method.
9.
If any damage is found, the stabilizer is not repairable. Replace the stabilizer.
10.
If no damage is found, continue with the inspection.
11.
Carefully move stabilizer (7) to the right side.
12.
Repeat step 7 to 9 for the R/H side of stabilizer in area of inserts (23).
13.
If any damage is found, the stabilizer is not repairable. Replace the stabilizer.
14.
If no damage is found, reinstall horizontal stabilizer in accordance with BHT-407-MM, Chapter 53.
15.
Make an entry in the helicopter records to indicate that inspection has been accomplished and that a recurring 600 hour or annual inspection is required until stabilizer is replaced.
407MM_04_0023
Figure 4-9. Horizontal Stabilizer (P/N 407-023-801-109) — Recurring 600 Hour/Annual Inspection
(Sheet 2 of 2)
29 SEP 2008 Rev. 25
4-00-00
Page 35/36
BHT-407-MM-2
MAINTENANCE MANUAL
VOLUME 2
HANDLING AND SERVICING
NOTICE
The instructions set forth in this manual, as supplemented or modified by Alert Service Bulletins (ASB) or other directions issued by Bell Helicopter Textron Inc. and Airworthiness Directives (AD) issued by the applicable regulatory agencies, shall be strictly followed.
COPYRIGHT NOTICE
COPYRIGHT 2008
BELL ® HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON
CANADA LTD.
ALL RIGHTS RESERVED
22 FEBRUARY 1996
REVISION 25 — 29 SEPTEMBER 2008
BHT-407-MM-2
PROPRIETARY RIGHTS NOTICE
These data are proprietary to Bell Helicopter Textron Inc. Disclosure, reproduction, or use of these data for any purpose other than helicopter operation or maintenance is forbidden without prior written authorization from
Bell Helicopter Textron Inc.
PN
Additional copies of this publication may be obtained by contacting:
Commercial Publication Distribution Center
Bell Helicopter Textron Inc.
P. O. Box 482
Fort Worth, Texas 76101-0482
Rev. 22 28 NOV 2005
BHT-407-MM-2
Figure 7-1. Lifting the Complete Helicopter (Sheet 2 of 2)
29 SEP 2008 Rev. 25
7-00-00
Page 5
BHT-407-MM-2
7-3.
LIFTING THE COMPLETE HELICOPTER
WITH THE MAST NUT
SPECIAL TOOLS REQUIRED
NUMBER
T102137-101
206-070-469-001
NOMENCLATURE
Lifting Eye Clevis
Tail Rotor Strap
WARNING
NOTE
If you lift the helicopter higher than you can hold it, attach a rope to the tail skid. Make a person hold the rope.
7.
Put a person at the tail skid. Tell the person to keep the helicopter in a level position when you lift the helicopter.
8.
Operate the hoist (1). Apply a constant force and slowly lift the helicopter from the ground.
7-4.
LIFTING THE HELICOPTER WITH THE
MAIN ROTOR AND THE MAST ASSEMBLY
REMOVED
DO NOT GO INTO OR CLIMB ONTO THE
HELICOPTER WHILE IT IS SUPPORTED
ON THE JACKS OR BEING RAISED. IF
Y O U D O N O T O B E Y, I N J U RY T O
PERSONNEL CAN OCCUR.
SPECIAL TOOLS REQUIRED
NUMBER
T103314-101
NOMENCLATURE
Lift Plate
CAUTION
WARNING
ALL OF THE LIFT EQUIPMENT MUST BE
RATED FOR A MINIMUM CAPACITY OF
7500 POUNDS (3400 KG). IF YOU USE
EQUIPMENT THAT IS NOT RATED FOR
THIS CA PA CITY, DAM AGE TO THE
HELICOPTER CAN OCCUR.
1.
Make the area around the helicopter safe. Put ropes around the area. Put warning signs up that have the message, "THIS HELICOPTER IS ON A HOIST".
2.
Remove the main rotor fairing and the Frahm damper (if installed) (Chapter 62).
3.
If applicable, use the tail rotor strap and safety the tail rotor (Chapter 10).
4.
Use the pin (7, Figure 7-1 ) to attach the lifting eye clevis (2) to the mast nut (4).
5.
Use the pin (3) and safety the pin (7).
6.
Attach the hoist (1) to the lifting tool (2). Operate the hoist (1) until the cable is tight.
7-00-00
Page 6 Rev. 22 28 NOV 2005
DO NOT GO INTO OR CLIMB ONTO THE
HELICOPTER WHILE IT IS SUPPORTED
ON THE JACKS OR BEING RAISED. IF
Y O U D O N O T O B E Y, I N J U RY T O
PERSONNEL CAN OCCUR.
CAUTION
ALL OF THE LIFT EQUIPMENT MUST BE
RATED FOR A MINIMUM CAPACITY OF
7500 POUNDS (3400 KG). IF YOU USE
EQUIPMENT THAT IS NOT RATED FOR
THIS CA PA CITY, DA MA GE TO THE
HELICOPTER CAN OCCUR.
1.
Make the area around the helicopter safe. Put ropes around the area. Put warning signs up that have the message, "THIS HELICOPTER IS ON A HOIST".
2.
Remove the main rotor hub (5, Figure 7-1 )
(Chapter 62) and the mast assembly (6) (Chapter 63).
BHT-407-MM-5
MAINTENANCE MANUAL
VOLUME 5
AIRFRAME
NOTICE
The instructions set forth in this manual, as supplemented or modified by Alert Service Bulletins (ASB) or other directions issued by Bell Helicopter Textron Inc. and Airworthiness Directives (AD) issued by the applicable regulatory agencies, shall be strictly followed.
COPYRIGHT NOTICE
COPYRIGHT 2008
BELL ® HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON
CANADA LTD.
ALL RIGHTS RESERVED
22 FEBRUARY 1996
REVISION 25 — 29 SEPTEMBER 2008
BHT-407-MM-5
PROPRIETARY RIGHTS NOTICE
These data are proprietary to Bell Helicopter Textron Inc. Disclosure, reproduction, or use of these data for any purpose other than helicopter operation or maintenance is forbidden without prior written authorization from
Bell Helicopter Textron Inc.
PN
Additional copies of this publication may be obtained by contacting:
Commercial Publication Distribution Center
Bell Helicopter Textron Inc.
P. O. Box 482
Fort Worth, Texas 76101-0482
Rev. 24 2 OCT 2007
BHT-407-MM-5
1
1
SECTION
A-A
1
1
3
2
SECTION
B-B
2
2
3 SEE VIEW
D
1
A
A
6.0 IN. (152.4 mm)
LONG. APPLY TO
FILL DIMPLED ZONE
2
B
1
2
B
SEE VIEW
C
1
1
2
VIEW
C
3
2
1.
2.
3.
Lower window
Screw
Washer
1
SEALANT (C-328)
2
SEALANT (C-392)
2
TO FILL ALONG
LOWER WINDOW
VIEW
D
2
TO FILL ALONG
LOWER WINDOW
407MM_52_0026_c1
Figure 52-14. Lower Window — Maintenance Practices
29 SEP 2008 Rev. 25
52-00-00
Page 57
BHT-407-MM-5
b. Examine the lower windows (1) for abrasions, scratches, cracks, holes, and other defects that can decrease visibility. Repair the lower windows, as specified in Figure 52-12 .
c. You can temporarily repair cracks, holes, or other damage by stop drilling, with repair patches, or by following other standard practices for acrylic plastic.
d. Examine the lower windows (1, Figure 52-14 ) for loose screws (2).
e. Make sure the sealant (C-392) has no cracks, holes, or deterioration that permit water leakage.
f. To stop water leaks, apply sealant (C-392) to the area of the leak.
52-67.
Lower Windows — Installation
6.
Install 33 screws (2) and washers (3) to attach the lower window (1) to the nose skin, floor panel, and console structure.
7.
Seal the window edges with sealant (C-308).
8.
Clean the lower window (Chapter 12).
52-68.
SKYLIGHT — DESCRIPTION
The two skylights are made of light grey tinted acrylic plastic. The skylights are flush mounted on the forward cabin roof and part of the sidebody fairings. The skylights are installed with double-sided self-adhesive tape and four screws. The skylights are sealed for water tightness.
52-69.
Skylight — Removal
MATERIALS REQUIRED
Refer to BHT-ALL-SPM for specifications.
NUMBER NOMENCLATURE
C-305
C-328
C-392
1.
Clean the fiberglass area of the window (1,
Figure 52-14 ) with aliphatic naphtha (C-305).
2.
Apply a layer of sealant (C-328) to the faying surfaces of the lower window (1).
3.
Clean the excess sealant (C-328) from the window.
4.
Apply sealant (C-392) at the locations shown on
Figure 52-14 .
5.
Put the lower window on the helicopter. Start at the aft inboard side and continue in a counterclockwise direction as seen from inside the helicopter. Make sure that the holes in the window frame align with the holes in the fiberglass reinforcement.
52-00-00
Page 58
Aliphatic Naphtha
Sealant
Sealant
Rev. 25 29 SEP 2008
MATERIALS REQUIRED
Refer to BHT-ALL-SPM for specifications.
NUMBER NOMENCLATURE
C-305
C-392
Aliphatic Naphtha
Sealant
C-405
C-407
Lockwire
Abrasive Pad
1.
Remove the nuts (2, Figure 52-15 ), washers (3), spacers (4), and screws (5).
2.
Lubricate a wire 0.025 inch (0.635 mm) in diameter with a soap/detergent and water solution.
3.
Push the wire in between the window and frame.
4.
Attach a wooden dowel at each end of the wire.
Safety the wire with lockwire (C-405).
5.
Pull on the dowel handles to separate the window from the frame.
6.
Remove the skylight window (1).
7.
Remove the adhesive tape (6) from the supports and skylight window (1).
BHT-407-MM-5
VERTICAL FIN
53-53. VERTICAL FIN
The vertical fin (3, Figure 53-26 ) is an aerodynamic surface that gives stability to the helicopter while it is in flight. The vertical fin is installed with the leading edge at a 9° angle (with reference to the helicopter longitudinal axis) outboard. This helps unload the tail rotor while the helicopter is in forward flight. The vertical fin is made of an aluminum honeycomb core with aluminum outer skins. The forward and the aft edge caps are made of formed aluminum alloy. The anti-collision light (1) is installed on the top fairing (2) of the vertical fin. The rubber bumper and the tail skid
(9) are installed on the lower edge of the vertical fin.
The tail skid absorbs shocks in a tail low landing.
(Refer to Chapter 32.)
On helicopters S/N 53057 and subsequent, and on helicopters S/N 53000 through 53056 Post TB
407-96-2, a flap is installed along the trailing edge of the vertical fin.
53-54.
VERTICAL FIN — REMOVAL
1.
Remove the tail rotor gearbox fairing ( paragraph
53-83 ).
2.
Disconnect the wires of the anti-collision light (1,
Figure 53-26 ) at the quick disconnects or at the terminal board on the tail rotor gearbox support casting (Chapter 96).
3.
Put insulation on the wire ends and stow wires to prevent damage.
4.
Hold the vertical fin (3). Remove the bolts (5) and the washers (6) from the fin.
5.
Remove the vertical fin (3) from the tail rotor gearbox support casting fin support (10).
6.
Put the vertical fin (3) on a soft surface to prevent damage.
53-55.
VERTICAL FIN — INSPECTION AND
REPAIR
WARNING
CORROSION AT THE FAYING EDGES
OF THE VERTICAL FIN FITTINGS CAN
CAUSE THE BOLTS (5, FIGURE 53-26 )
TO LOOSEN. THIS CAN RESULT IN
DAMAGE TO THE FITTINGS AND THE
B O L T S A N D T H E L O S S O F T H E
VERTICAL FIN.
NOTE
For information on damage and repair procedures not contained in this chapter, please contact Product Support
Engineering for assistance.
1.
Examine the attachment area of the vertical fin (3,
Figure 53-26 ) for cracks, extended bolt holes, loose bolts, distortion, and corrosion.
NOTE
Step 2 applies if the vertical fin has been removed.
2.
Examine the bolts (5), the barrel nuts (13), and the nut retainers (12) for wear or thread damage.
3.
On helicopters S/N 53057 and subsequent, and on helicopters S/N 53000 through 53056 Post TB
407-96-2, examine the flap (4) for cracks, dents, damage, and loose rivets.
4.
Examine the vertical fin for scratches, cracks, dents, corrosion, and other damage. Inspect for security of attachment to tail rotor gearbox support casting (10). Refer to Figure 53-27 for negligible damage and repair limits.
5.
Examine the anti-collision light (1) for damage. If it is damaged, replace it (Chapter 96).
2 OCT 2007 Rev. 24
53-00-00
Page 93
BHT-407-MM-5
SEE DETAIL
A
1
2
3
4
1
3 (4 PLACES)
1.
2.
3.
4.
5.
6.
7.
8.
9.
10.
11.
12.
13.
Anticollision light
Fairing
Vertical fin
Flap
Bolt
Thin aluminium washer
Sleeve and plug
Former tail rotor gearbox fairing
Tail skid assembly
Tail rotor gearbox support casting and fin support
Tailboom
Nut retainer
Barrel nut
5
6
7
8
9
SEE DETAIL
B
DETAIL
A
CORROSION PREVENTIVE COMPOUND (C-104) 2
75 TO 95 IN-LBS
(8.5 TO 10.7 Nm)
1
NOTES
Helicopter S/N 53000 through 53056 Post TB 407-96-2 and helicopter S/N 53057 and subsequent.
2 Apply a coating of corrosion preventive compound (C-104) to all bolt shanks prior to installation. Do not apply corrosion preventive compound to bolt threads.
12
3 The faying surfaces between the vertical fin (3) and the tail rotor gearbox support casting and fin support (10) are to be coated with primer only.
13
3
DETAIL
B
10
(4 PLACES)
10
3
11
407_MM_53_0053_c1
Figure 53-26. Vertical Fin Assembly
53-00-00
Page 94 Rev. 25 29 SEP 2008
BHT-407-MM-6
MAINTENANCE MANUAL
VOLUME 6
MAIN ROTOR/MAIN ROTOR DRIVE SYSTEM
NOTICE
The instructions set forth in this manual, as supplemented or modified by Alert Service Bulletins (ASB) or other directions issued by Bell Helicopter Textron Inc. and Airworthiness Directives (AD) issued by the applicable regulatory agencies, shall be strictly followed.
COPYRIGHT NOTICE
COPYRIGHT 2008
BELL ® HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON
CANADA LTD.
ALL RIGHTS RESERVED
22 FEBRUARY 1996
REVISION 25 — 29 SEPTEMBER 2008
BHT-407-MM-6
PROPRIETARY RIGHTS NOTICE
These data are proprietary to Bell Helicopter Textron Inc. Disclosure, reproduction, or use of these data for any purpose other than helicopter operation or maintenance is forbidden without prior written authorization from
Bell Helicopter Textron Inc.
PN
Additional copies of this publication may be obtained by contacting:
Commercial Publication Distribution Center
Bell Helicopter Textron Inc.
P. O. Box 482
Fort Worth, Texas 76101-0482
Rev. 24 2 OCT 2007
BHT-407-MM-6
63-1
63-2
63-3
63-4
63-5
63-6
63-7
63-8
63-9
63-10
63-11
63-12
63-13
63-14
63-15
63-21
63-22
63-23
63-24
63-25
63-26
63-27
63-28
63-29
63-30
63-31
63-32
63-33
63-34
Paragraph
Number
63-16
63-17
63-18
63-19
63-20
CHAPTER 63 — MAIN ROTOR DRIVE SYSTEM
TABLE OF CONTENTS
Title
Chapter/Section
Number
MAIN ROTOR DRIVE SYSTEM
Main Rotor Drive System ...............................................................
Main Rotor Drive System — Operational Check ......................
Main Rotor Drive System — Scheduled Inspection..................
Main Rotor Drive System —Condition and Security Inspection
Transmission Top Case — Torque Check Special
Instructions........................................................................
Mast — 12 Month Inspection ............................................
Mast — 60 Month Inspection ............................................
Transmission — 60 Month Inspection...............................
Freewheel — 60 Month Inspection ...................................
Main Rotor Drive System — Permitted Leakage Rate .............
Serviceability Check.......................................................................
Transmission and Freewheel Assemblies — Serviceability
Check........................................................................................
Unwanted Particles ........................................................................
Unwanted Particles — Visual Identification ..............................
Unwanted Particles — Chemical Identification .........................
MAST ASSEMBLY
Mast Assembly...............................................................................
Mast Assembly — Removal......................................................
Mast Assembly — Cleaning......................................................
Mast Assembly — Inspection and Repair.................................
Mast Assembly — Installation...................................................
TRANSMISSION ASSEMBLY
Transmission Assembly .................................................................
Transmission — Removal.........................................................
Transmission — Cleaning.........................................................
Transmission — Inspection and Repair....................................
Transmission — Installation......................................................
Transmission — Input Seal.......................................................
Transmission Input Magnetic Carbon Seal — Removal ...
Transmission Input Magnetic Carbon Seal — Cleaning ...
Transmission Input Magnetic Carbon Seal — Inspection and Repair ........................................................................
Transmission Input Magnetic Carbon Seal — Installation
Transmission Input Lip Seal — Removal..........................
Transmission Input Lip Seal — Cleaning..........................
Transmission Input Lip Seal — Inspection........................
Transmission Input Lip Seal — Installation.......................
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
63-00-00
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Page
Number
33
33
36
37
37
25
25
25
26
26
26
27
21
24
9
9
27
28
28
31
49
52
55
57
57
59
44
47
47
49
39
39
43
43
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Page 1
BHT-407-MM-6
63-51
63-52
63-53
63-54
63-55
63-56
63-57
63-58
63-59
63-60
63-61
63-43
63-44
63-45
63-46
63-47
63-48
63-49
63-50
63-35
63-36
63-37
63-38
63-39
63-40
63-41
63-42
Paragraph
Number
63-62
63-63
63-64
63-65
63-66
63-67
63-68
63-69
63-70
63-71
63-72
63-73
TABLE OF CONTENTS (CONT)
Title
Chapter/Section
Number
Page
Number
PYLON ASSEMBLY
Pylon Assembly .............................................................................
Pylon Beam Assembly..............................................................
Pylon Beam Assembly — Removal ..................................
Pylon Beam Assembly — Cleaning ..................................
Pylon Beam Assembly — Inspection ................................
Pylon Beam Assembly — Repair......................................
Pylon Beam Assembly — Installation ...............................
Pylon Stop Transmission Fitting ...............................................
Pylon Stop Transmission Fitting — Removal....................
Pylon Stop Transmission Fitting — Cleaning....................
Pylon Stop Transmission Fitting — Stripping....................
Pylon Stop Transmission Fitting —Inspection ..................
Pylon Stop Transmission Fitting — Painting.....................
Pylon Stop Transmission Fitting — Installation.................
Pylon Stop Deck Fitting ............................................................
Pylon Stop Deck Fitting — Removal.................................
Pylon Stop Deck Fitting — Cleaning.................................
Pylon Stop Deck Fitting — Inspection...............................
Pylon Stop Deck Fitting — Installation..............................
Restraint Spring Assembly .......................................................
Restraint Spring Assembly — Removal............................
Restraint Spring Assembly — Disassembly......................
Restraint Spring Assembly — Cleaning............................
Restraint Spring Assembly — Inspection..........................
Restraint Spring Assembly — Repair ...............................
Restraint Spring Assembly — Assembly ..........................
Restraint Spring Assembly — Installation.........................
ENGINE TO TRANSMISSION DRIVESHAFT
Engine to Transmission Driveshaft ................................................
Engine to Transmission Driveshaft — Removal .......................
Engine to Transmission Driveshaft — Cleaning .......................
Engine to Transmission Driveshaft — Inspection .....................
Engine to Transmission Driveshaft — Repair...........................
Engine to Transmission Driveshaft — Installation ....................
FREEWHEEL ASSEMBLY
Freewheel Assembly......................................................................
Freewheel Assembly — Removal.............................................
Freewheel Assembly — Cleaning.............................................
Freewheel Assembly — Inspection and Repair........................
Freewheel Assembly — Installation..........................................
Freewheel — Aft Support Assembly Seal.................................
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113
113
120
120
120
123
97
97
105
105
110
110
88
90
90
90
86
86
86
88
90
93
93
84
84
86
86
83
83
84
84
70
80
80
83
69
69
69
70
63-00-00
Page 2 Rev. 23 7 NOV 2006
(
5.
Examine the engine to transmission driveshaft paragraph 63-65
63-5.
1.
Refer to BHT-ALL-SPM for specifications.
NUMBER NOMENCLATURE
C-561
).
Transmission Top Case — Torque Check
Special Instructions
MATERIALS REQUIRED
Thixogrease
Remove the Frahm assembly (Chapter 62).
NOTE
Do not disturb the upper cone to make it easier to do the mast nut torque check.
2.
Remove the mast nut (Chapter 62).
3.
Remove the plastic cap plug (10, Figure 63-14 ) from the mast inside diameter.
6.
If the protective coating is removed, scratched, or cracked, remove the mast (1) and refer to the
BHT-407-CR&O for protective coating application.
7.
Install the plastic cap plug (10).
CAUTION
BHT-407-MM-6
4.
Use a bright light and examine the mast (1) inside diameter for corrosion and condition of the protective coating.
5.
If there is corrosion, the mast (1) must be removed, inspected, and repaired as per the
BHT-407-CR&O.
MATERIALS REQUIRED
Refer to BHT-ALL-SPM for specifications.
NUMBER NOMENCLATURE
C-308 Sealant
1.
Apply the minimum torque required for the nut
(11, Figure 63-14 ). If the fastener does not move, the test is completed.
2.
If the fastener moved, examine the split line at the top case, the ring gear gearshaft support case, and the lower case to make sure that there are no cracks in the sealant. If there is a crack in the sealant, clean the old sealant and apply sealant (C-308).
3.
If the fastener moved, do the transmission top case torque check again at the same scheduled interval (Chapter 5) until the fastener does not move.
63-6.
Mast — 12 Month Inspection
DO NOT APPLY THIXOGREASE TO THE
THREADS OF THE MAST OR MAST NUT.
8.
Apply a thin layer of Thixogrease (C-561) to the surface of the upper cone that contacts the mast nut
(Chapter 62).
NOTE
Do not disturb the lower cone hardware.
9.
Install the mast nut (Chapter 62).
10. If applicable, install the Frahm assembly
(Chapter 62).
11. Do a mast nut torque check (Chapter 5).
63-7.
Mast — 60 Month Inspection
1.
Do this inspection if an overhaul inspection has not been done in the last 60 months. Do not do a non destructive inspection.
2.
If the primer protective coating of the mast inside is damaged, refer to the BHT-407-CR&O for primer stripping and application instructions.
3.
Remove the mast assembly ( paragraph 63-17 ).
4.
Disassemble the mast assembly
(BHT-407-CR&O).
5.
Examine the mast parts for condition
(BHT-407-CR&O).
29 SEP 2008 Rev. 25
63-00-00
Page 25
BHT-407-MM-6
6.
Assemble the mast assembly (BHT-407-CR&O).
7.
Apply paint on the outer surface of the mast where damaged (BHT-407-CR&O).
8.
Install the mast assembly ( paragraph 63-20 ).
63-8.
Transmission — 60 Month Inspection
1.
Do this inspection if an overhaul inspection has not been done in the last 60 months. Do not do a non destructive inspection. It is not necessary to paint strip the parts.
2.
Remove the transmission assembly ( paragraph
63-22 ).
3.
Remove the transmission top case, the planetary assembly, and the sungear (BHT-407-CR&O).
4.
Do not disassemble the planetary assembly.
5.
Examine all of the visible parts and assembled components for condition (BHT-407-CR&O).
6.
Insert a mirror between the spiral bevel gear edge and the lower case and use a bright light to examine the input pinion gear and surrounding areas for condition.
7.
Remove the transmission lower chip detector and chip detector housing and screen ( paragraph 63-195 ).
Use a bright light and a mirror to examine the spiral bevel gear, the gearshaft, the accessory drive gear and the surrounding areas for condition.
8.
Install the transmission chip detector housing and screen and the lower chip detector ( paragraph
63-197 ).
9.
Install the sungear, the planetary and the top case (BHT-407-CR&O).
10. Apply paint where damaged (BHT-407-CR&O).
11. Install the transmission assembly ( paragraph
63-25 ).
12. Do a transmission top case assembly torque check (Chapter 5).
63-00-00
Page 26 Rev. 25 29 SEP 2008
63-9.
Freewheel — 60 Month Inspection
1.
Do this inspection if an Overhaul Inspection has not been done in the last 60 months. Do not do a Non
Destructive Inspection. It is not necessary to paint strip the parts.
2.
Remove the freewheel assembly ( paragraph
63-69 ).
3.
Disassemble the freewheel assembly
(BHT-407-CR&O).
4.
Inspect the freewheel parts for condition
(BHT-407-CR&O).
5.
Assemble the freewheel assembly
(BHT-407-CR&O).
6.
Apply paint where damaged (BHT-407-CR&O).
7.
Install the freewheel assembly ( paragraph 63-72 ).
63-10.
MAIN ROTOR DRIVE SYSTEM —
PERMITTED LEAKAGE RATE
Table 63-1. Maximum Permitted Leakage Rates for the Transmission and Freewheel Assembly
COMPONENT
Transmission and
Freewheel
Assembly
TYPE
Static
LEAKAGE RATE
The input quill leakage must not be more than 5 drops per minute.
Transmission and
Freewheel
Assembly
The total transmission leakage at all sources with the freewheel assembly must not be more than 10 drops per minute.
Dynamic One quart per 3 hours of operation time.
BHT-407-MM-6
63-11. SERVICEABILITY CHECK
63-12.
TRANSMISSION AND FREEWHEEL
ASSEMBLIES — SERVICEABILITY
CHECK
1.
Drain the transmission and freewheel oil system
(Chapter 12) through a paper filter and save the paper filter for the inspection. Flush the transmission with clean oil.
2.
Examine the electrical magnetic element of the transmission top chip detector for unwanted particles.
3.
Examine the electrical magnetic element of the transmission lower chip detector and the screen of the chip detector housing for unwanted particles.
4.
Examine the electrical magnetic element of the freewheel chip detector for unwanted particles.
5.
Collect and identify the unwanted particles
( paragraph 63-13 ).
6.
Install the transmission top chip detector.
7.
Install the transmission lower chip detector.
8.
Install the freewheel chip detector.
9.
Remove the filter element. Cut open the filter element and examine for unwanted particles.
10. Replace the filter element with a serviceable filter element.
11. If the indicating bypass valve assembly is pushed out, do the steps that follow:
a. Remove the oil cooler and replace it with a serviceable unit (Chapter 79).
b. Fully flush all the oil system lines.
c. Remove the freewheel lubrication restrictor.
d. Clean the restrictor.
e. Install the freewheel lubrication restrictor.
f. Connect the hose assemblies.
g. Remove the jet assemblies No. 1, No. 2, No. 3 and No. 4.
h. Clean and examine the jet assemblies No. 1,
No. 2, No. 3 and No. 4.
i. Install the jet assemblies No. 1, No. 2, No. 3 and No. 4.
j. Disconnect the hose from the union filter.
k. Remove the union filter.
l. Clean and examine the union filter.
m. Install the union filter.
n. Connect the hose to the union filter.
12. Fill the transmission and the freewheel oil system
(Chapter 12).
13. Do an operational check ( paragraph 63-2 ).
14. Remove and examine the transmission top chip detector for unwanted particles.
15. Remove and examine the transmission lower chip detector for unwanted particles.
16. Remove and examine the freewheel chip detector for unwanted particles.
17. Drain the transmission and freewheel oil system
(Chapter 12) through a clean paper filter. Compare the filter with the previous results.
NOTE
If you do the serviceability check because of metal particle contamination and the number of particles has increased or the particles are large enough to be identified as chips from the gear or bearing, replace the effected component. If only the freewheel chip detector is contaminated with unwanted particles, replace the freewheel assembly. If the number of particles has decreased and you find only minute particles, the transmission assembly and freewheel assembly are serviceable.
18. Install the transmission top chip detector.
29 SEP 2008 Rev. 25
63-00-00
Page 27
BHT-407-MM-6
19. Install the transmission lower chip detector.
20. Install the freewheel chip detector.
21. Fill the transmission and freewheel oil system
(Chapter 12).
63-13. UNWANTED PARTICLES
2.
Use a magnetic retraction tool to examine the unwanted particles. The magnet will collect only ferrous metal particles.
3.
If you find a small amount of particles, they can be signs of normal wear. When the particles are large enough to be identified as part of a component, replace the component. If the particle is too small to be identified visually, refer to paragraph 63-15 for the chemical identification.
63-14.
UNWANTED PARTICLES — VISUAL
IDENTIFICATION
4.
If you find metal particles or you are not sure of the serviceability of the component, do a serviceability check ( paragraph 63-11 ).
NOTE
When you find metal particles, or if you are not sure of the serviceability of the transmission or freewheel, do a serviceability check ( paragraph 63-12 ).
5.
The chip detectors will trigger an annunciator on the caution and warning panel if there are ferrous metal particles caught in them.
1.
Visually examine the unwanted particles you find after the serviceability check or troubleshooting procedure.
6.
See Table 63-2 and Figure 63-13 to identify the unwanted particles that you can find on the chip detectors or in the oil filter or paper filter that you used for the serviceability check.
Table 63-2. Identification of Unwanted Materials
MATERIAL
Aluminum and
Magnesium
Silver
Copper
(Bronze) form.
MATERIAL YOU CAN FIND IN THE FILTERS OR SCREENS
DESCRIPTION
The particles are in granular form or look like miniature lathe turnings.
Small flakes or powder.
The particles are in granular
NECESSARY PROCEDURE
No procedure is necessary if the quantity is small and you find it at the first inspection after the overhaul or major maintenance.
If you find particles at the subsequent inspection or if the quantity is large, replace the component.
No procedure is necessary if you find it during the first 100 hours of operation, overhaul or at first inspection. If you find it after the first inspection of the first 100 hours and the quantity is large, replace the component.
If the quantity is large, replace the component.
CAUSE
This can be the result from the use of mallets or drifts at assembly. It can also show wear of the oil pump interior surfaces or unusual interference.
This can be result from wear of the silver plated components such as bearing cages and input pinion gear teeth. The quantity can be relatively large until the components fully "break-in".
This can be an indication of too much wear to the oil pump sleeve bearings or the bronze cages.
Phenolic
Rubber
Chips.
Chips, flakes or powder.
Different shapes and sizes, usually have one rounded side.
None.
None.
None.
Results from the use of mallets or drifts at assembly.
Results from the use of mallets or drifts at assembly.
Material cut from the packings at assembly.
63-00-00
Page 28 Rev. 25 29 SEP 2008
BHT-407-MM-6
TRANSMISSION ASSEMBLY
63-21. TRANSMISSION ASSEMBLY
The transmission assembly is made up of a top case, support, and lower case that contain a bevel gear and shaft arrangement, an input pinion, a ring gear, a planetary gear train, sun gear and an accessory gear drive. The components that are attached to the transmission assembly are the mast assembly, swashplate, hydraulic pump, engine to transmission driveshaft, monopole sensor, two chip detectors, four oil jets, oil pump and the oil filter manifold and housing.
The oil filter manifold and housing contain the temperature switch, thermostatic valve, temperature bulb, bypass valve and the bypass valve and indicator.
MATERIALS REQUIRED
Refer to BHT-ALL-SPM for specifications.
NUMBER NOMENCLATURE
C-428 Caps and/or Plugs
1.
Remove the external power from the helicopter.
Disconnect the battery power (Chapter 96).
The transmission assembly is attached to the roof of the helicopter, forward of the engine by the pylon beam assembly. The pylon beams are attached to the transmission assembly at two points on each side of the transmission top case by elastomeric corner mounts and at two points on each side of the lower case by the elastomeric restraints.
2.
Remove the forward fairing assembly, transmission cowling assembly, and engine air inlet cowling (Chapter 53) to get access to the transmission assembly (2, Figure 63-15 ).
3.
Drain the transmission oil system (Chapter 12).
4.
Drain the hydraulic fluid (Chapter 12).
The transmission assembly gives a two stage reduction of 15.29 to 1.0 (6317 to 413 RPM). The first stage is the bevel gear arrangement with a reduction of 3.26 to 1.0 (6317 to 1936 RPM). The second stage is from the planetary gear train that gives a reduction of 4.69 to 1.0 (1936 to 413 RPM). A complete hydraulic system power pack is mounted on the left-hand side of the transmission on the adapter housing that is mounted on the oil pump. The oil pump is driven by transmission accessory gear that gives a reduction of 1.42 to 1.0 (6317 to 4445 RPM).
5.
Disconnect the pitch links from the main rotor hub
(Chapter 67).
6.
Remove the Frahm damper from the main rotor hub (Chapter 62).
7.
Remove the main rotor hub and blade assembly
(Chapter 62).
63-22.
TRANSMISSION — REMOVAL
SPECIAL TOOLS REQUIRED
8.
Disconnect the cyclic control tubes and the collective control link from the swashplate assembly and the bellcranks on the transmission control support assembly (Chapter 67).
NUMBER
T102102
T102137-107
T103314-101
NOMENCLATURE
Dehydrator
Lifting Tool
Cover and Lift Plate
9.
Disconnect the collective and cyclic control tubes from the servo actuators and the bellcranks on the transmission control support assembly (Chapter 67).
10. Remove the transmission control support assembly with the bellcranks (Chapter 67).
7 NOV 2006 Rev. 23
63-00-00
Page 39
BHT-407-MM-6
SEE DETAIL
G
SEE DETAIL
A
SEE DETAIL
B
SEE DETAIL
C
SEE DETAIL
D
SEE DETAIL
E
SEE DETAIL
F
DETAIL
A
DETAIL
B
DETAIL
C
DETAIL
D
DETAIL
E
DETAIL
F
Figure 63-15. Transmission Assembly — Removal/Installation (Sheet 1 of 2)
407_MM_63_0050_c1
63-00-00
Page 40 Rev. 25 29 SEP 2008
BHT-407-MM-7
MAINTENANCE MANUAL
VOLUME 7
TAIL ROTOR/TAIL ROTOR DRIVE SYSTEM
NOTICE
The instructions set forth in this manual, as supplemented or modified by Alert Service Bulletins (ASB) or other directions issued by Bell Helicopter Textron Inc. and Airworthiness Directives (AD) issued by the applicable regulatory agencies, shall be strictly followed.
COPYRIGHT NOTICE
COPYRIGHT 2008
BELL ® HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON
CANADA LTD.
ALL RIGHTS RESERVED
22 FEBRUARY 1996
REVISION 25 — 29 SEPTEMBER 2008
BHT-407-MM-7
PROPRIETARY RIGHTS NOTICE
These data are proprietary to Bell Helicopter Textron Inc. Disclosure, reproduction, or use of these data for any purpose other than helicopter operation or maintenance is forbidden without prior written authorization from
Bell Helicopter Textron Inc.
PN
Additional copies of this publication may be obtained by contacting:
Commercial Publication Distribution Center
Bell Helicopter Textron Inc.
P. O. Box 482
Fort Worth, Texas 76101-0482
Rev. 24 2 OCT 2007
BHT-407-MM-7
I
H
B
G
C
D
K
J F A E
SPIRAL BEVEL PINION (406-040-420-103)
MATERIAL: ALLOY STEEL
NO.
1
2
3
4
5
6
7
REF
LTR
A
A
B
C
D
E
F
CHARACTERISTIC
Spline wear
Corrosion/scratches
Packing groove damage
Corrosion/mechanical
Corrosion, nicks, dents, scratches
Thread damage
Inside diameter corrosion
INSPECTION
METHOD
Measure above
0.120 inch
(3.05 mm) diameter pins
Measure
Visual
Measure
Measure
Visual/measure
Measure
LIMIT
Minimum 1.4267 inches (36.2 mm).
Maximum 0.004 inch (0.102 mm).
Depth 1/3 of tooth surface, 2 teeth maximum. Finish to be 32 RMS with a
0.1 inch (2.54 mm) minimum radius.
None permitted.
7
0.005 inch (0.127 mm) maximum depth.
0.010 inch (0.254 mm) maximum depth, 0.25 inch (6.35 mm), maximum length, two for each quadrant. Finish to be 32 RMS with a 0.1 inch
(2.54 mm) minimum radius.
One thread pitch, 1/3 thread depth,
0.250 inch (6.35 mm) maximum length.
Maximum 0.010 inch (0.254 mm) for
180 degrees of circumference.
6
407MM_65_0050
Figure 65-9. Spiral Bevel Pinion Inspection (Sheet 1 of 2)
7 NOV 2006 Rev. 23
65-00-00
Page 25
BHT-407-MM-7
NO.
8
9
10
11
12
REF
LTR
G
H
I
J
K
CHARACTERISTIC
Corrosion/mechanical
Chipped, missing teeth
Indentation or cuts on face of teeth
Duplex bearing journal
Roller bearing journal
INSPECTION
METHOD
Visual
Visual
Feel
Measure
Measure
LIMIT
None permitted.
None permitted.
If you find indentations or cuts on the faces of the teeth that can be felt with a probe that has a 0.020 inch
(0.508 mm) radius spherical point, replace the part.
Minimum diameter 1.3782 inches
(35.0 mm).
1
Light circumferential scratches that can be felt with a probe that has a
0.010 inch (0.254 mm) radius spherical point are permitted. Axial scratches are not permitted.
3
4
2
5
1
2
3
4
5
6
7
NOTES
The shaft journal is harder than the bearing inner ring. The inner ring can creep and/or fret and keep excess material on the journal. The material deposit can cause some galling or tearing of the journal surface when the bearing is removed. If there are signs of movement or damage, the dimension is given to make the local rework easier. Remove only the raised material from the surface of the journal. Polish/blend all axial scratches that are galling. Make them smooth with the initial journal surface or to a maximum width of 0.010 inch (0.254
mm), length of 0.100 inch (2.54 mm) and a radius of 0.010 inch (0.254 mm). No more than one rework for each adjacent quadrant in a 0.5 inch (12.7 mm) length of journal or four for each shaft journal is permitted.
In areas where the tooth contact pattern extends to the top edge of the tooth, remove nicks from the top edge of the teeth that extend to a maximum of 0.020 inch (0.508 mm) into the pattern. In areas outside the tooth contact pattern, remove nicks from the top edge of the teeth that extend to a maximum of 0.050 inch
(1.27 mm) onto the face of the tooth. Do the rework with a fine India stone. Finish to be 32 RMS with a
0.1 inch (2.54 mm) radius.
No rework is permitted on the active bearing surface. Nicks or indentations on the race lead in chamfer can be reworked by removal of the raised material. This can show damage to the bearing roller or retainer, it is recommended that you discard the bearing if you find these signs.
Axial scratches to the raised surface edge because of bearing removal and installation is permitted. Blend the scratches with a 400 grit aluminum oxide paper with a minimum diameter of 1.378 inches (35.0 mm).
The tooth pattern that shows pitting, scoring and spalling is cause for rejection. Pitting that causes patchy wear to the contact zone of the tooth can be identified as micropitting and is cause for rejection.
Rework evenly to prevent balance problems. After the rework, fill with primer (C-204), drain and put in a vertical position to dry.
Remove the wear step between the worn and unworn areas of the spline teeth with an India stone. This makes a smooth change.
407MM_65_0051_c1
Figure 65-9. Spiral Bevel Pinion Inspection (Sheet 2 of 2)
65-00-00
Page 26 Rev. 25 29 SEP 2008
BHT-407-MM-7
8.
Slide the splined adapter (5) or the splined flywheel adapter (23) off the oil cooler blower shaft.
65-27. Forward Short Shaft Assembly — Paint
Removal and Application
CAUTION
MA K E S URE TH AT THE OR DE R I N
WHICH THE DISC SEGMENTS ARE
STACKED DOES NOT CHANGE AFTER
T H E C O U P L I N G D I S C PA C K H A S
OPERATED. FIGURE 65-12 CONTAINS
INSTRUCTIONS FOR MAINTENANCE
AND REPAIR.
9.
Put lockwire (C-405) through an empty bolt hole in the coupling disc pack (22) to make sure that the order of the discs are not changed.
NOTE
Complete step 10 or step 11 as it applies to the helicopters configuration.
10. For helicopters without air conditioning, remove the nuts (4), the flat washers (3), the coupling washers
(2), and the bolts (1) that attach the coupling disc pack
(22) to the splined adapter (5) or the splined flywheel adapter (23).
11. For helicopters with air conditioning and with a balance plate (24), remove the nuts (34, Figure 65-11 , sheet 4), the flat washers (31), the bevel washers (30 and 29), the flat washers (31) and the bolts (33).
12. Put lockwire (C-405) through an empty bolt hole in the coupling disc pack (19, Figure 65-11 ) to make sure the order of the discs are not changed.
13. Remove the nuts (15), the flat washers (14), the coupling washers (13), and the bolts (12) that attach the coupling disc pack (19) to the freewheel output adapter (16).
65-26. Forward Short Shaft Assembly — Cleaning
1.
To clean the forward short shaft assembly (11,
Figure 65-11 ), refer to the BHT-ALL-SPM.
2.
To clean the coupling disc packs (22), refer to
Figure 65-12 .
NUMBER
C-208
C-305
C-309
C-344
C-436
MATERIALS REQUIRED
Refer to BHT-ALL-SPM for specifications.
NOMENCLATURE
Epoxy/Zinc Coating
Aliphatic Naphtha
MEK
Alcoholic Phosphoric
Cleaner
Paint Remover
NOTE
You do not need to remove the primer from the inside of the shaft to do the inspection.
1.
If the primer is damaged or if necessary, remove the primer from the inside of the shaft with paint remover (C-436). Refer to the BHT-ALL-SPM for the procedures.
2.
To paint the inside of the shaft, do the steps that follow:
a. Clean the inside of the shaft with aliphatic naphtha (C-305) or MEK (C-309).
b. Treat the shaft surface with alcoholic phosphoric cleaner (C-344). Refer to the
BHT-ALL-SPM for the procedures.
c. Install a tapered nonmetallic plug in one end of the shaft. Flow epoxy/zinc coating (C-208) along the inside of the shaft while you turn the shaft. Make sure the internal surface has a full layer of epoxy/zinc coating (C-208).
d. Put the shaft in a vertical position and remove the tapered plug. Permit the epoxy/zinc coating
(C-208) to drain out. Two layers of epoxy/zinc coating
(C-208) are required. The shaft must stay in a vertical position to stop the formation of epoxy/zinc coating
(C-208) puddles and to keep dynamic balance.
7 NOV 2006 Rev. 23
65-00-00
Page 39
BHT-407-MM-7
Figure 65-12. Coupling Disc Packs — Inspection and Repair
65-00-00
Page 40 Rev. 25 29 SEP 2008
10
11
2
1
3
3
6
BHT-407-MM-7
7
5
9
12
0.100 IN.
(2.54 mm)
4
4
90°
12 O'CLOCK POSITION,
6 O'CLOCK 180°
OPPOSITE
3
8
5.
6.
7.
8.
1.
2.
3.
4.
9.
10.
11.
12.
Bolt
Bolt, T103388-103
Coupling washer
Coupling disc pack
Aft short shaft
Flat washer
Nut
Dial indicator
Oil cooler blower shaft
Dial indicator holding sleeve
Post, T103388-105
Splined adapter
150 TO 180 IN-LBS
(16.95 TO 20.34 Nm)
NOTES
1. Maximum total indicator reading 0.035 inch (0.89 mm) for S/N 53000 through 53554.
Nominal total indicator reading 0.024 inch (0.61 mm) for S/N 53000 through 53554.
2.
Maximum total indicator reading 0.035 inch (0.89 mm) for S/N 53555 and subsequent.
Nominal total indicator reading 0.031 inch (0.79 mm) for S/N 53555 and subsequent.
3
4
5.
The dial indicator is shown at this location for clarity. Take readings at the 6 and 12 o'clock positions.
Shield not shown for clarity.
Balance plate (if installed) not shown for clarity. Refer to Figure 65-28 for installed position of balance plate.
6
7
6
6
The allowable tolerance is dictated by the maximum value shown. Nominal value is provided for reference only.
If tailboom P/N 407-030-801-205 or subsequent is installed on helicopter S/N 53000 through
53554, use indicator values applicable to S/N 53555 and subsequent.
7
407MM_65_0081_c1
Figure 65-25. Oil Cooler Blower Shaft — Aft Alignment
29 SEP 2008 Rev. 25
65-00-00
Page 77
BHT-407-MM-7
c. If the disc pack gap exceeds 0.015 inch (0.381
mm), or if the other bolts are loose, tighten all the bolts
( paragraph 65-60 ).
d. Install the post (11) T103388-105, and the
196K dial indicator holding sleeve (10).
e. Remove the threaded tip from the Starrett
196B dial indicator (8).
f. Install the Starrett 196B dial indicator (8) on the dial indicator holding sleeve (10) as shown in
Figure 65-25 .
g. Set the travel of the Starrett 196B dial indicator (8) so that it has a minimum range of ±0.04
inch (±1.02 mm) and does not reach the end of travel in that range.
CAUTION
E N S U R E T H A T T H E M A I N R O T O R
BLADES AND THE TAIL ROTOR BLADES
ARE CLEAR BEFORE YOU ROTATE THE
OIL COOLER BLOWER SHAFT. IF THE
BLADES HIT AN OBJECT, THEY CAN BE
DAMAGED.
h. Use an inspection mirror to read the Starrett
196B dial indicator (8). Adjust the dial indicator to read zero at the 12 o’clock position ( Figure 65-25 , Note 3).
Turn the shaft one-half turn to the 6 o’clock position.
Record the reading off the dial indicator. Refer to
Figure 65-25 , Notes 1 and 2, for limits.
i. If the value is not within the limits shown in
Figure 65-25 , take no remedial action until you measure the alignment of the forward bearing hanger bracket as specified in step 4 .
j. Remove the Starrett 196B dial indicator (8), holding sleeve (10), post (11) and bolt (2). Install the bolt (1), the coupling washers (3), the flat washer (6), and the nut (7). Do not tighten the nut (7) at this time.
4.
Measure the alignment of the oil cooler blower forward bearing hanger bracket as follows:
a. Remove one nut (1, Figure 65-26 ), the flat washer (11), the coupling washers (10), and the bolt
(9).
65-00-00
Page 78 Rev. 23 7 NOV 2006
on
CAUTION
INSTALL THE COUPLING WASHERS
WITH THE BEVELLED SIDE AGAINST
THE COUPLING DISC PACK.
b. Install the bolt (8), T103388-103, the coupling washers (10), the flat washer (11), and the nut (1)
T
Make sure the short end of the bolt (8) faces forward.
.
c. If, after you tighten the disc pack, the gap exceeds 0.015 inch (0.381 mm), or if the other bolts are loose, tighten all the bolts and nuts ( paragraph
65-60 ).
d. Install the Post (7), T103388-105, and the
Starrett dial indicator holding sleeve (6), 196K, on the bolt (8).
e. Install the Starrett 196B dial indicator (5) on the holding sleeve (6) as shown in Figure 65-26 .
f. Set the travel of the Starrett 196B dial indicator (5) so that it has a range of ±0.04 inch (±1.02
mm) and does not reach the end of travel in that range.
CAUTION
E N S U R E T H A T T H E M A I N R O T O R
BLADES AND THE TAIL ROTOR BLADES
ARE CLEAR BEFORE YOU ROTATE THE
OIL COOLER BLOWER SHAFT. IF THE
BLADES HIT AN OBJECT, THEY CAN BE
DAMAGED.
g. Use an inspection mirror to read the Starrett
196B dial indicator (5). Adjust the dial indictor to read zero at the 12 o’clock position ( Figure 65-26 , Note 3).
Turn the shaft one-half turn to the 6 o’clock position.
Record the reading off the dial indicator. Refer to
Figure 65-26
196B dial indicator (5, at this time.
, Notes 1 and 2, for limits.
h. If both ends of the shaft are in the limits shown
Figure 65-25 and Figure 65-26 , remove the Starrett
Figure 65-26 ) and the bolt (8).
Install the bolt (9), the coupling washers (10), the flat washer (11) and the nut (1). Do not tighten the nut (1)
BHT-407-MM-7
10
11
9
10
1
8
4
6
7
2
3
0.76 IN.
(19.30 mm)
12 O'CLOCK
POSITION,
6 O'CLOCK
3
180° OPPOSITE
1.
2.
3.
4.
5.
6.
7.
8.
9.
90°
Nut
Splined adapter (helicopter
S/N 53443 and subsequent, or
S/N 53000 through 53442
Post BHT-407-II-30)
Forward short shaft
Coupling disc pack
Dial indicator
Dial indicator holding sleeve
Post, T103338-105
Bolt, T103338-103
Bolt
5
3
12
10.
Coupling washer
11.
Flat washer
12.
Splined flywheel adapter
(helicopter S/N 53000 through
53442 Pre BHT-407-II-30)
150 TO 180 IN-LBS
(16.95 TO 20.34 Nm)
NOTES
1.
2.
3
Maximum total indicator reading 0.039 inch (0.99 mm) for S/N 53000 through 53554.
Nominal total indicator reading 0.028 inch (0.71 mm) for S/N 53000 through 53554.
Maximum total indicator reading 0.035 inch (0.89 mm) for S/N 53555 through subsequent.
Nominal total indicator reading 0.028 inch (0.71 mm) for S/N 53555 through subsequent.
The dial indicator is shown at this location for clarity. Take readings at the 6 and 12 o'clock positions.
4. Balance plate (if installed) not shown for clarity. Refer to Figure 65-11 for installed position of balance plate.
5 The allowable tolerance is dictated by the maximum value shown. Nominal value is provided for reference only.
6 If tailboom P/N 407-030-801-205 or subsequent is installed on helicopter S/N 53000 through
53554, use indicator values applicable to S/N 53555 and subsequent.
5 6
5
407MM_65_0082_c1
Figure 65-26. Oil Cooler Blower Shaft — Forward Alignment
29 SEP 2008 Rev. 25
65-00-00
Page 79
BHT-407-MM-7
FWD
A
B
A
B
NOTE
Measure distance
A
and
B
. Difference between
A
and
B
cannot exceed 0.030 inch (0.762 mm).
Figure 65-27. Oil Cooler Blower Shaft — Centering
65-00-00
Page 80 Rev. 23 7 NOV 2006
407MM_65_0083
BHT-407-MM-7
i. Loosen the three other nuts (1) on the coupling disc pack (4). Torque the nuts (1) in accordance with the torque sequence procedure.
paragraph 65-60 .
j. Loosen the three other nuts (7, Figure 65-25 ) on the coupling disc pack (4). Torque the nuts (7) in accordance with the torque sequence procedure, paragraph 65-60 .
NOTE
Adjustment to the shims of the brackets (19 or 40, Figure 65-15 ) can affect the tail rotor driveshaft segment assembly. If the oil cooler blower shims are changed, do the tail rotor driveshaft segment alignment
( paragraph 65-57 ).
Prior to any adjustment of shims, contact product support engineering. Product support can evaluate the values measured and provide appropriate recommendations.
k. If a measurement is less than the nominal total indicator reading shown in Figure 65-25 and
Figure 65-26 , the angle between the oil cooler blower shaft and the next shaft segment is below average but is acceptable. As the allowable tolerance is dictated by a maximum value only, this nominal value is to be used as a reference only.
NOTE
Prior to any adjustment of shims, contact product support engineering. Product support can evaluate the values measured and provide appropriate recommendations.
l. If a measurement is greater than the maximum total indicator reading shown on
Figure 65-25 and Figure 65-26 , the angle between the oil cooler blower shaft and the next shaft segment is too large. Add shims under the affected bracket, or remove shims from under the bracket on the opposite end.
5.
Following the procedures for the alignment of the oil cooler blower shaft, a final check is required to ensure that end (axial) play still exists:
a. Grasp the blower shaft (51), one hand on each end. Move the shaft fore and aft. An end (axial) play, however slight, must be felt. This indicates that the bearings are not loaded axially.
b. If no end (axial) play exists, follow the procedures in paragraph 65-38 , step 5 . for helicopter
S/N 53000 through 53442 Pre TB 407-02-35 or paragraph 65-38 , step 6 for helicopter S/N 53443 and subsequent or S/N 53000 through 53443 Post TB
407-02-35.
6.
Lockwire the screws (36) and the bolts (38) with lockwire (C-405).
7.
Make sure the oil cooler blower shaft (51) is vertically centered in the oil cooler blower housing inlet covers (26 and 50). The oil cooler blower shaft (51) must not be more than 0.030 in. (0.762 mm) off center
( Figure 65-28 ). If the oil cooler blower shaft is more than 0.030 inch (0.762 mm), peel or replace the shims
(25, Figure 65-15 ). If you replace the shims (25), they must be bonded to the helicopter structure with adhesive (C-317).
a. Tighten the screws (23 and 43)
T
.
b. Lockwire the screws (23 and 43) with lockwire
(C-405).
65-40. Oil Tank and Oil Cooler — Installation
1.
Install the oil tank (Chapter 79).
2.
Install the oil cooler and transition duct (Chapter
79).
3.
Connect the oil inlet and outlet lines on the transmission and engine oil coolers (Chapter 79).
4.
Service the engine oil systems (Chapter 12).
5.
Do an operational check ( paragraph 65-2 ).
6.
Install the cowling (Chapter 53).
29 SEP 2008 Rev. 25
65-00-00
Page 81
BHT-407-MM-7
BEVELLED SIDE MUST
FACE COUPLING DISC PACK
DETAIL
A
2
3
4
1
5
SEE DETAIL
A
20
19
18
21
20
22
17
3
16
15
SEE DETAIL
A
20
21
20
22
10
11
12
14
13
1
1
2
3
4
5
6
7
10
9
6
7
9
8
11
12
18
23
4
14
13
15
16
26
25
24
17
3
2
SEE DETAIL
B
8
Figure 65-28. Aft Short Shaft Assembly — Removal/Installation (Sheet 1 of 2)
407MM_65_0084
65-00-00
Page 82 Rev. 23 7 NOV 2006
BHT-407-MM-9
MAINTENANCE MANUAL
VOLUME 9
POWER PLANT
NOTICE
The instructions set forth in this manual, as supplemented or modified by Alert Service Bulletins (ASB) or other directions issued by Bell Helicopter Textron Inc. and Airworthiness Directives (AD) issued by the applicable regulatory agencies, shall be strictly followed.
COPYRIGHT NOTICE
COPYRIGHT 2008
BELL ® HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON
CANADA LTD.
ALL RIGHTS RESERVED
22 FEBRUARY 1996
REVISION 25 — 29 SEPTEMBER 2008
BHT-407-MM-9
PROPRIETARY RIGHTS NOTICE
These data are proprietary to Bell Helicopter Textron Inc. Disclosure, reproduction, or use of these data for any purpose other than helicopter operation or maintenance is forbidden without prior written authorization from
Bell Helicopter Textron Inc.
PN
Additional copies of this publication may be obtained by contacting:
Commercial Publication Distribution Center
Bell Helicopter Textron Inc.
P. O. Box 482
Fort Worth, Texas 76101-0482
Rev. 24 2 OCT 2007
BHT407-MM-9
71-28.
ENGINE PTP SEAL — REMOVAL
SPECIAL TOOLS REQUIRED
NUMBER
406-240-009-115
406-240-001-101
NOMENCLATURE
Puller set
Pusher set
2.
Examine the surface of the PTP shaft, where the seal is installed, for condition. If the PTP shaft is damaged, do not install the seal (refer to the
Rolls-Royce 250-C47B Operation and Maintenance
Manual, Publication CSP21001).
3.
Make sure that there are no unwanted materials in the engine.
71-30.
ENGINE PTP SEAL — INSTALLATION
1.
Drain the engine and transmission oils
(Chapter 12).
2.
Remove the freewheel assembly (Chapter 63).
3.
Install the puller set (1, Figure 71-4) on seal (8).
Make sure that you attach the puller set to the seal correctly.
MATERIALS REQUIRED
Refer to BHT-ALL-SPM for specifications.
NUMBER NOMENCLATURE
C-001 Grease
NOTE
Discard the seal (8) after you remove it. Do not install a seal that was removed.
4.
Remove the seal (8) with the slide hammer (2) on the puller set (1).
5.
Compress the puller set (1) with the hose clamp
(3) to remove the seal (8) from the puller set. Discard the seal.
71-29.
ENGINE PTP SEAL — INSPECTION AND
REPAIR
1.
Examine the seal lips for condition. Ensure the following:
• Lips are smooth
• There are no nicks on the lips or there is no missing material from the lips
• Tension spring is not damaged or loose on the seal
NOTE
Mobil 28 grease is recommended in the following step.
1.
Put grease (C-001) on the lip of the seal (8,
Figure 71-4) and on the surface of the PTP shaft where you install the seal. The grease will prevent damage to the seal lip when you first start the engine.
2.
Use the pusher assembly (4) to install the seal
(8). Make sure that the metal part of the seal points out from the engine.
3.
After you install the seal (8), examine it for condition (paragraph 71-29). Make sure of the following:
• Seal is in the correct position
• Lip is not damaged
• Spring is not loose
4.
Install the freewheel assembly (Chapter 63).
5.
Service the engine and transmission
(Chapter 12).
29 SEP 2008 Rev. 25
71-00-00
Page 21
BHT-407-MM-9
Figure 71-4.
71-00-00
Page 22
Figure 71-4. Engine PTP seal — Installation and removal (Sheet 1 of 2)
Rev. 7
BHT-407-MM-9
76-18
76-19
76-20
76-21
76-6
76-7
76-8
76-9
76-10
76-11
76-12
76-13
76-14
76-15
76-35
Paragraph
Number
76-1
76-2
76-3
76-4
76-5
CHAPTER 76 — ENGINE CONTROLS
TABLE OF CONTENTS
Title
Chapter/Section
Number
GENERAL
Full Authority Digital Electronic Control (FADEC) ..........................
FADEC System Acronyms........................................................
FADEC Control Features..........................................................
FADEC System Components ...................................................
Power Up Mode and Built In Test .............................................
OPERATION
Start In Auto Mode — Description .................................................
Start in Auto Mode — Alternate Start .......................................
Start In Manual Mode — Description .............................................
FADEC Manual Check — Description ...........................................
In-flight — Auto Mode Operation ...................................................
NDOT Control ...........................................................................
Gas Generator (N
G
Power Turbine (N
) Governor..................................................
) Governor ..................................................
Engine Auto Relight .......................................................................
Auto Mode — Auto Relight (N
Above 50%)............................
Auto Mode — Pilot Assisted In-flight Restart (N
and 50%) ..................................................................................
Manual Mode — Relight ................................................................
Engine Overspeed and Protection — Description .........................
N
P
Overspeed................................................................................
Power Up Functional Check .....................................................
Continuous Functional Check...................................................
Overspeed System Failure Annunciation .................................
Overspeed System Shutdown Check .......................................
Overspeed................................................................................
Engine Shutdown...........................................................................
HMU Manual Piston Parking Procedure ...................................
FADEC System Faults..............................................................
Category 1 — FADEC Failure ..................................................
Auto to Manual Mode Transition...............................................
Category 2 — FADEC Degraded..............................................
Category 3 — FADEC Fault .....................................................
Category 4 — Restart Fault ......................................................
Category 5 — Maintenance Advisory, FADEC System Faults
— Engine Shutdown .................................................................
FADEC Faults/Exceedances — Recording Procedure..................
FADEC Faults/Exceedances — Clearing Procedure.....................
Recorded Exceedances/Conversion Factors.................................
, MGT, and Torque (Q) Exceedances Recorded
by FADEC (ECU) ...........................................................................
76-00-00
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76-00-00
76-00-00
76-00-00
Page
Number
76-00-00
76-00-00
76-00-00
76-00-00
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76-00-00
76-00-00
5
6
5
5
7
20
20
20
20
21
21
15
16
16
19
19
19
23
23
31
31
32
15
15
15
15
15
11
13
13
14
14
32
32
32
33
34
76-00-00
Page 1
BHT-407-MM-9
Paragraph
Number
76-38
76-39
76-40
76-41
76-42
76-43
76-44
76-45
76-46
76-54
76-57
76-58
76-62
76-64
TABLE OF CONTENTS (CONT)
Title
Chapter/Section
Number
Determination of Faults/Exceedances and Required
Troubleshooting Steps ...................................................................
Maintenance Mode — Procedure for Viewing FADEC Fault
Codes Using Caution Panel Flashing Display ...............................
FADEC Fault Codes — Procedure to Determine Last Engine
Run Faults From “Current” Faults ..................................................
Check Run Procedure....................................................................
FADEC Download Intervals ...........................................................
Engine Start — Troubleshooting....................................................
76-00-00
76-00-00
76-00-00
76-00-00
76-00-00
76-00-00
HYDROMECHANICAL
Hydromechanical Unit (HMU) ........................................................
Hydromechanical Unit (HMU) — Removal ...............................
Hydromechanical Unit (HMU) — Installation ............................
MECHANICAL
Throttle/Fly Detent Rigging Procedure...........................................
Throttle/Fly Detent Friction Check .................................................
Throttle/Fly Detent Friction Adjustment..........................................
Throttle Control Cable....................................................................
Throttle Control Cable — Removal...........................................
Throttle Control Cable — Inspection.........................................
Throttle Control Cable — Installation........................................
ELECTRICAL
Electrical Engine Controls — General ...........................................
Electronic Control Unit (ECU) ...................................................
Electronic Control Unit (ECU) — Removal S/N 53000
Through 53749 Pre TB 407-07-75 ....................................
Electronic Control Unit (ECU) — Removal S/N 53000
Through 53749 Post TB 407-07-75 and S/N 53750 and
Subsequent.......................................................................
Electronic Control Unit (ECU) — Inspection .....................
Electronic Control Unit (ECU) — Installation S/N 53000
Through 53749 Pre TB 407-07-75 ....................................
Electronic Control Unit (ECU) — Installation S/N 53000
Through 53749 Post TB 407-07-75 and 53750 and
Subsequent.......................................................................
Collective Pitch Transducer (CPT) ...........................................
Collective Pitch Transducer (CPT) — Removal................
Collective Pitch Transducer (CPT) — Inspection..............
Collective Pitch Transducer (CPT) — Installation/Rigging
Collective Pitch Transducer (CPT) — Functional Test .....
Compressor Inlet Temperature (CIT) Sensor ...........................
76-00-00
76-00-00
76-00-00
76-00-00
76-00-00
76-00-00
76-00-00
76-00-00
Page
Number
49
49
49
78
78
79
79
79
83
84
55
59
59
61
61
67
67
71
71
71
74
74
77
35
36
44
44
44
45
BHT-407-MM-9
Paragraph
Number
76-67
76-68
76-69
76-70
TABLE OF CONTENTS (CONT)
Title
Chapter/Section
Number
Compressor Inlet Temperature (CIT) Sensor
— Removal .......................................................................
Compressor Inlet Temperature (CIT) Sensor
— Inspection .....................................................................
Compressor Inlet Temperature (CIT) Sensor
— Installation ....................................................................
Compressor Inlet Temperature (CIT) Sensor
— Functional Test.............................................................
76-00-00
76-00-00
76-00-00
76-00-00
Page
Number
84
85
85
85
Figure
Number
76-1
76-2
76-3
76-11
76-12
76-13
76-14
76-15
76-18
FIGURES
Title
FADEC Control System Schematic ....................................................................
HMU Schematic ..................................................................................................
Engine Overspeed Light and ECU N
Recording Activation Table for FADEC
Example of Engine History Data — Maintenance...............................................
Instrument Panel — FADEC System Switches, Caution/Warning Panel............
Auto to Manual Transition at Low Fuel Flow.......................................................
Auto to Manual Transition at Intermediate Fuel Flow .........................................
Auto to Manual Transition at High Fuel Flow ......................................................
FADEC/ECU Maintenance Button and FADEC/ECU Maintenance Terminal
Engine Control Rigging .......................................................................................
Rigging Fly Detent ..............................................................................................
Collective Throttle Friction ..................................................................................
Throttle Control Cable.........................................................................................
Electronic Control Unit (ECU) — Removal/Installation S/N 53000 Through
53749 Pre TB 407-07-75 ....................................................................................
Electronic Control Unit (ECU) — Removal/Installation S/N 53000 Through
53749 Post TB 407-07-75 and S/N 53750 and Subsequent...............................
Collective Pitch Transducer (CPT) — Removal/Installation................................
Compressor Inlet Temperature (CIT) Sensor — Removal/Installation ...............
Page
Number
17
18
22
25
27
29
8
9
72
75
80
86
37
50
51
57
60
62
Table
TABLES
Time to Power Change — (DRTM).....................................................................
250-C47B FADEC Software Version 5.202 Fault Code Display.........................
250-C47B FADEC Software Versions 5.356 and 5.358 (Reversionary Governor)
Page
Number
31
38
41
76-00-00
Page 3
BHT-407-MM-9
Table
Number
76-4
TABLES (CONT)
Title
Throttle Rigging Parameters ...............................................................................
Page
Number
56
76-00-00
Page 4 Rev. 25 29 SEP 2008
BHT-407-MM-9
GENERAL
76-1.
FULL AUTHORITY DIGITAL
ELECTRONIC CONTROL (FADEC)
This chapter contains information pertaining to the
FADEC engine control system and its associated airframe inputs. Electronic Control Unit (ECU)
Software Versions 5.202 and 5.356 or 5.358
(Reversionary Governor) with Direct Reversion to
Manual (DRTM) are addressed.
For additional information on the FADEC system, refer to the Rolls-Royce 250-C47B Operation and
Maintenance Manual, Publication CSP 21001.
Additionally, reference to Rolls-Royce Commercial
Service Letter CSL-6069 will provide general FADEC system maintenance guidelines.
76-2.
FADEC SYSTEM ACRONYMS
• Compressor Inlet Temperature (CIT)
• Combined Engine Filter Assembly (CEFA)
• Collective Pitch (CP)
• Collective Pitch Transducer (CPT)
• Caution/Warning/Advisory Panel (CWAP)
• Direct Reversion To Manual System (DRTM)
• Electronic Control Unit (ECU)
• Engine Monitoring System (EMS)
• Full Authority Digital Electronic Control
(FADEC)
• Volatile Memory – Data Lost When Power is
Removed
• Line Replaceable Unit (LRU)
• Nonvolatile Memory – Data Not Lost When
Power is Removed
• Hydromechanical Unit (HMU)
• Measured Gas Temperature (MGT)
• Gas Generator Speed (N
G
)
• Power Turbine Speed (N
P
)
• Rate of Change of N
G
Speed (NDOT)
• Main Rotor Speed (N
R
)
• Power Lever Angle (PLA) – Controlled by
Throttle Position
• Permanent Magnet Alternator (PMA)
• Resistance Temperature Device (RTD)
• Torque Meter Oil Pressure (TMOP)
• Fuel Flow (Wf)
76-3.
FADEC CONTROL FEATURES
The engine control and monitoring systems provide the following features:
1.
Fault Detection – The ECU monitors the FADEC system for faults and makes appropriate accommodation to continue operation. When the engine is operating, fault codes are stored inside the
ECU in nonvolatile memory.
2.
Automatic Start – The ECU provides for automatic control of fuel flow during engine starts to control the rate of acceleration and limit engine start temperature. The control provides a hot start abort feature that cuts fuel flow off to prevent an overtemperature start.
3.
NDOT Control System – The engine control law is based on a Gas Generator rate of acceleration
(NDOT) system that maximizes engine performance while maintaining safe engine operation.
4.
Electronically Controlled Gas Generator (N
G
)
Governor – Allows engine power control and modulation, from shutoff to maximum power, via the throttle twist grip.
5.
Electronically Controlled Power Turbine (N
P
Governor – Provides constant power turbine speed
) governing.
29 SEP 2008 Rev. 25
76-00-00
Page 5
BHT-407-MM-9
6.
Autorelight – The ECU detects engine flameout and initiates an automatic engine relight sequence.
7.
Overspeed System – The ECU protects against
Gas Generator (N
G
) (FADEC Software Version 5.202) and Power Turbine (N
P
) overspeed. The ECU provides a power up self-test and a pilot initiated test at engine shutdown to verify proper operation.
8.
Failure Annunciation – The operational status of the engine control system is automatically and continually monitored by the ECU. Should a failure be detected it will be annunciated to the pilot. When a control system fault occurs that prevents continued operation of the Auto Mode control, the FADEC FAIL will be annunciated. The FADEC DEGRADE will be annunciated when a fault has occurred that may affect engine performance, or if maintenance action is required following shutdown. The FADEC FAULT will be annunciated when a fault has occurred that does not affect engine performance, but may affect an operating feature such as engine limiting. The
RESTART FAULT will be annunciated when a fault is detected that may affect the ability of the engine to start in the Auto Mode. This failure information will be recorded by the ECU.
9.
Engine Condition Monitoring – The ECU provides an Engine Monitoring System (EMS) to record and log
FADEC system faults and engine overspeed limit exceedances.
10. Exceedance Limiting – Automatic limiting functions accomplished by the FADEC include MGT temperature limiting, N
G
speed limiting and N
P
speed limiting. In the Model 407, the FADEC provides engine
MGT limiting at the engine maximum transient (1661°F
(905°C)).
In regards to MGT limiting, the FADEC system interfaces to the MGT harness to measure engine temperature. When the engine is approaching its maximum transient temperature limit, the FADEC reduces fuel flow to prevent limit exceedance (1661°F
(905°C)). A smooth, controlled transition between governing and temperature limiting is accomplished by the FADEC.
11. Surge Detection and Recovery – The FADEC detects engine surge by comparing the rate of change of N
G
speed to a predetermined boundary rate. If the boundary is exceeded and MGT is increasing, the surge will be detected and recorded by the internal
76-00-00
Page 6 Rev. 25 29 SEP 2008
ECU Engine Monitor System (EMS). The surge will be recorded in the ECU's memory relative to the N
G speed at which it occurred. Without pilot action, the
FADEC will reduce fuel flow during the surge and reduce the maximum acceleration schedule against the current acceleration in order to recover from the surge. The FADEC will then lower the acceleration schedule at the range of N
G
where the surge occurred to avoid subsequent surge. The acceleration schedule is reset to the original schedule at the next FADEC power up.
12. Manual Mode – In manual mode, the pilot's PLA input is tied hydromechanically to the fuel flow metering valve in the HMU. Manual mode is engaged by de-energizing the auto/manual solenoid in the
HMU. This allows the pilot to vary fuel flow to the engine by moving the PLA, via the throttle twist grip.
This manual mode fuel flow is altitude compensated to allow a consistent PLA/horsepower relationship versus altitude. At 100% throttle travel, the manual mode will provide sea level rated takeoff power at sea level. The fuel flow slew rate is mechanically limited to avoid blowout and to provide proper responsiveness for helicopter operation.
13. Maintenance Mode – The ECU provides a maintenance mode function. This function identifies, by a series of flashing lights, the suspect LRU when a
FADEC fault has been indicated. The maintenance mode can also be used to identify recorded overspeed exceedances. This function is only operational on the ground and is a guide to troubleshooting.
76-4.
FADEC SYSTEM COMPONENTS
The Model 407 power plant is made up of a
Rolls-Royce 250-C47B engine with an electronic control system. The control system is based upon a single channel Full Authority Digital Electronic Control
(FADEC) that controls, monitors, and limits engine power while maintaining helicopter rotor speed. There is also a manual mode hydromechanical backup.
The FADEC system has two main components. The airframe mounted Electronic Control Unit (ECU) and the engine mounted Hydromechanical Unit (HMU).
The ECU monitors numerous internal and external inputs to modulate fuel flow and therefore control engine speed, acceleration rate, temperature and other engine parameters. The ECU provides inputs to the HMU to modulate fuel flow based on the
BHT-407-MM-9
continuous monitoring of the following: Measured Gas
Temperature (MGT), Gas Generator speed (N
G
Power Turbine speed (N
P
), Main Rotor speed (N
R
),
),
Engine Torque Meter Oil Pressure (TMOP), Collective
Pitch (CP) and rate, Compressor Inlet Temperature
(CIT), Ambient Pressure (P1), and Power Lever Angle
(PLA)/throttle position ( Figure 76-1 ).
On helicopters with FADEC Software Version 5.356 or
5.358 installed, the engine uses a digital electronic control system based on two electronic governors called primary channel and reversionary governor. The primary channel is a full-authority digital electronic control (FADEC) that controls, monitors, and limits engine power while maintaining helicopter rotor speed.
The reversionary governor can automatically take control over the engine in the event of a primary channel failure. The reversionary governor uses a limited set of inputs and provides basic electronic governing.
To be more specific, the 5.356 and 5.358 FADEC
Reversionary (backup) Governor consists primarily of a backup channel that is contained in the ECU and is isolated from the primary governor by a firewall for
EMI. It provides basic power turbine speed governing
in the event of a hard fault occurring in the primary governor. Failure of the FADEC into reversionary governor mode is indicated by the illumination of the following three lights; FADEC FAULT, FADEC
DEGRADE, and RESTART FAULT. This type of failure causes a degradation in performance and can cause
NR droop or NR lag. Operations should be continued in AUTO mode and helicopter is to be flown smoothly and non-aggressively. Applicable maintenance action will be required prior to next flight ( paragraph 76-38 ).
The engine gearbox mounted Hydromechanical Unit
(HMU) consists of an engine driven fuel pump assembly and a fuel metering unit assembly combined into one unit ( Figure 76-2 ).
The fuel pump assembly of the HMU consists of a side channel liquid ring boost stage and a gear main stage.
The fuel pump assembly contains a high pressure relief valve.
The fuel metering unit assembly of the HMU consists of a stepper motor controlled flat plate metering valve, a metering head regulator valve, a windmill bypass valve, a minimum flow bypass valve, a metering head altitude compensation valve, a pressurizing and shutoff valve, an overspeed solenoid valve, a hot start fuel solenoid valve, an auto/manual changeover solenoid valve, and a manual Wf/P1 servomechanism.
Pump discharge fuel flow passes through the metering valve and, in parallel, through the minimum flow path and out to the fuel nozzle. Excess pump discharge fuel flow is returned back to the pump gear stage inlet by the metering head regulator valve. In the auto mode, the metering valve flow area is set by stepper motor position. In the manual mode, the metering valve flow area is set as a function of PLA/throttle position. In either mode, engine shutdown is accomplished by retarding the PLA/throttle to the cutoff position. This action causes the windmill bypass valve to dump metering valve discharge pressures which, in turn, causes the pressurizing and shutoff valve to close.
Prior to exiting the HMU, fuel must pass through the overspeed solenoid valve. When energized, the overspeed solenoid valve closes, causing fuel flow to go to a minimum flow (sub-idle) condition.
76-5.
POWER UP MODE AND BUILT IN TEST
The FADEC system incorporates logic and circuitry to perform self-diagnostics. In general, sensors are checked for continuity, rate, and proper range.
Discrete inputs are checked for continuity and output drivers are monitored for current demand to sense failed actuators and open or shorted circuits. A FADEC power up check exercises output drivers and actuators to ensure system functionality and readiness.
The brief appearances of light indications and their respective horn observed immediately after application of power are normal and part of the
FADEC system’s designed initialization process.
If any faults are detected during the self-test, the appropriate FADEC caution panel light will illuminate.
The helicopter 28 VDC bus supplies electrical power to the FADEC ECU until the engine achieves 85% N
P
Above this speed, the FADEC ECU will select between
.
the 28 VDC bus and the engine-driven Permanent
Magnet Alternator (PMA) as its primary power source.
The higher voltage source will be selected. In the event of a primary power source failure, the alternate source will be selected.
29 SEP 2008 Rev. 25
76-00-00
Page 7
BHT-407-MM-9
HYDRO-
MECHANICAL
UNIT
(HMU)
N
G
FUEL NOZZLE
FUEL
FILTER
N
P
(PMA)
PERMANENT MAGNET
ALTERNATOR
(CIT)
COMPRESSOR
INLET
TEMPERATURE
GAS GENERATOR
POWER
LEVER
POWER TURBINE
FUEL
OUT
ENGINE
FUEL IN
THROTTLE
(PLA) LINKAGE
(MGT)
MEASURED
GAS
TEMPERATURE
TORQUE METER
OIL PRESSURE
(TMOP)
AIRFRAME
COCKPIT TYPICAL
60
AMBIENT
PRESSURE
(P1)
ELECTRONIC
CONTROL
UNIT (ECU)
VIBRATION
ISOLATORS
IGNITER RELAY
AUTO RELIGHT
STARTER RELAY
LATCHING
MAIN ROTOR
FADEC/ECU MTCE
PORT (MTCE
TERMINAL
CONNECTOR)
+28 VDC BATTERY/
AIRFRAME POWER
(MECHANICAL LINKAGE)
60
N /N
R
MGT
ON
OVERSPEED
TEST SWITCH
AUTO
MANUAL
FADEC MODE SWITCH
CAUTION
PANEL LIGHTS
-FADEC FAIL
-FADEC MANUAL
-FADEC DEGRADED
-FADEC FAULT
-RESTART FAULT
-ENGINE OVSPD
-ENGINE OUT
-AUTO RELIGHT
THROTTLE
POWER LEVER
ANGLE (PLA)
COLLECTIVE
PITCH (CP)
AND RATE
Figure 76-1. FADEC Control System Schematic
76-00-00
Page 8 Rev. 25 29 SEP 2008
407MM_76_0001_c02+
BHT-407-MM-9
LEGEND
PRESSURE IN
PRESSURE BEFORE FILTER
PRESSURE AFTER FILTER
FUEL
INLET
PUMP DISCHARGE PRESSURE
REGULATED PRESSURE
METERED FUEL PRESSURE
NOZZLE PRESSURE
SIDE CHANNEL
LIQUID RING
BOOST PUMP
STANDPIPE
SCREEN
METERING
VALVE
FEEDBACK
POTENTI-
OMETER
GEAR PUMP
AUTO/MANUAL
CHANGEOVER
SOLENOID VALVE
NORMALLY
ENERGIZED
CLOSED
LOAD PISTON
ORIFICES
PLA
FEEDBACK
POTENTIOMETER
FAIL FIXED
ORIFICE
MANUAL
SERVOMECHANISM
CONTROL
MANUAL PISTON
UPSTREAM
ORIFICE
MANUAL FOLLOWER
PISTON (FAST)
(SHOWN RETRACTED)
MANUAL LOAD
PISTON (SLOW)
(SHOWN RETRACTED)
POWER
LEVER
METERING VALVE
LEVER
FLAT PLATE
METERING VALVE
INC
INC
DAMPING
ORIFICE
PRESSURIZING
AND
SHUTOFF
VALVE
TO
NOZZLE
CEFA FUEL
FILTER
HIGH
PRESSURE
RELIEF VALVE
EXTERNAL
SCREEN
STEPPER MOTOR
AND GEARHEAD
MHR DAMPING
ORIFICE
MIN
FLOW
ORIFICE
METERING HEAD ( P)
REGULATOR VALVE
WINDMILL
BYPASS
VALVE
MIN FLOW
BYPASS
VALVE
NORMALLY
DE-ENERGIZED
OPEN
OVERSPEED
SOLENOID VALVE
ENERGIZED CLOSED
ALTITUDE
COMPENSATION
VALVE
EVACUATED
BELLOWS
P1 AIR
START/HOT START
ABORT SOLENOID VALVE
(NORMALLY
ENERGIZED CLOSED)
FUEL SYSTEM HYDROMECHANICAL (HMU) SCHEMATIC - AUTO MODE SHOWN
250-C47B ENGINE FADEC SYSTEM
Figure 76-2. HMU Schematic
407MM_76_0002
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76-00-00
Page 9/10
BHT-407-MM-9
OPERATION
76-6.
START IN AUTO MODE —
DESCRIPTION
The FADEC control system provides automatic start sequencing and engine control during the engine starting cycle. This involves controlling fuel flow until stabilized idle Gas Generator (N
G
) speed is reached.
Starting is initiated by the pilot by placing the power lever in the idle position and momentarily activating the start switch. The automatic start cycle can only be actuated by engaging the helicopter starter within 60 seconds of the PLA being moved to idle. Once the required lightoff (N
G
) speed is achieved, the FADEC introduces fuel to the engine. The engine fuel flow is then regulated to control the (N
G
) turbine rate of acceleration (NDOT) and to maintain a turbine temperature (MGT) within limits while accelerating to idle. Pilot fuel modulation is not required or possible.
Additionally, the control can prevent most overtemperature starts by automatically cutting fuel flow off should MGT reach hot start abort limits during the start.
NOTE
Refer to Chapter 96 for detailed start system information and troubleshooting.
The following paragraphs provide additional information on automatic starts:
1.
To ready the system for an automatic start, the
FADEC MODE switch must be set to AUTO, and the throttle set to the idle position. The start switch is then momentarily positioned to START. Observe START and AUTORELIGHT lights are illuminated before releasing START switch. Throttle modulation of fuel flow is not required.
2.
Although the start sequence is automatic, the pilot is responsible for monitoring the start and taking appropriate action if required. Therefore, it is recommended that both the throttle and start switch are guarded until the start is completed. Do not initiate a start if FADEC related caution panel lights are illuminated unless appropriate maintenance investigation or successful corrective action has been carried out and no “current” faults are shown.
NOTE
After the throttle is set to idle, the momentary contact start switch must be activated within 60 seconds to initiate the start and engage the latching feature. The latching feature of the start will engage when the FADEC ECU senses momentary activation (1 second) of the start switch or upon sensing an N
G
speed of 5%. If a start is attempted following a delay of more than
60 seconds, the FADEC system will not allow the starter to latch following the release of the start switch, and will not introduce fuel if the start switch is held to
START. Therefore, if a delay of more than
60 seconds has occurred, the system must be reset. To reset the system, the throttle must be repositioned to cutoff and then back to idle. In addition, if electrical power is interrupted prior to initiating the start, with the throttle at idle, the throttle must be repositioned to cutoff and then back to idle after power is restored to re-enable the latching feature. A normal automatic start sequence may then commence. To allow for cooler starts and reduce the possibility of reaching hot start abort limits, it is recommended that residual MGT be below
302°F (150°C), when below 10,000 feet Hp or below 149°F (65°C) when above 10,000 feet Hp prior to start. To reduce residual
MGT, a Dry Motoring Run may be performed in accordance with the
BHT-407-FM-1.
3.
Activating the automatic start mode engages the airframe mounted FADEC/start relay, which is then latched by the FADEC ECU until the N
G
speed reaches 50%. The FADEC/start relay places the MGT indicator into the start mode, signals the generator control unit/voltage regulator to inhibit generator output, flashes the shunt field for the duration of the start, and activates the starter relay. The starter relay activates the starter, illuminates the START advisory segment on the caution panel, and activates the igniter relay. The igniter relay activates the engine igniter system and illuminates the AUTO RELIGHT advisory segment of the caution panel. While in start mode, the
MGT indicator is programmed to record start MGT
29 SEP 2008 Rev. 25
76-00-00
Page 11
BHT-407-MM-9
exceedances, should they occur, which differ from normal operational MGT exceedance limits.
NOTE
The following is applicable if helicopter is configured with FADEC Software Version
5.356 or 5.358 (Reversionary Governor).
4.
Once N
G
speed reaches 10% for ambient temperatures of 80°F (26.6°C) or below, or 12% for ambient temperatures above 80°F (26.6°C) the
FADEC system will introduce fuel, detect the lightoff, and smoothly accelerate the engine to idle while limiting MGT if necessary. At inlet temperatures below
-18°C (0°F), the FADEC will increase start acceleration from 2% to 6% N
G
per second. The increase in the acceleration rate will be noticeable during start and improves start performance at colder temperatures and at higher altitudes.
NOTE
The following is applicable if helicopter is configured with FADEC Software Version
5.202.
Once N
G
speed reaches 10% for ambient temperatures of 20°F (-6.7°C) or below, or 12% for ambient temperatures above 20°F (-6.7°C) the
FADEC system will introduce fuel, detect the lightoff, and smoothly accelerate the engine to idle while limiting MGT if necessary. The ECU increments the start-counter to the next number when lightoff is detected. The start acceleration rate is increased from
2% to 6% N
G
per second for inlet temperatures below
0°F (-18°C). The increase in the acceleration rate will be noticeable during start and improves start performance at colder temperatures and at higher altitudes.
5.
Upon reaching an engine N
G
speed of 50%, the
FADEC ECU unlatches the FADEC/start relay, terminating the start sequence.
6.
Above 50% N
G
(FADEC Software Versions 5.202
and 5.356) or 55% N
G
(FADEC Software Version
5.358), the FADEC ECU carries out a self-test of the auto relight system and continues to energize the igniter relay until 60 ±1% N
G
. During this time the
AUTO RELIGHT light will be illuminated. The engine will continue to accelerate until reaching a stabilized idle of 63 ±1% N
G
.
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Page 12 Rev. 25 29 SEP 2008
7.
The FADEC system also incorporates “Hot Start
Abort Logic”, up to 50% N
G
during start. This feature will cut off fuel flow to the engine fuel nozzle if any of the following conditions occur:
a. Start MGT exceeds 1550°F (843°C) (FADEC
Software Versions 5.202 and 5.356) or 1625°F
(885°C) (FADEC Software Version 5.358), at pressure altitudes less than 10,000 feet and if ECU determined residual MGT was less than 180°F (82.2°C) at initiation (5% N
G
) of start.
b. Start MGT exceeds 1675°F (912°C) (FADEC
Software Version 5.202) or 1700°F (927°C) (FADEC
Software Versions 5.356 and 5.358), at pressure altitudes greater than 10,000 feet or if ECU determined residual MGT was greater than 180°F
(82.2°C) at initiation (5% N
G
) of start.
c. Voltage to FADEC ECU drops below 10.3
VDC. As a significant momentary voltage drop occurs at initiation of the start, ensuring a battery voltage of
24 VDC or above prior to start, in conjunction with appropriate battery maintenance, will reduce the possibility of voltage dropping to 10.3 VDC.
8.
If a FADEC aborted start occurs or the pilot manually initiates a start abort by positioning the throttle to cut off, the FADEC is designed to automatically keep the starter engaged for up to 60 seconds from initiation of the start, to reduce MGT to
302°F (150°C). Once 302°F (150°C) MGT is obtained, the starter will disengage.
NOTE
Momentarily positioning the start switch to
DISENG will only deactivate the FADEC/ start relay which in turn disables the starter and igniter circuits. In the event deactivation of the starter and igniter circuits occurs after engine light-off, but below 50% N
G
, the FADEC ECU will either modulate fuel flow to provide a start if N
G speed is sufficient, or cut off fuel flow if
MGT exceeds hot start abort limits due to low N
G
speed. Therefore, positioning the throttle to cutoff is the appropriate method to manually stop the start sequence.
9.
If external power was used to power the start and the battery switch was left in the “OFF” position, it is important to position the battery switch to “ON” prior to removing the external power source (refer to
BHT-407-MM-9
BHT-407-FM-1, Normal Procedures). If all sources of electrical power are removed from the ECU with the engine at idle in AUTO Mode, the start solenoid valve in the HMU will open, causing the engine to decelerate and possibly flame out. If the battery switch is inadvertently left “OFF” and the external power source is removed, do not attempt to reapply power when a decrease in N
G
speed is noted. Throttle should be positioned to cutoff. Reapplication of electrical power could cause an overtemperature condition due to the reduced N
G
speed and reintroduction of fuel by the
FADEC system.
76-7.
START IN AUTO MODE — ALTERNATE
START
For helicopters that operate in approximately 80°F
(26.6°C) and above temperatures, or at high altitudes and experience a hot start abort event, the start procedure listed below is approved and has been demonstrated to overcome this issue in most instances.
For hot and/or high altitude environments, and when prior troubleshooting has not revealed any engine maintenance issues, this procedure can be used to alleviate the possibility of aborted hot starts.
1.
With the collective in the full down position and the throttle in CUTOFF, the pilot can initiate the start by pressing the starter switch. This will energize the starter motor and turn on the ignition exciter (continue to hold starter switch).
2.
Once N
G
has reached approximately 16%, the pilot is to cycle the throttle from CUTOFF to IDLE.
Once throttle is moved to IDLE position, the starter will latch and the starter switch can be released once lightoff is detected. The engine will detect lightoff and smoothly accelerate to ground idle while limiting the
MGT if necessary.
76-8.
START IN MANUAL MODE —
DESCRIPTION
In accordance with the Rolls-Royce 250-C47B
Operation and Maintenance Manual, Publication CSP
21001, Manual Mode starting on the ground is not authorized except for use under emergency conditions or under special permit from the local aviation authority. Refer to the Rolls-Royce 250-C47B
Operation and Maintenance Manual, Publication CSP
21001 to ensure all Manual Mode Operational
Procedures are followed. Automatic hot start abort features are not available in MANUAL mode.
1.
To ready the system for a MANUAL start, the
FADEC MODE switch is set to MAN and the throttle positioned to cutoff. In MANUAL, the FADEC will not latch the FADEC/start relay or control the fuel scheduling during the start. The start switch must be held in the START position until 50% N
G
. When the N
G speed reaches 12 to 15% (refer to Rolls-Royce
250-C47B Operation and Maintenance Manual,
Publication CSP 21001 for appropriate N
G
speed for
OAT), slowly advance the throttle out of cutoff and stop when the engine lights off. Allow the MGT to peak and then increase fuel flow by modulating throttle to maintain MGT within limits. Once the engine has been started, position the throttle to the idle detent and monitor the N
G
speed.
NOTE
The engine idle speed may reduce to the point where the engine out light and horn are activated.
NOTE
If N
G
increases to more than 75% N
G
with throttle positioned in idle detent, maintenance action is required. Refer to the Rolls-Royce 250-C47B Operation and
Maintenance Manual.
Idle detent speed in MANUAL may not stabilize at 63
±1% N
G
.
• If N
G
speed stabilizes below 63 ±1%, adjust throttle to maintain 63 ±1% N
G
.
• If N
G
speed stabilizes between 63 ±1% and
75%, this is acceptable provided N
P
speed does not fall into avoid steady state range of
68.4 to 87.1% N
P
. If this occurs, adjust throttle to avoid 68.4 to 87.1% N
P
.
• If N
G
stabilizes above 75% N
G
, maintenance action is required.
2.
Once a MANUAL mode start has been initiated, it may be terminated at any time by rotating the throttle to cutoff. The pilot should continue motoring the engine until the MGT has stabilized at an acceptable level. Releasing the start switch from the START
29 SEP 2008 Rev. 25
76-00-00
Page 13
BHT-407-MM-9
position will disengage the FADEC/start relay and disable the starter and igniter circuits.
3.
After a successful MANUAL mode start, and idle has been achieved and is stable, perform a momentary switch back to AUTO mode. If engine flameout occurs, subsequent starts/flight are prohibited until the FADEC ECU has been replaced.
4.
Engine acceleration from idle to 100% is achieved by positioning the throttle towards full open.
This is achieved through hydromechanical control of the HMU fuel metering valve.
76-9.
FADEC MANUAL CHECK —
DESCRIPTION
Following start and prior to increasing engine speed and main rotor RPM to 100%, a FADEC MANUAL check is required. The purpose of this check is to ensure the engine responds to throttle movement while in MANUAL mode. Prior to positioning the
FADEC mode switch to MANUAL for the purposes of this check, ensure the throttle is positioned to idle and
N
G
speed is at 63 ±1%.
NOTE
The engine idle speed may reduce to the point where the engine out light and horn are activated.
NOTE
If N
G
increases to more than 75% N
G
with throttle positioned in idle detent, maintenance action is required. Refer to the Rolls-Royce 250-C47B Operation and
Maintenance Manual.
Once the FADEC mode switch is positioned to
MANUAL, the FADEC MANUAL and AUTO RELIGHT lights will illuminate. In addition, the N
G
speed may change from 63 ±1%. An increase or decrease in N
G speed may be noticed. The change in N
G
speed is due to the fact that the FADEC system is not temperature compensated in regard to fuel flow in MANUAL mode.
Engine load and HMU calibration will also play a part in determining if the N
G
speed will change.
Once N
G
is stabilized, slowly increase throttle approximately 5 to 10% N
G
to ensure engine responds and then return throttle to idle position. Position
76-00-00
Page 14 Rev. 25 29 SEP 2008
FADEC Mode switch to AUTO and ensure FADEC
MANUAL and AUTO RELIGHT lights extinguish.
Engine should return to an N
G
speed of 63 ±1%.
76-10. IN-FLIGHT — AUTO MODE
OPERATION
NOTE
To avoid rapid engine acceleration, roll the throttle smoothly and slowly from the idle to the FLY position.
Engine acceleration from idle to 100% N
P
/N
R
in AUTO mode is achieved by smoothly increasing the throttle to full open or to the detented FLY position. As the throttle is positioned from idle (PLA 30° to 40°) to full open, or the detented FLY position, electrical signals are sent to the ECU from the HMU – PLA potentiometer. These signals dictate the amount of authority the ECU has to control maximum fuel flow
(N
G
limiting) based on throttle position, and in turn controls engine N
G
speed. Therefore, as the throttle is increased from idle to the detented FLY position, the fuel flow is electronically increased until 100% N
P is obtained.
/N
R
The ECU has complete control over engine operation to maintain N
R
within Power On limits found in the
BHT-407-FM-1 when the PLA is between 62 and 100°.
The PLA will be approximately 100° when the throttle is positioned to “full open” or approximately 70° PLA when the throttle is positioned to the detented FLY position. Although the FLY position is the appropriate throttle position for flight operations, 100% N
P
/N
R
will be maintained when the PLA is between 62 and 100°.
To maintain the appropriate N
R
speed, the ECU receives engine and airframe inputs, processes them and modulates the HMU stepper motor driven fuel metering valve to achieve desired engine performance.
If required, as may be the case in certain Emergency
Procedures, an alternate means of engine control is also available to the pilot. This can be achieved by manipulating the throttle below 62° PLA. As the throttle is positioned between 40 and 62° PLA, electrical signals are sent to the ECU from the HMU –
PLA potentiometer. These signals dictate the amount of authority the ECU has to control maximum fuel flow
(N
G
limiting), and in turn engine N
G
speed. Therefore, as the throttle is varied between 40 and 62° PLA, the
BHT-407-MM-9
engine N
G
speed can be manipulated to achieve desired engine performance.
To provide additional information on how the FADEC operates, the following paragraphs discuss the NDOT control, Gas Generator (N
G
) governor and Power
Turbine (N
P
) governor.
76-11.
NDOT CONTROL
The basis for this control system is a NDOT controller that regulates the acceleration rate of the gas generator and thereby controls engine power. Each governor and limiter in the control sends a signal to the
NDOT controller requesting that more or less power is output by the engine. These requests, which are in the terms of demand for N
G
acceleration/deceleration, are then compared by the NDOT control.
This comparison consists of examining the current actual rate of acceleration to that demanded by each governor and limiter. Each governor and limiter NDOT is compared to the maximum acceleration schedule and lowest wins to avoid surge. The lowest demanded
NDOT is then selected and a final comparison is done with the engine deceleration limit on a highest wins basis. The winning parameter is then converted into a command to move the fuel flow metering valve causing an increase/decrease in fuel flow and, accordingly, N
G
speed. Fuel flow is limited between a minimum and maximum limit.
This system provides for consistent engine acceleration and deceleration rates regardless of engine condition. The programmed limits are established to avoid compressor stall, turbine overtemperature (1661°F (905°C) in-flight), and combustion blowout.
76-12.
GAS GENERATOR (N
G
) GOVERNOR
In auto mode the pilot's PLA (throttle position) controls the set point for the N
G
governor. This allows the pilot to limit engine power as desired and provides smooth transition from N
G speed N
P
governing at idle to power turbine
governing at 100% rotor speed.
76-13.
POWER TURBINE (N
P
) GOVERNOR
The control governs power turbine speed at 100%.
The control utilizes isochronous speed (constant speed) governing with gains and compensation optimized for the engine installation.
A collective pitch position analog input signal provided by the Collective Pitch Transducer (CPT) provides load anticipation for the N anticipation initiates N
G
P
speed governor. This
acceleration after CPT movement (increase in collective), prior to actual load increase, to reduce rotor droop.
The rotor speed input frequency signal provided to the control by the N
R
monopole pickup enhances autorotation recovery. Using the rotor speed input and collective pitch, the FADEC changes N
G
acceleration/ deceleration to change NP speed to match rotor speed
(N
R
) during the reapplication of rotor load, thus minimizing rotor speed droop/overshoot.
76-14. ENGINE AUTO RELIGHT
76-15.
AUTO MODE — AUTO RELIGHT (N
G
ABOVE 50%)
In AUTO mode, the FADEC is capable of detecting an engine flameout by measuring an N
G
deceleration rate greater than the predetermined flameout boundary rate. If a flameout is detected, the ENGINE OUT warning light and horn will be activated by the FADEC
ECU. Without pilot action, the auto relight sequence is initiated, a fuel flow rate is established, and the ignition system is activated. If a relight is achieved, the FADEC will control the MGT and accelerate the engine back to its commanded speed of operation. The engine out light and warning horn will turn off after a minimum
NDOT (N
G
acceleration speed) or increasing MGT is established.
The automatic auto relight sequence will initiate from detection of flameout until the N
G
50%. Once the N
G
speed decays to
decays below 50% the FADEC will no longer attempt to relight the engine.
76-16.
AUTO MODE — PILOT ASSISTED
IN-FLIGHT RESTART (N
AND 50%)
G
BETWEEN 9.5
In addition to the above mentioned engine relight features, the FADEC system also incorporates specific relight logic, for engine out conditions, when the NG speed is between 9.5 and 50%. Pilot action is required to initiate an in-flight restart at N
G
speeds below 50%.
When appropriate procedures found in the
BHT-407-FM-1 are followed, the in-flight restart logic will introduce fuel scheduling based on the existing N
G speed. Should a relight be achieved, the FADEC will accelerate the engine to an idle speed of 63% N
G
. As
29 SEP 2008 Rev. 25
76-00-00
Page 15
BHT-407-MM-9
the priority of the in-flight relight logic is to help achieve an engine start in an emergency condition, the hot start abort function is disabled.
Therefore, to help reduce the possibility of an overtemperature condition from occurring, procedures found in the BHT-407-FM-1 require that the throttle be initially positioned to the closed position and the start switch be positioned to “START”. Once the starter is assisting to maintain or increase the N
G
speed, the throttle can be positioned to idle and the FADEC will introduce fuel scheduling. Ignition will be provided in conjunction with activation of the starter.
As the in-flight restart logic is designed for N between 9.5 and 50% N
G initiated below 9.5% N
G
G
speeds
, if an in-flight restart is
, normal start logic will be used to introduce fuel based upon 5.202, 5.356, or 5.358
FADEC software parameters ( paragraph 76-6 ). Should a relight occur, the FADEC will accelerate the engine to idle. In addition, hot-start abort logic will be enabled for starts initiated at N
G
speeds below 9.5%.
76-17. MANUAL MODE — RELIGHT
In MANUAL mode, the FADEC controlled auto relight circuit is disabled. In this mode the ignition system has been designed to operate continuously at engine Gas
Generator (N
G
) speeds of 55% or greater to reduce the possibility of flameout.
When the FADEC system is in MANUAL mode, the N
G gauge operates as a trigger device for the engine out horn and light when N
G
drops below 55%. In the event of a power loss to the N
G
indicator while operating in
MANUAL mode, the failure mode will provide
continuous ignition regardless of N
G
speed, but the engine out light and horn will not be activated when
N
G
drops below 55%.
76-18. ENGINE OVERSPEED AND
PROTECTION — DESCRIPTION
N
P
overspeed limiting is available in both the AUTO and MANUAL modes by independent analog circuits integral to the ECU. With FADEC Software Version
5.202, N
G
overspeed limiting is available in the AUTO and MANUAL modes through software control of the independent analog circuits. In the event of a FADEC
FAILURE, it is possible that overspeed protection will not be available. N
G
overspeed protection is not
applicable with FADEC Software Versions 5.356 and
5.358.
76-00-00
Page 16 Rev. 25 29 SEP 2008
The FADEC ECU continuously monitors for N
G
N
P
, N
P
or
versus torque (5.202 FADEC software only) overspeed conditions in both AUTO and MANUAL mode.
The ENGINE OVSPD light will illuminate if the FADEC detects a N
G
overspeed of 110 ±1% (FADEC Software
Version 5.202) or a N
P
overspeed of 118.5 ±1%.
Illumination occurs when the overspeed solenoid valve is activated within the HMU. With FADEC Software
Version 5.202, the ENGINE OVSPD light will also illuminate when N maximum continuous limit (102.1% N
P to 108.6% N
P
P
versus TORQUE is above the
at 100% torque
at 0% torque). With FADEC Software
Version 5.356 or 5.358, the ENGINE OVSPD light will also illuminate when N
P
is above the maximum continuous limit of 102.1% for 2.5 seconds or immediately when N
P
reaches or exceeds 107.3%.
For FADEC Software Version 5.358, the FADEC
FAULT light will illuminate for the remainder of the flight and will trigger the ENGINE OVSP light to be illuminated on shutdown if N
P
exceeded a maintenance limit per the following conditions:
CONDITION
1
2
3
NP
THRESHOLD
> 102.1%
> 107.3%
> 113.3%
TIME
LIMIT
15 s
0 s
0 s
NO. OF
EVENTS
> 0
> 5
> 0
The light will also momentarily illuminate during the overspeed system test when the overspeed solenoid valve closes ( Figure 76-3 ).
As applicable to FADEC Software Version 5.202, if 15 seconds is exceeded with N
P
versus torque above a line between 102.1% N
P
at 100% torque to 108.6% N
P at 0% torque (line 1) and a line between 104.2% N
P
100% torque to 113.3% N
at
P
at 0% torque (line 2), ECU recording of an overspeed will occur. The ECU will also record an overspeed anytime N
P exceeds a line from 104.2% N
P
113.3% N
P
versus torque
at 100% torque to
at 0% torque (line 2) ( Figure 76-3 ).
As applicable to FADEC Software Versions 5.356 and
5.358, the ECU will record an N
P
overspeed when operating for greater than 15 seconds between
102.1% N
P
and 107.3% N
P
, anytime operations occur between 107.3% N
113.3% N
P
P
and 113.3% N
P
, or anytime
is exceeded. The FADEC also records the
number of overspeed occurrences between 107.3%
N
P
and 113.3% N
P
.
BHT-407-MM-9
120
115
113.3
110
108.6
105
LINE 2
LINE 1
104.2% N
P
102.1% N
P
100
95
100%
90
0 8.9
17.9
26.8
35.7
44.6
53.6
62.4
71.4
TORQUE % (BELL HELICOPTER % TORQUE)
80.3
89.2
98.2
107.1
ENGINE OVERSPEED LIGHT ACTIVATION:
ECU OVERSPEED RECORDING:
NOTE
This information is only applicable with 5.202 FADEC software installed.
Figure 76-3. Engine Overspeed Light and ECU N
P
Recording Activation Table for
FADEC Software Version 5.202
407MM_76_0003
29 SEP 2008 Rev. 25
76-00-00
Page 17
BHT-407-MM-9
Figure 76-4. Example of Engine History Data — Maintenance
76-00-00
Page 18 Rev. 25 29 SEP 2008
BHT-407-MM-9
If the ENGINE OVSPD light is activated during engine operation, due to an exceedance, it will be recorded by the ECU, and the pilot will be provided with a maintenance advisory on shutdown in the form of a
FADEC DEGRADED light. The FADEC DEGRADED light will illuminate when N
G
speed decays below
9.5%. If the pilot fails to recognize illumination of the
FADEC DEGRADED light on shutdown, it will be illuminated the next time electrical power is applied following the FADEC system self-test. When the
FADEC DEGRADED light is illuminated as a maintenance advisory, maintenance investigation is required prior to further flight. Peak values of exceedances are located on the Engine History Data page of the Maintenance Terminal ( Figure 76-4 ).
To determine if maintenance action is required following a recorded overspeed, refer to Chapter 5 of this manual and refer to the overspeed limits in the
Rolls-Royce 250-C47B Operation and Maintenance
Manual, Publication CSP 21001.
NOTE
As applicable to 5.202 FADEC software, if
N
P
overspeed was recorded as
NpQNppkExLm (exceedance limit), the duration of the overspeed was less than 15 seconds. If N
P
overspeed was recorded as
NpQNppkRnLm (run limit), the duration of the overspeed was greater than 15 seconds.
76-19. N
P
OVERSPEED
When the engine reaches 118.5 ±1% N
P
, the ENGINE
OVSPD warning light will illuminate and overspeed limiting will occur. The analog overspeed limiting feature will activate the overspeed solenoid valve, which reduces fuel to the engine to a minimum flow condition (sub-idle value of 34 to 45 pph). The minimum fuel flow increases the likelihood of the engine remaining running and recovering from the overspeed. Once the N
P
speed drops to 112.5 ±2%, the overspeed solenoid valve will be deactivated and fuel flow will return to its previously commanded value.
In the event the overspeed cannot be controlled after fuel flow is reintroduced, the overspeed limiting feature will control the overspeed between the activation trip point of 118.5 ±1% N
P
112.5 ±2% N
P
and the deactivation point of
. If this occurs, attempt to control engine and rotor speed with throttle and collective. Refer to the ENGINE OVERSPEED procedure in the
BHT-407-FM-1.
In addition to the ENGINE OVSP warning light illuminating during the actual overspeed conditions,
FADEC Software Version 5.358 will also trigger the
FADEC FAULT light to illuminate for the remainder of the flight and will trigger the ENGINE OVSP light to be illuminated on shutdown if N
P maintenance limit ( paragraph 76-18 ).
exceeded a
Additional information on the N
P follows:
Overspeed System
1.
The overspeed limit control design incorporates four analog speed sensing circuits driven by two N
P speed signals. Two of the sensing circuits are independently capable of sourcing current to the overspeed solenoid valve in the HMU, two are independently capable of providing a ground to the overspeed solenoid valve. False trips are unlikely since a false trip requires that two independent sensing circuits fail. Additionally, the availability of the overspeed protection is high since up to two sensing circuit failures can occur without affecting capability.
The Power Turbine (N
P
) overspeed limiter operates while the ECU is in either the automatic or manual
mode. Functionality of the overspeed system is
evaluated by three methods, power up check, continuous checks, and a pilot-initiated overspeed test at engine shutdown.
2.
The power supply for the power turbine overspeed limiting circuits is redundant to the power supply for the remaining ECU circuits and is sourced by both the helicopter power bus and the engine mounted PMA.
76-20.
POWER UP FUNCTIONAL CHECK
The power up check occurs when the ECU is first turned on. This check ensures electrical continuity of the overspeed circuit and the ability of the ECU to power the overspeed solenoid. This test is performed by turning on each of the overspeed solenoid drivers and measuring the voltage across and current draw through the overspeed solenoid valve. The measured voltage and current are then compared to limits.
76-21.
CONTINUOUS FUNCTIONAL CHECK
Continuous checks occur during normal engine operation. These checks monitor the functionality of
29 SEP 2008 Rev. 25
76-00-00
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BHT-407-MM-9
the N
P
speed signals that supply the overspeed system. The two N
P
speed signals that supply the overspeed system are continuously compared for differences, and should a difference become larger than a predetermined limit, a fault is declared.
76-22.
OVERSPEED SYSTEM FAILURE
ANNUNCIATION
The operational status of the overspeed system is automatically and continually monitored by the ECU circuits to detect latent failures that could result in false trips or nonoperation should one or more additional failures occur. Should a failure be detected by the automatic test or the continuous checks, a fault will be declared.
76-23.
OVERSPEED SYSTEM SHUTDOWN
CHECK
Functionality of the overspeed system is checked during FADEC power up and thereafter continuously by the ECU. Operation of the overspeed solenoid is check periodically by the pilot through the use of the
OVERSPEED SHUTDOWN test procedure. The
OVERSPEED SHUTDOWN test procedure will shut down the engine only if collective pitch is below 10%, throttle position is at idle, N
G and N
P
is between 60 and 66%
is less than 75%. The OVERSPEED test button must be pressed and held for a minimum of 1.0
second but not more than 10.0 seconds. Once the test button is released, the OVERSPEED test is completed as follows. The FADEC ECU signals the overspeed solenoid valve to close and the ENGINE OVSPD light to come on. Once the FADEC ECU senses an N
G decrease greater than 0.5%, the overspeed solenoid valve is opened, the ENGINE OVSPD light goes off, and the engine is shut down by FADEC ECU activation of the hot start abort feature.
If the overspeed test is unsuccessful, the engine will continue to operate at idle power, the FADEC FAULT caution light will illuminate, and a normal shutdown procedure must be carried out.
76-24. N
G
OVERSPEED
NOTE
The following information is only applicable to FADEC Software Version 5.202.
In auto mode, a software implemented overspeed system is provided. Should the software detect a N
G
76-00-00
Page 20 Rev. 25 29 SEP 2008
overspeed, the protection feature will be activated. In addition, if the FADEC ECU has not failed, N
G overspeed protection will be available in MANUAL mode.
When the engine reaches 110 ±1% N
G
, the ENGINE
OVSPD warning light will illuminate and overspeed limiting will occur. The software controlled overspeed limiting feature will activate the overspeed solenoid valve, which reduces fuel to the engine to a minimum flow condition (sub-idle value of 34 to 45 pph). The minimum fuel flow increases the likelihood of the engine remaining running and recovering from the overspeed. Once the N
G
speed drops to 107 ±1%, the overspeed solenoid valve will be deactivated and fuel flow will return to its previously commanded value.
In the event the overspeed cannot be controlled after fuel flow is reintroduced, the overspeed limiting feature will control the overspeed between the activation point of 110 ±1% N
G
N
G
and the deactivation point of 107 ±1%
. If this occurs, attempt to control engine and rotor speed with throttle and collective. Refer to the
ENGINE OVERSPEED procedure in the
BHT-407-FM-1.
76-25. ENGINE SHUTDOWN
Pilot control of engine speed from 100% N
P
/N
R
to idle in AUTO mode is controlled electrically through throttle movement. As the throttle is positioned from full open to idle, electrical signals are sent to the ECU from the
HMU – PLA potentiometer. These signals dictate the amount of authority the ECU has to control maximum fuel flow (N
G
limiting), and in turn, engine speed.
Therefore, as throttle is decreased, the maximum fuel flow that can be delivered to the engine is electrically reduced by positioning the fuel metering valve to control engine N
G
speed/power.
In the unlikely event that a system fault occurs that does not allow a reduction in engine speed by positioning the throttle to idle, complete the 2-minute cool down at 100% flat pitch. After the 2-minute cool down, the engine is to be shut down by rolling the throttle to CUT–OFF.
Pilot control of engine speed from 100% N
P
/N
R
to N
G idle in MANUAL mode is controlled hydromechanically through throttle movement. Idle speed in MANUAL mode may not stabilize at 63 ±1% N
G maintain idle speed at 63 ±1% N
G
. If this occurs,
with throttle.
BHT-407-MM-9
Following the appropriate cool down period at idle, the engine may be shut down in either the AUTO or
MANUAL mode by positioning the throttle to cutoff.
As applicable to FADEC Software Version 5.202, do not reposition the throttle out of cutoff unless N
G
has decayed to zero. If the throttle is positioned out of cutoff prior to the N
G
speed decreasing through 9.5%, the FADEC “In-flight” restart logic ( paragraph 76-16 ) will introduce fuel and activate the igniter. This can cause a relight and possible overtemperature condition. If relight occurs, the pilot must immediately position the throttle to cutoff and activate the starter.
FADEC Software Versions 5.356 and 5.358 will not introduce fuel or activate the igniter under these conditions unless the start switch is activated.
Additionally, if shutting down in auto mode, the pilot must also allow N
G
speed to decay to 0% prior to positioning the battery switch to off. The reason it is important to wait for the N
G
to decay to zero prior to removing battery power is because of the auto/manual solenoid in the HMU. When you remove electrical power, this solenoid opens and allows fuel to flow to the manual mode pistons. As the HMU fuel pump is very capable of providing high pressure fuel at very low N
G
0% N
speeds, if battery power is removed prior to
G
, the auto/manual solenoid will open and high pressure fuel will flow to the manual pistons and extend them from their parked position. This in turn, will set the open metering valve fault (openMvFlg), and cause the restart fault light to illuminate the next time electrical power is applied. An HMU manual piston parking procedure will then be required per paragraph
76-26 or as described in the Rolls-Royce 250-C47B
Operation and Maintenance Manual, Publication CSP
21001.
76-26.
HMU MANUAL PISTON PARKING
PROCEDURE
Starting with the HMU manual mode pistons in the wrong position may result in a hot start of the engine.
The reason you may get a hot start when the pistons are not parked is because the extended pistons may restrict the movement of the metering valve. Under normal conditions, the metering valve is positioned to a start position of approximately 45 to 50 pph until lightoff occurs. Once lightoff is detected, the system cuts back fuel flow to maintain the required start
acceleration rate. If the metering valve cannot be
positioned to the lower fuel flow after lightoff because the position of the manual pistons is restricting the
ECU's command of the HMU metering valve, a hot start could occur.
When personnel are (i.e., following maintenance) not certain of the position of the pistons or have received a
Maintenance Mode Advisory that the pistons are out of position, the following procedure will assure the pistons are in the correct position (fully retracted) for engine starting.
1.
Position throttle to cutoff.
2.
Pull igniter circuit breaker.
3.
BATT — ON.
4.
Power up check — Complete.
5.
FADEC Mode switch — MANUAL.
6.
Motor the engine (with throttle in cutoff) for 10 seconds.
7.
Wait for N
G
to decay to 0%.
8.
FADEC Mode switch – AUTO.
9.
Motor the engine (with throttle in cutoff) for an additional 10 seconds.
10. Wait for N
G
to decay to 0%.
11. BATT — OFF.
12. Push in igniter circuit breaker.
76-27.
FADEC SYSTEM FAULTS
CAUTION
B E L L H E L I C O PT E R R E Q U I R E S
MAINTENANCE ACTION, PRIOR TO
FLIGHT, WHEN A FADEC RELATED
LIGHT IS ILLUMINATED.
There are eight lights in the caution/warning/advisory panel that are controlled by the FADEC: FADEC FAIL,
FADEC MANUAL, FADEC DEGRADED, FADEC
FAULT, RESTART FAULT, ENGINE OVSPD, ENGINE
OUT, and AUTO RELIGHT ( Figure 76-5
29 SEP 2008 Rev. 25
).
76-00-00
Page 21
BHT-407-MM-9
FADEC FAULT
LIGHT
CAUTION/WARNING
PANEL
RESTART FAULT
LIGHT
AUTO RELIGHT
LIGHT
START LIGHT
FADEC FAIL
LIGHT
FADEC MANUAL
LIGHT
FADEC DEGRADED
LIGHT
ENGINE OVRSPD
LIGHT
ENGINE OUT
LIGHT
C/W LT TEST
SWITCH
OVSPD TEST
SWITCH
FADEC MODE
SWITCH
407MM_76_0005_c01
Figure 76-5. Instrument Panel — FADEC System Switches, Caution/Warning Panel
76-00-00
Page 22 Rev. 25 29 SEP 2008
BHT-407-MM-9
The FADEC ECU continuously monitors the FADEC system for faults and makes appropriate accommodations to continue operation. Fault codes have been preassigned to those parameters being monitored by the FADEC ECU.
If any failure occurs in the ECU/HMU or in one of the input/output signals that significantly impacts the ECU or control of the HMU, the pilot will be alerted via the
FADEC FAIL warning horn and the FADEC FAIL/
FADEC MANUAL warning lights. With FADEC
Software Version 5.356 or 5.358 installed, the reversionary (backup) governor will be activated under certain fault conditions to eliminate a FADEC FAIL condition. This will allow operations in a degraded mode while remaining in AUTO mode.
If the detected failure does not significantly impair the functioning of the ECU, the pilot will be alerted via a
FADEC DEGRADED, FADEC FAULT, RESTART
FAULT caution light, or combination of, depending on the nature of the fault.
If the fault is minor in nature, it will not be communicated to the cockpit with the engine running.
These faults are identified as maintenance advisory faults and will be displayed during shutdown when the throttle is placed in the cutoff position and N
G
speed decays below 9.5%. This will be in the form of a
FADEC DEGRADED light. The BHT-407-FM-1 provides the appropriate action required by the pilot for each light or light/horn condition.
All FADEC faults have been categorized into five types. The first four relate to in-flight faults and the fifth relates to Maintenance Advisory faults with the engine shut down. Maintenance Advisory faults displayed during shutdown will be discussed under the heading
FADEC SYSTEM FAULTS — ENGINE SHUTDOWN.
76-28.
CATEGORY 1 — FADEC FAILURE
NOTE
The following information is only applicable to Model 407 helicopters S/N 53390 and subsequent and those that have complied with ASB 407-99-31.
With the Direct Reversion to Manual System, faults that require pilot action and transition to the MANUAL mode will be displayed immediately when detected by the ECU. The FADEC FAIL horn (chime tone) will activate in conjunction with the FADEC FAIL and
FADEC MANUAL warning lights.
In addition, with FADEC Software Version 5.356 or
5.358 installed, the RESTART FAULT light will also be displayed with a FADEC FAILURE.
The Direct Reversion to Manual system ensures all
FADEC failures revert directly to manual. Fail Fixed failures do not exist with this system. In addition, the system incorporates a throttle that is detented at the
90% bezel or FLY position.
The main intent of the Direct to Manual system is to simplify pilot procedures in the event of a FADEC failure. This is accomplished by allowing the pilot to keep his hands on the controls during a FADEC failure and enable an increase or decrease in throttle from the detented FLY position as required. The pilot will only have to remove his hand from the collective to press the FADEC MODE switch and silence the horn when firmly established in MANUAL mode.
It is the pilots responsibility to control the helicopter during the transition to MANUAL mode.
In this situation, reversion to MANUAL mode will occur independent of the position of the FADEC MODE switch on the instrument panel. The fuel flow will initially be failed fixed and reversion to the MANUAL mode will begin immediately.
As the FADEC SYSTEM has initiated the transition to
MANUAL mode, the pilot must be aware that an increase or decrease in N
P
/N
R
will most likely occur within 7 seconds following a direct failure to MANUAL.
If this occurs, collective will have to be used to control
RPM.
76-29.
AUTO TO MANUAL MODE TRANSITION
NOTE
The following information is only applicable to Model 407 helicopters S/N 53390 and subsequent and those that have complied with ASB 407-99-31.
The objective of transition to MANUAL mode is to provide the pilot with throttle control of the fuel metering valve and in turn, control engine speed.
MANUAL mode allows the pilot to control N
P
/N
R
with coordinated control of the collective and throttle.
29 SEP 2008 Rev. 25
76-00-00
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BHT-407-MM-9
The following procedural steps are required:
1.
Throttle – If time permits, the system can be manipulated for a smoother transition from AUTO to
MANUAL mode. This can be accomplished by matching the throttle and bezel to the actual indicated
N
G
speed. This procedure permits the smoother transition to MANUAL. This is due to the fact that the actual N
G
speed and fuel metering valve position prior to switching to MANUAL will be very close to that following the transition to MANUAL mode, resulting in little, if any, RPM change.
2.
Control rotor (N
R collective, only.
) and engine (N
P
) RPM with the
a. It is most important to ensure that N
P
/N
R
is monitored and properly controlled during and following the transition to MANUAL. N
P
/N
R
may begin to increase or decrease very rapidly, within 7 seconds following a failure direct to MANUAL. This will require collective inputs to control RPM. The Model 407 rotor system is very responsive to collective inputs and can be controlled by the pilot should a N
P
/N underspeed tendency arise.
R
overspeed/
b. To complete the transition from AUTO to
MANUAL, it will normally take approximately 2 to 7 seconds. The transition will not be completed until the fuel metering valve in the HMU can be manually controlled by the pilot through use of the throttle on the collective. Therefore, use of throttle to control N
R
/N
P will be ineffective until the transition to MANUAL mode is complete.
c. There are two pistons within the HMU, a
Manual Load Piston (slow piston) and a PLA Follower
Piston (fast piston), which must hydromechanically extend to contact opposite sides of the fuel metering valve shaft lever. The two pistons move at different rates toward the fuel metering valve lever. It takes approximately 2.0 seconds for both pistons to make contact with the fuel metering valve lever following a transition from an initial condition of low fuel flow.
Similarly, up to 7 seconds may be required for the two pistons to make contact following a transition from an initial condition of high fuel flow. Refer to Figure 76-6 ,
Figure 76-7 , and Figure 76-8 for additional information on AUTO to MANUAL transitions.
76-00-00
Page 24 Rev. 25 29 SEP 2008
d. An increase in N
P
/N
R
speed may be experienced while in transition to MANUAL from a condition of low to higher fuel flow or high fuel flow to a higher fuel flow. This will be seen if the throttle to bezel selection made by the pilot in step 1 of the procedure is higher than the actual N
G
speed at the time of the
FADEC FAILURE condition. This will occur during the period when the HMU Manual Load Piston (slow piston) engages the fuel metering valve lever and moves it to a more open position until the PLA
Follower Piston (fast piston) is contacted.
e. Inversely, a decrease in N
P
/N
R
speed may be experienced during the transition to MANUAL from a condition of low to lower fuel flow or high fuel flow to a lower fuel flow. This will be seen if the throttle to bezel selection made by the pilot in step 1 of the procedure is lower than the actual N
G
speed at the time of the
FADEC FAILURE condition. This will occur during the period when the PLA Follower Piston (fast piston)
engages the fuel metering valve lever and moves it to
a more closed position as dictated by throttle to bezel position.
f. The approximate time to detect a power change during the transition to MANUAL is summarized in Table 76-1 . In simpler terms, there will be a time delay and possible change in engine power while the system transitions to manual. The length of the delay and degree of power change during the transitions depend on engine power at the time of the transition. As stated previously, the degree of power change can be minimized by matching the throttle and
bezel to the actual indicated N
G
speed in step one of the FADEC FAILURE procedure. This permits the smoother transition to MANUAL due to the fact that the actual N
G
speed and fuel metering valve position prior to switching to MANUAL will be very close to that following the transition to MANUAL mode. This will result in little, if any, power/RPM change.
g. Once both pistons contact the lever on the fuel metering valve shaft, the transition to MANUAL Mode
will be complete. The pilot will have slew rate limited
control of the fuel metering valve via throttle position without any delay.
h. Throttle may now be used, in conjunction with collective, to maintain rotor and engine RPM within 95 to 100%.
BHT-407-MM-9
1.0 SECOND
2.0 SECONDS
HIGH
PRESSURE
FUEL
1 AUTO/MANUAL
CHANGEOVER
SOLENOID VALVE
VALVE
SHOWN
OPEN
(DE-ENERGIZED)
POWER
LEVER
LOW
PRESSURE
FUEL
MANUAL LOAD
PISTON (SLOW) 2 PLA FOLLOWER
PISTON (FAST)
METERING
VALVE
LEVER
MIN FUEL
FLOW
MAX FUEL
FLOW
VARIABLE
ORIFICE
PARTIAL CROSS SECTION OF HMU
407MM_76_0006
Figure 76-6. Auto to Manual Transition at Low Fuel Flow (Sheet 1 of 2)
29 SEP 2008 Rev. 25
76-00-00
Page 25
BHT-407-MM-9
AUTO TO MANUAL TRANSITION AT LOW FUEL FLOW
INITIAL CONDITION:
Auto mode, low engine fuel flow, throttle at "FLY'' position.
After FADEC mode switch manual selection or initiation of direct reversion to manual:
Auto/manual changeover solenoid valve is de-energized, allowing high pressure fuel to manual load piston and PLA follower piston.
Manual load piston "slowly'' extends. Engages metering valve lever in approximately 2.0 seconds and begins to drive it to meet PLA follower piston.
Concurrently, PLA follower piston "rapidly'' extends in approximately 1.0 second to a position that is a function of throttle (PLA) position.
itself very close to the actual position of the metering valve at the time of the transition. This will minimize the fuel flow change during the transition. Positioning the throttle to a bezel setting that during the transition. The increase in fuel flow will be caused as the manual load piston engages the metering valve lever and drives it towards the PLA follower piston. Inversely, positioning the produce a decrease in fuel flow during the transition. The decrease in fuel flow will be caused as the PLA follower piston engages the metering valve lever and drives it towards the manual load piston.
After both pistons engage, manual mode is established and no delay exists between throttle
(PLA) movement and fuel flow change.
Slew rate limiting is achieved by hydraulic dynamics.
1
NOTES
Auto/manual changeover solenoid valve normally closed (energized) in auto mode.
Auto/manual changeover solenoid valve is opened (de-energized) for transition to and during manual mode operation. With the valve open, fuel pressure is used to position the manual load piston and PLA follower piston.
2
PLA follower piston is controlled by throttle (PLA) position during transition to manual and when in manual mode. PLA follower piston position is regulated by fuel pressure bleed through variable orifice. This provides the means to increase or decrease fuel flow by altering the position of the fuel metering valve. Manual load piston ensures metering valve lever is held against PLA follower piston.
When in automatic mode, both the PLA follower piston and manual load piston are retracted from the metering valve lever. They are held in the retracted position by fuel pressure when the auto/manual changeover solenoid valve is closed (energized).
Figure 76-6. Auto to Manual Transition at Low Fuel Flow (Sheet 2 of 2)
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407MM_76_0007
BHT-407-MM-9
3 SECONDS
0.5 SECOND
HIGH
PRESSURE
FUEL
1
AUTO/MANUAL
CHANGEOVER
SOLENOID VALVE
VALVE
SHOWN
OPEN
(DE-ENERGIZED)
POWER
LEVER
LOW
PRESSURE
FUEL
MANUAL LOAD
PISTON (SLOW)
2
PLA FOLLOWER
PISTON (FAST)
METERING
VALVE
LEVER
MIN FUEL
FLOW
MAX FUEL
FLOW
VARIABLE
ORIFICE
PARTIAL CROSS SECTION OF HMU
Figure 76-7. Auto to Manual Transition at Intermediate Fuel Flow (Sheet 1 of 2)
407MM_76_0008
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AUTO TO MANUAL TRANSITION AT INTERMEDIATE (CRUISE) FUEL FLOW
INITIAL CONDITION:
Auto mode, intermediate engine fuel flow, throttle at "FLY'' position.
After FADEC mode switch manual selection or initiation of direct reversion to manual:
Auto/manual changeover solenoid valve is de-energized allowing high pressure fuel to manual load piston and PLA follower piston.
Manual load piston "slowly'' extends. Engages metering valve lever in approximately 3.0 seconds and begins to drive it to meet PLA follower piston.
Concurrently, PLA follower piston "rapidly'' extends in approximately 0.5 second to a position that is a function of throttle (PLA) position.
itself very close to the actual position of the metering valve at the time of the transition. This will minimize the fuel flow change during the transition. Positioning the throttle to a bezel setting that during the transition. The increase in fuel flow will be caused as the manual load piston engages the metering valve lever and drives it towards the PLA follower piston. Inversely, positioning the produce a decrease in fuel flow during the transition. The decrease in fuel flow will be caused as the PLA follower piston engages the metering valve lever and drives it towards the manual load piston.
After both pistons engage, manual mode is established and no delay exists between throttle
(PLA) movement and fuel flow change.
Slew rate limiting is achieved by hydraulic dynamics.
1
NOTES
Auto/manual changeover solenoid valve normally closed (energized) in auto mode.
Auto/manual changeover solenoid valve is opened (de-energized) for transition to and during manual mode operation. With the valve open, fuel pressure is used to position the manual load piston and PLA follower piston.
2
PLA follower piston is controlled by throttle (PLA) position during transition to manual and when in manual mode. PLA follower piston position is regulated by fuel pressure bleed through variable orifice. This provides the means to increase or decrease fuel flow by altering the position of the fuel metering valve. Manual load piston ensures metering valve lever is held against PLA follower piston.
When in automatic mode, both the PLA follower piston and manual load piston are retracted from the metering valve lever. They are held in the retracted position by fuel pressure when the auto/manual changeover solenoid valve is closed (energized).
Figure 76-7. Auto to Manual Transition at Intermediate Fuel Flow (Sheet 2 of 2)
407MM_76_0009
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BHT-407-MM-9
0.1 SECOND
6 TO 7 SECONDS
HIGH
PRESSURE
FUEL
1
AUTO/MANUAL
CHANGEOVER
SOLENOID VALVE
VALVE
SHOWN
OPEN
(DE-ENERGIZED)
POWER
LEVER
LOW
PRESSURE
FUEL
MANUAL LOAD
PISTON (SLOW)
2
PLA FOLLOWER
PISTON (FAST)
METERING
VALVE
LEVER
MIN FUEL
FLOW
MAX FUEL
FLOW
VARIABLE
ORIFICE
PARTIAL CROSS SECTION OF HMU
407MM_76_0010
Figure 76-8. Auto to Manual Transition at High Fuel Flow (Sheet 1 of 2)
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AUTO TO MANUAL TRANSITION AT HIGH FUEL FLOW
INITIAL CONDITION:
Auto mode, high engine fuel flow, throttle at "FLY'' position.
After FADEC mode switch manual selection or initiation of direct reversion to manual:
Auto/manual changeover solenoid valve is de-energized allowing high pressure fuel to manual load piston and PLA follower piston.
Manual load piston "slowly'' extends. Engages metering valve lever in approximately 6 to 7 seconds and begins to drive it to meet PLA follower piston.
Concurrently, PLA follower piston "rapidly'' extends in approximately 0.1 second to a position that is a function of throttle (PLA) position.
close to the actual position of the metering valve at the time of the transition. This will minimize the fuel speed at the time of the transition will produce an increase in fuel flow during the transition. The increase in fuel flow will be caused as the manual load piston engages the metering valve lever and drives it towards the PLA follower piston. Inversely, positioning the throttle to a bezel setting that is lower than decrease in fuel flow will be caused as the PLA follower piston engages the metering valve lever and drives it towards the manual load piston.
After both pistons engage, manual mode is established and no delay exists between throttle (PLA) movement and fuel flow change.
Slew rate limiting is achieved by hydraulic dynamics.
1
NOTES
Auto/manual changeover solenoid valve normally closed (energized) in auto mode.
Auto/manual changeover solenoid valve is opened (de-energized) for transition to and during manual mode operation. With the valve open, fuel pressure is used to position the manual load piston and PLA follower piston.
2 PLA follower piston is controlled by throttle (PLA) position during transition to manual and when in manual mode. PLA follower piston position is regulated by fuel pressure bleed through variable orifice.
This provides the means to increase or decrease fuel flow by altering the position of the fuel metering valve. Manual load piston ensures metering valve lever is held against PLA follower piston.
When in automatic mode, both the PLA follower piston and manual load piston are retracted from the metering valve lever. They are held in the retracted position by fuel pressure when the auto/manual changeover solenoid valve is closed (energized).
Figure 76-8. Auto to Manual Transition at High Fuel Flow (Sheet 2 of 2)
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407MM_76_0011
BHT-407-MM-9
i. Once in MANUAL mode, the pilot will have complete control of N
P
/N
R
by flight control manipulation and the throttle on the collective. The fuel flow slew rate is hydromechanically limited to provide proper responsiveness for helicopter operation. Fuel flow will be a function of the pilot controlled fuel metering valve orifice size. Maximum Continuous
Power will be available for all ambient conditions.
MANUAL mode fuel flow is altitude compensated to allow a consistent horsepower/altitude relationship without throttle adjustment by the pilot. Fuel flow in the
MANUAL mode, however, is not temperature compensated. Because of this, there may be temperatures at which maximum fuel flow in MANUAL will not be sufficient to produce Takeoff Power.
3.
ENGINE POWER AT TIME OF
FAILURE
Low Power
Low Power
High Power
High Power
Table 76-1. Time to Power Change — (DRTM)
DESIRED POWER AS
SELECTED BY THROTTLE
POSITION
Higher Power
Lower Power
Higher Power
Lower Power
APPROX. TIME TO DETECT
POWER CHANGE DURING
TRANSITION TO MANUAL
2.0 Seconds
1.0 Seconds
6.0 to 7.0 Seconds
0.1 Second
NOTE
In the event engine (N
P
) overspeed system is activated at 118.5 ±1% N
P
during transition to or operation in manual mode, control system is designed to keep engine running. Engine may oscillate between
112.5 and 118.5% N
P
until corrective action is taken with throttle and collective. In addition, if the FADEC ECU is operational, it will track HMU operation, perform diagnostics, monitor engine functions and provide overspeed limiting for both N
P
N
G
and
. Surge detection and avoidance will not be available. If an engine surge is encountered, decrease the throttle until the surge condition clears, then slowly increase the throttle to the desired power level.
Rapid power changes should be avoided.
FADEC Mode switch — Depress one time. This will silence FADEC fail warning horn (chime tone).
4.
Land as soon as practical. Applicable maintenance action will be required prior to next flight.
5.
Normal shutdown if possible. If normal shutdown cannot be completed by rolling throttle to closed position, fuel shutoff valve can be positioned to off.
76-30.
CATEGORY 2 — FADEC DEGRADED
FADEC DEGRADED faults represent a loss of some feature of the FADEC system that may cause a degradation in performance. This may result in N
R droop, N
R
lag, or reduced maximum power capability.
These faults will be displayed immediately when detected by the ECU. Operations should be continued in AUTO mode and helicopter is to be flown smoothly and nonaggressively. In conjunction with the FADEC
DEGRADED light, the RESTART FAULT light may also activate under certain fault conditions.
With FADEC Software Version 5.356 or 5.358
installed, the reversionary (backup) governor will be activated under certain fault conditions that will also allow operations in a degraded mode.
Applicable maintenance action will be required prior to next flight.
76-31.
CATEGORY 3 — FADEC FAULT
FADEC FAULT indicates that PMA, MGT, N
P
, or N
G automatic limiting circuit(s) may not be functional. In conjunction with activation of the FADEC FAULT light, the RESTART FAULT light may also activate under certain fault conditions. These faults will be displayed immediately when detected by the ECU. Operations should continue in AUTO mode. If both lights (FADEC
FAULT and RESTART FAULT) are illuminated, this indicates the MGT automatic limiting circuit (1661°F
(905°C) in-flight), may not be functional. In addition,
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BHT-407-MM-9
with FADEC Software Version 5.358, the FADEC
FAULT light will also be illuminated continuously whenever N
P
has exceeded a maintenance limit
( paragraph 76-18 ). The pilot should follow the appropriate procedures as set out in the
BHT-407-FM-1.
Applicable maintenance action will be required prior to next flight.
76-32.
CATEGORY 4 — RESTART FAULT
RESTART FAULT indicates a subsequent automatic engine start may not be possible. The fault does not require immediate action by the pilot and should not affect performance of the helicopter. It is recommended that the pilot plan the landing site accordingly. These faults will be displayed immediately when detected by the ECU and displayed as
RESTART FAULT. Do not attempt a subsequent start until applicable maintenance action has been completed.
If the engine shutdown procedures are not properly followed, the manual mode pistons may begin to engage during the shutdown. The FADEC may then be unable to prevent a hot start on the next start and will indicate a RESTART FAULT to warn the pilot.
Refer to engine shutdown paragraph 76-25 .
76-33.
CATEGORY 5 — MAINTENANCE
ADVISORY, FADEC SYSTEM FAULTS —
ENGINE SHUTDOWN
Maintenance Advisory Faults are those detected by the ECU that are considered minor in nature and are not communicated to the cockpit with the engine running. The FADEC DEGRADED light serves as the maintenance advisory light. This light will be illuminated if any fault or exceedance has been detected during the last engine run or if a current fault exists. This will indicate that maintenance action is required prior to the next flight. Maintenance advisory faults will display during shutdown when the throttle is placed in the cutoff position and the N
G
speed decays below 9.5%. If the pilot misses the maintenance
advisory on shutdown, it will be reilluminated at the
next application of electrical power.
76-34. FADEC FAULTS/EXCEEDANCES —
RECORDING PROCEDURE
Faults and Exceedances can be recorded under the following conditions:
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1.
Engine Operating.
When the engine is operating (i.e., lightoff has been detected), the FADEC will automatically record Faults/
Exceedances as Current, Last Engine Run,
Accumulated, and Time Stamped as they occur. N
G
N
P
, MGT and Torque exceedance values will also be recorded by the FADEC.
,
2.
Engine Not Operating.
When the engine is not operating, but electrical power is applied, the FADEC will only display Current Faults as they occur. Faults will not be recorded.
76-35. FADEC FAULTS/EXCEEDANCES —
CLEARING PROCEDURE
CAUTION
FAULTS/EXCEEDANCES ARE NOT TO
BE ERASED UNLESS APPROPRIATE
MAINTENANCE ACTIONS HAVE BEEN
CARRIED OUT. DO NOT ATTEMPT TO
CLEAR ANY FAULT OR EXCEEDANCE
WHILE ENGINE IS OPERATING.
1.
Current Faults.
Current faults may be cleared by performing a power reset (battery switch OFF/ON). If fault is no longer detected, associated FADEC DEGRADED light will be extinguished.
NOTE
To erase Last Engine Run and
Accumulated Faults/Exceedances with the
Maintenance Terminal, refer to the
Maintenance Terminal User's Guide for operating instructions.
2.
Last Engine Run Faults/Exceedances.
Last Engine Run faults or exceedances may be cleared by performing a successful engine start or with the use of the Maintenance Terminal.
3.
Accumulated Faults/Exceedances.
Accumulated faults or exceedances may only be cleared with the use of the Maintenance Terminal.
Additionally, N
P
exceedance values may only be cleared from engine history data page with the use of the Maintenance Terminal.
BHT-407-MM-9
76-36. RECORDED EXCEEDANCES/
CONVERSION FACTORS
The FADEC monitors and records N
G
, N
P
, MGT, and
Torque exceedances. The helicopter cockpit gauges monitor and record N
G
, MGT, and Torque exceedances. To determine required maintenance actions following a recorded exceedance, Bell
Helicopter and Rolls-Royce agreed that the helicopter cockpit gauges would be used for N
G
, MGT, and
Torque (Q) exceedances and the FADEC would be used for N
P
overspeed exceedances.
To determine maintenance actions, N
G
, MGT, and
Torque (Q) exceedance limits are to be monitored and recorded by either the helicopter cockpit gauges or the pilot. If a discrepancy between the helicopter cockpit gauges or the pilot is noted, the indication of an exceedance by a serviceable helicopter cockpit gauge/indicating circuit is to be used for determining maintenance action. If cockpit gauge or indicating circuit caused exceedance to be recorded, maintenance action in regards to the exceedance is not required. Refer to Chapter 95 for information to download peak exceedance and duration (time) of cockpit gauge recorded exceedances.
If an operator wishes to determine if the FADEC ECU also recorded a N
G
, MGT, or Torque (Q) exceedance, the Maintenance Terminal may be used ( paragraph
76-37 ).
In regards to N
P
(N
2
) overspeed exceedances, these are monitored and recorded by the FADEC. To determine peak N
P
exceedance information, the ECU must be downloaded with the Maintenance Terminal
(Engine History Data screen). To determine the duration of the N
P
exceedance (time), the ECU must be downloaded using the Windows version of the
Maintenance Terminal.
If the Windows version of the Maintenance Terminal is not available, the following information will help to determine the duration of the overspeed. If the N
P overspeed was recorded by the FADEC as
NpQNppkExLm (Exceedance Limit) (5.202 software),
the duration of the overspeed was less than 15
seconds. If the N
P
overspeed was recorded by the
FADEC as NpQNppkRnLm (Run Limit) (5.202
software), the duration of the overspeed was more than 15 seconds.
When the FADEC or Litton instruments record exceedance values, conversion factors may be required to determine the maintenance actions as specified in the applicable sections of the Bell
Helicopter and Rolls-Royce manuals. This is specifically true when the helicopter cockpit gauge records a Torque (Q) exceedance. Although the recorded exceedance value can be applied directly to determine overtorque maintenance actions in accordance with Chapter 5 of this manual, a conversion to Rolls-Royce % torque and foot-pounds
(ft/lbs) is required to determine maintenance actions in accordance with the Rolls-Royce 250-C47B Operation and Maintenance Manual, Publication CSP 21001.
The following will provide the necessary information:
1.
Convert Bell Helicopter cockpit gauge % torque value to Rolls-Royce FADEC % torque as follows:
• Divide Bell Helicopter cockpit gauge % torque by 1.0535 (100% Bell Helicopter cockpit gauge torque = 94.92% Rolls-Royce FADEC torque. Therefore, 1.0535% Bell Helicopter torque = 1.0% Rolls-Royce FADEC torque).
• Example: 115.3% Bell Helicopter cockpit gauge torque ÷ 1.0535 = 109.44%
Rolls-Royce FADEC torque.
2.
Convert Rolls-Royce FADEC % torque to foot-pounds (ft/lbs) as follows:
• Multiply Rolls-Royce FADEC % torque by 5.9
(100% Rolls-Royce FADEC torque = 590 ft/ lbs. Therefore, 1% Rolls-Royce FADEC torque = 5.9 ft/lbs).
• Example: 109.44% Rolls-Royce FADEC torque X 5.9 = 645.70 ft/lbs.
3.
Convert Rolls-Royce FADEC % torque to Bell
Helicopter cockpit gauge % torque as follows:
• Multiply Rolls-Royce FADEC % torque by
1.0535 (100% Rolls-Royce FADEC torque =
105.35% Bell Helicopter cockpit gauge % torque. Therefore, 1% Rolls-Royce FADEC torque = 1.0535% Bell Helicopter torque).
• Example: 119.3% Rolls-Royce FADEC torque
X 1.0535 = 125.68% Bell Helicopter cockpit gauge torque.
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BHT-407-MM-9
76-37. DOWNLOADING N
G
, MGT, AND
TORQUE (Q) EXCEEDANCES
RECORDED BY FADEC (ECU)
To determine if the FADEC ECU has recorded a N
G
MGT, or Torque (Q) exceedance, the Windows version
, of the Maintenance Terminal may be used.
NOTE
Values recorded by the ECU will be of a higher value and longer duration than the values recorded by the cockpit indicators.
This is normal and is due to the fact that the cockpit indicators are dampened. It is also important to note that the times recorded by the FADEC ECU are cumulative (i.e., times are added from one exceedance to the next until cleared). As stated in paragraph
76-36 , the cockpit indicators are to be used to determine maintenance actions following a N
G
, MGT, or Torque (Q) exceedance.
1.
Install FADEC download cable between computer and FADEC ECU maintenance port on left side of lower console.
2.
Apply electrical power to helicopter and open
Maintenance Terminal program (Mterm version 2.00 or subsequent).
3.
From the Maintenance Terminal menu click File and click Connect to ECU.
4.
Click OK when the Component serial number
5.
Activate the user defined real time display by clicking Real Time then clicking User Defined.
NOTE
If user defined screen is full (15 parameters maximum) highlight one of the existing parameters and use the Edit – “Change” feature to alter the highlighted parameter.
You may also highlight and use the Edit –
“Delete” feature to remove existing parameters.
6.
Use Edit menu to add or insert the following exceedance parameters as desired. Case sensitivity is
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not important, but you have to be connected to the
ECU with the Maintenance Terminal. As each exceedance parameter is entered, exceedance values will be immediately displayed on the screen if they have been previously recorded.
NOTE
In specific regards to Torque (Q) exceedances, FADEC recorded torque values are not the same as Cockpit
Instrument recorded values. Refer to paragraph 76-36 for conversion factors.
MGTSRLmTm
NgLmPk
NgLmTm
NgRLmPk
NgRLmTm
QLmPk
QLmTm
QRLmPk
QRLmTm
EXCEEDANCE
PARAMETER
NAME
MGTLmPk
MGTLmTm
MGTRLmPk
MGTRLmTm
MGTSLmPk
MGTSLmTm
MGTSRLmPk
DESCRIPTION
MGT limit exceedance peak value
MGT limit exceedance time
MGT run limit exceedance peak value
MGT run limit exceedance time
MGT start limit exceedance peak value
MGT start limit exceedance time
MGT start run limit exceedance peak value
MGT start run limit exceedance time
N
G
limit exceedance peak value
N
G
limit exceedance time
N
G
run limit exceedance peak value
N
G
run limit exceedance time
Torque limit exceedance peak value
Torque limit exceedance time
Torque run limit exceedance peak value
Torque run limit exceedance time
BHT-407-MM-9
NOTE
Ensure computer is connected to ECU prior to saving data. Do not interrupt helicopter power prior to saving data.
7.
Save data as required.
NOTE
Ensure saved data can be opened with computer prior to proceeding to step 8 .
8.
Once saved, it is recommended that all exceedance data be deleted from the ECU. This will ensure future exceedances may be referenced without possible confusion from previously recorded events.
This may be accomplished as follows:
a. Go to Engine History Page.
b. Write down the existing values for Engine Run
Time (EngRnTm) and Number of Engine Starts
(NumStrt).
c. Use the “Clear all” command and erase all
Engine History Data.
d. Use “Edit” command to re-enter Engine Run
Time (EngRnTm) and Number of Engine Starts
(NumStrt).
76-38. DETERMINATION OF FAULTS/
EXCEEDANCES AND REQUIRED
TROUBLESHOOTING STEPS
The following information is provided to assist maintenance personnel troubleshooting FADEC system faults. In addition to the steps and information provided below, the Fault Isolation Manual contains detailed information on troubleshooting. The Fault
Isolation Manual is the primary source of FADEC troubleshooting information and may be referenced in
Chapter 73-25-04 of the Rolls-Royce 250-C47B
Operation and Maintenance Manual, Publication CSP
21001. It is recommended that operators familiarize themselves with the Fault Isolation Manual.
Faults detected by the FADEC can be displayed via a
FADEC FAIL, FADEC MANUAL, FADEC DEGRADED,
FADEC FAULT, RESTART FAULT, or by a combination of these lights. Faults are also used to identify exceedances. For additional information on exceedances, refer to paragraph 76-36 .
1.
When a FADEC related light has illuminated in-flight or on the ground, maintenance action is required. Complete step a or step b depending on your download capability.
a. The preferred method to determine FADEC faults or exceedances is with the Maintenance
Terminal. The Maintenance Terminal (Windows version) is capable of providing information on Current
Faults, Last Engine Run Faults, Accumulated Faults,
Time Stamped Faults (Fault History screen) and N
P overspeed exceedance information (Engine History screen). Referring to the applicable Maintenance
Terminal Users Guide for operating instructions, download and note all fault codes (maintenance message codes) and/or exceedance data.
b. If a maintenance terminal is not available, identify faults or exceedances using the Maintenance
Mode feature of the FADEC system. This feature
allows operators to determine faults through a
sequence of flashing light displays on the cockpit caution panel. Refer to paragraph 76-39 for procedures on using the Maintenance Mode feature.
In addition, refer to paragraph 76-40 for procedures to determine if faults are current.
2.
Conduct FADEC system troubleshooting in accordance with Fault Isolation Manual. The Fault
Isolation Manual is the primary source of FADEC troubleshooting information and may be referenced in
Chapter 73-25-04 of the Rolls-Royce 250-C47B
Operation and Maintenance Manual, Publication CSP
21001. As multiple systems interface or may indirectly affect the FADEC, additional troubleshooting information may be found within this chapter and
Chapter 96 of this manual as well as various chapters of the Rolls-Royce 250-C47B Operation and
Maintenance Manual, Publication CSP 21001.
The Fault Isolation Manual may be used as follows:
a. If the Maintenance Terminal was used to determine the FADEC fault(s), refer to the
Maintenance Message Code List for software version
5.202, 5.356, or 5.358 in the Fault Isolation Manual.
The intent of the Maintenance Message Code List is to direct you to the appropriate troubleshooting procedure in the Fault Isolation Manual. This may be accomplished by looking up the maintenance
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BHT-407-MM-9
message code(s), determined from step 1 , and referencing the associated “Go To FIM Task”. Locate the referenced FIM Task(s) in the Fault Isolation
Manual and carry out troubleshooting procedures. In addition to troubleshooting, the FIM task will also provide a description of the fault, background information on the fault, possible causes, and verification procedures.
b. If the Maintenance Mode was used to determine the fault(s) through the caution panel flashing display, refer to the Alert/Status Message List in the Fault Isolation Manual for FADEC Software
Version 5.202, 5.356, or 5.358. The Alert/Status
Message List will provide a cross reference between the flashing code(s) displayed and the maintenance message code(s). Once the maintenance message code(s) are determined, refer to the Maintenance
Message Code List in the Fault Isolation Manual. The
intent of the Maintenance Message Code List is to direct you to the appropriate troubleshooting procedure in the Fault Isolation Manual. This may be accomplished by looking up the maintenance message code(s) and referencing the associated “Go
To FIM Task”. Locate the referenced FIM Task(s) in the
Fault Isolation Manual and carry out troubleshooting procedures. In addition to troubleshooting, the FIM task will also provide a description of the fault, background information on the fault, possible causes and verification procedures.
3.
Following corrective action, ensure no Current
Faults are present. In addition, if Maintenance
Terminal is available, erase Exceedances, Last Engine
Run, Accumulated and Time Stamped faults.
Complete step a or step b depending on download capability.
a. If Maintenance Terminal is available, apply electrical power and allow FADEC to complete self-test. Following completion of FADEC self-test, view Maintenance Terminal Fault History and confirm that no Current Faults exist. Erase Exceedances
(Engine History screen), Last Engine Run faults,
Accumulated faults, and Time Stamped faults (Fault
History screen) as required.
b. If a maintenance terminal is not available, confirm Current Faults do not exist. This may be accomplished by positioning battery switch from ON to
OFF to ON (a power reset is required — battery switch
OFF to ON) and allowing FADEC to complete self-test.
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Page 36 Rev. 25 29 SEP 2008
Position throttle to the idle position. If a “current” fault(s) exist, it will be displayed on the caution panel via the FADEC Degraded light.
If a maintenance terminal is not available,
Exceedances, Last Engine Run, Accumulated and
Time Stamped faults can not be erased at this time.
Although Last Engine Run faults will be erased by the
FADEC when lightoff is detected on the first start, erasal of Exceedances, Accumulated faults and Time
Stamped faults will require the use of the Maintenance
Terminal. Make arrangements to clear Exceedances and faults at the next available opportunity. If the faults are not erased and subsequent faults are recorded, the previously recorded faults will complicate the troubleshooting of the subsequent faults.
4.
Perform Check Run procedure in accordance with paragraph 76-41 .
76-39. MAINTENANCE MODE —
PROCEDURE FOR VIEWING FADEC
FAULT CODES USING CAUTION
PANEL FLASHING DISPLAY
The caution panel fault display may be operated as follows:
NOTE
Displayed faults may be “Last Engine Run” or “Current” faults. Following this procedure, refer to paragraph 76-40 to determine if faults are “current.”
1.
Engine must be shut down and the FADEC
MODE switch positioned to MANUAL. Place the collective full down (below 10%) and the throttle in the cutoff position.
NOTE
If the throttle or collective is moved during the above procedure or the FADEC MODE switch is positioned to AUTO, the FADEC
ECU will exit the fault code reporting mode.
2.
Depress and release FADEC ECU maintenance button on the left hand side of the lower pedestal to
enter the fault code reporting mode (
Figure 76-9 ).
BHT-407-MM-9
INSTALLED WITH
KLN 89B GPS KIT
FADEC/ECU
MAINTENANCE
BUTTON
FADEC/ECU
GROUND
MAINTENANCE
CONNECTOR
LEFT SIDE LOWER CONSOLE
407MM_76_0012+
Figure 76-9. FADEC/ECU Maintenance Button and FADEC/ECU Maintenance Terminal Connector
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BHT-407-MM-9
3.
FADEC DEGRADED, FADEC FAULT, and the
RESTART FAULT lights will simultaneously flash five times to indicate that the maintenance mode has been entered by the ECU ( Figure 76-5 ).
4.
Depress and release the FADEC ECU maintenance button. If a fault is present, it will be displayed by a specified number of FADEC
DEGRADED caution panel light segment flashes.
5.
Depress and release FADEC/ECU maintenance button to flash the next fault code.
6.
Steady illumination of the FADEC DEGRADED caution panel light segment indicates that no other faults exist for this light.
7.
Continue to depress and release the FADEC/
ECU maintenance button to step through the FADEC
FAULT and RESTART FAULT caution panel light
segments as above, to reveal existing fault codes.
8.
If no fault code exists for display by the selected caution panel segment, the caution light will illuminate continuously when the FADEC/ECU maintenance button is released.
9.
When interrogation is complete, the next push of the FADEC maintenance button will cause the FADEC
DEGRADED, FADEC FAULT, and RESTART FAULT caution light segments to flash simultaneously five times and then extinguish. This indicates that the
FADEC ECU has exited the maintenance mode.
NOTE
Refer to Table 76-2 for FADEC Software
Version 5.202 and to Table 76-3 for FADEC
Software Versions 5.356 and 5.358
(Reversionary Governor) Fault Code
Displays.
10. To determine fault description and maintenance message code(s) from caution panel flashing display, refer to Table 76-2 or Table 76-3 .
Table 76-2. 250-C47B FADEC Software Version 5.202 Fault Code Display
STATUS MESSAGE FAULT DESCRIPTION MAINTENANCE MESSAGE CODE
FADEC DEGRADED cockpit lamp flashes 1 time
FADEC DEGRADED cockpit lamp flashes 2 times
ECU Failure has occurred
This Status Message is not used.
Verify that message is indicated by lamp.
Np-Q Exceedance
AD12bitFlt, AD8bitFlt, ECUOTFlt,
GainFlt, HLRfFLt, OffsFlt, PROMFlt,
PW10Flt,
RAMFlt, V15Flt, V5Flt, WDTFlt,
CJCFlt, OrDiodeFlt, BacCompFlt,
EEPROMFlt, ForCompFlt, P1Flt,
SWIntFlt, TestCelFlt, UARTFlt,
UUIntFlt, WDTOutFlt, OSVFlt,
NpOSFlt, OR28Flt, WDTTimeOut
Not Used
NpQExLmAdv FADEC DEGRADED cockpit lamp flashes 3 times
FADEC DEGRADED cockpit lamp flashes 4 times
FADEC DEGRADED cockpit lamp flashes 5 times
This Status Message is not used.
Verify that message is indicated by lamp.
MGT Indication Failure
Not Used
MGTFlt
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BHT-407-MM-9
Table 76-2. 250-C47B FADEC Software Version 5.202 Fault Code Display (Cont)
STATUS MESSAGE FAULT DESCRIPTION MAINTENANCE MESSAGE CODE
FADEC DEGRADED cockpit lamp flashes 6 times
FADEC DEGRADED cockpit lamp flashes 7 times
FADEC DEGRADED cockpit lamp flashes 8 times
FADEC DEGRADED cockpit lamp flashes 9 times
FADEC DEGRADED cockpit lamp flashes 10 times
FADEC DEGRADED cockpit lamp flashes 11 times
FADEC DEGRADED cockpit lamp flashes 12 times
FADEC FAULT cockpit lamp flashes 1 time
FADEC FAULT cockpit lamp flashes 2 times
This Status Message is not used.
Verify that message is indicated by lamp.
Failure to Control HMU – Auto/
Manual Solenoid
CIT Temperature Indication
Failure
Metering Valve not in Start
Position
Starter Relay Interface
Nr Sensor – Rotor Decay
Anticipation
Incorrect Overspeed Test Switch
Indication
HMU – Failure to Control Fuel
Flow
This Status Message is not used.
Verify that message is indicated by lamp.
Np-Q Run Limit Advisory
Not Used
AMSolFlt
T1AFlt, T1BFlt, or T1ABFlt
OpenMvFlg
StrFlt
NrFlt
OSTstSwFlt
WfLimFlag
Not Used
NpQRnLmAdv FADEC FAULT cockpit lamp flashes 3 times
FADEC FAULT cockpit lamp flashes 4 times
FADEC FAULT cockpit lamp flashes 5 times
FADEC FAULT cockpit lamp flashes 6 times
FADEC FAULT cockpit lamp flashes 7 times
FADEC FAULT cockpit lamp flashes 8 times
FADEC FAULT cockpit lamp flashes 9 times
FADEC FAULT cockpit lamp flashes 10 times
Failure in HMU (Metering Valve position reading)
Collective Pitch Potentiometer
Indication Failure
Failure to control HMU (Stepper motor)
This Status Message is not used.
Verify that message is indicated by lamp.
Failure to control HMU
(Overspeed Solenoid)
Engine Surge Event
Ng Speed Indication Failure
WfMvFlt or WfStFlt
CPFlt
StepCntFlt, SMFlt, or AMSolFlt
Not Used
OSFlt
SgFlag
Ng1Flt, Ng2Flt, or Ng12Flt
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BHT-407-MM-9
Table 76-2. 250-C47B FADEC Software Version 5.202 Fault Code Display (Cont)
STATUS MESSAGE FAULT DESCRIPTION MAINTENANCE MESSAGE CODE
FADEC FAULT cockpit lamp flashes 11 times
FADEC FAULT cockpit lamp flashes 12 times
RESTART FAULT cockpit lamp flashes 1 time
Airframe Power Supply Failure
Engine Overspeed
AF28Flt
OSFlag
RESTART FAULT cockpit lamp flashes 2 times
RESTART FAULT cockpit lamp flashes 3 times
RESTART FAULT cockpit lamp flashes 4 times
RESTART FAULT cockpit lamp flashes 5 times
RESTART FAULT cockpit lamp flashes 6 times
RESTART FAULT cockpit lamp flashes 7 times
RESTART FAULT cockpit lamp flashes 8 times
This Status Message is not used.
Verify that message is indicated by lamp.
TMOP Sensor – Torque Indication
Failure
This Status Message is not used.
Verify that message is indicated by lamp.
Not Used
QFlt
Not Used
This Status Message is not used.
Verify that message is indicated by lamp.
Failure in HMU (PLA
Potentiometer) Position Reading
PMA Power Supply Failure
Not Used
PLA1Flt, PLA2Flt, PLA12Flt or
PLARfFlt
Al28Flt
Failure to control HMU (Hot Start
Abort Solenoid)
This Status Message is not used.
Verify that message is indicated by lamp.
Failure to control Ignition Relay
Interface
Np Speed Indication Failure
StSFlt or StSIFlt
Not Used
IgnFlt or IgnIFlt
Np1Flt, Np2Flt, or Np12Flt
RESTART FAULT cockpit lamp flashes 9 times
RESTART FAULT cockpit lamp flashes 10 times
RESTART FAULT cockpit lamp flashes 11 times
RESTART FAULT cockpit lamp flashes 12 times
FADEC DEGRADED, FADEC
FAULT, and RESTART FAULT cockpit lamps on steady
Incorrect Auto/Manual Switch
Indication
Quiet Mode Switch Fault
All faults have been displayed
AMSwFlt
QMSwFlt
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BHT-407-MM-9
Table 76-3. 250-C47B FADEC Software Versions 5.356 and 5.358 (Reversionary Governor)
Fault Code Display
STATUS MESSAGE FAULT DESCRIPTION MAINTENANCE MESSAGE CODE
FADEC DEGRADED cockpit lamp flashes 1 time
Primary Governor Failed
FADEC DEGRADED cockpit lamp flashes 2 times
Reversionary Governor Failed
Np Exceedance
AD12bitFlt, AD8bitFlt, ECUOTFlt,
GainFlt, HLRfFLt, OffsFlt, PROMFlt,
PW10Flt, RAMFlt, V15Flt, V5Flt,
WDTFlt, CJCFlt, OrDiodeFlt,
BacCompFlt, EEPROMFlt,
ForCompFlt, P1Flt, SWIntFlt,
TestCelFlt, UARTFlt, UUIntFlt, OSVFlt,
NpOSFlt, OR28Flt, WDTTimeOut,
ARINCFlt, ARINCHWFlt, SWCfgFlt,
RGSDFlt,QRawFlt, T1BRawFlt,
ESWRGFlt, ESW2RGFlt, ESW3RGFlt,
ESW4RGFlt, ESW5RGFlt,
SWConfigRGFlt, PwrRstFlt,
RGSelSwFlt or NDOTWRCdRGFlt
AD10bitFltRG, PROMFltRG,
RAMFltRG, RGOTFltRG,
SWConfigFltRG, V10FltRG, V15nFltRG,
V15pFltRG, V5qFltRG,
WDTTimeOutRG, ARINCHdFltRG,
ARINCFltRG, ARINCHWFltRG,
BacCompFltRG, ForCompFltRG,
Or28FltRG, OrDiodeFltRG,
PW10LoFltRG, RGTempFltRG,
SPITempFltRG, UARTFltRG,
WDTFltRG, PGSDHdFltRG,
PGSDFltRG, WfCorrPGHdFltRG,
WfCorrPGFltRG, EngRnCtPGFltRG,
EngRnTmPGFltRG, ESWPGFltRG,
NpIncPGFltRG, Np2RawPGFltRG,
P1RawPGFltRG, T1ARawPGFltRG,
SwPwrFltRG
NpLmTOut FADEC DEGRADED cockpit lamp flashes 3 times
FADEC DEGRADED cockpit lamp flashes 4 times
Reversionary Governor did not
Govern When Primary Governor
Failed
MGT Indication Failure
ECUGovFltRG
MGTFlt FADEC DEGRADED cockpit lamp flashes 5 times
FADEC DEGRADED cockpit lamp flashes 6 times
FADEC DEGRADED cockpit lamp flashes 7 times
This Status Message is not used.
Verify that message is indicated by lamp.
Failure to control Auto/Manual
Solenoid
Not Used
AMSolFlt, AMSolFltRG
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BHT-407-MM-9
Table 76-3. 250-C47B FADEC Software Versions 5.356 and 5.358 (Reversionary Governor)
Fault Code Display (Cont)
STATUS MESSAGE FAULT DESCRIPTION MAINTENANCE MESSAGE CODE
FADEC DEGRADED cockpit lamp flashes 8 times
FADEC DEGRADED cockpit lamp flashes 9 times
FADEC DEGRADED cockpit lamp flashes 10 times
FADEC DEGRADED cockpit lamp flashes 11 times
FADEC DEGRADED cockpit lamp flashes 12 times
FADEC FAULT cockpit lamp flashes 1 time
FADEC FAULT cockpit lamp flashes 2 times
CIT Temperature Indication
Failure
Metering Valve not in start position
Starter Relay Interface
Nr Sensor – Rotor Decay
Anticipation
Incorrect Overspeed Test Switch
Indication
HMU – Failure to control fuel flow
This Status Message is not used.
Verify that message is indicated by lamp.
Np Run Limit
T1AFlt, T1ABFlt, T1BFltRG or
T1DFltRG
OpenMvFlg
StrFlt
NrFlt
OSTstSwFlt
WfLimFlag, WfLimFlagRG
Not Used
NpRLmTOut FADEC FAULT cockpit lamp flashes 3 times
FADEC FAULT cockpit lamp flashes 4 times
FADEC FAULT cockpit lamp flashes 5 times
FADEC FAULT cockpit lamp flashes 6 times
FADEC FAULT cockpit lamp flashes 7 times
FADEC FAULT cockpit lamp flashes 8 times
FADEC FAULT cockpit lamp flashes 9 times
FADEC FAULT cockpit lamp flashes 10 times
FADEC FAULT cockpit lamp flashes 11 times
Failure in HMU Metering Valve
Potentiometer Reading
Collective Pitch Potentiometer
Indication Failure
Failure to control HMU (Stepper motor)
This Status Message is not used.
Verify that message is indicated by lamp.
Failure to control HMU
(Overspeed Solenoid)
Engine Surge Event
Ng Speed Indiction Failure
Airframe Power Supply Failure
WfMvFlt, WfStFlt or WfStFltRG
CPFlt
StepCntFlt, SmFlt, AMSolFlt or
SmFltRG
Not Used
OSFlt
SgFlag
Ng1Flt, Ng2Flt, Ng12Flt or Ng1FltRG
AF28Flt or AF28FltRG
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BHT-407-MM-9
Table 76-3. 250-C47B FADEC Software Versions 5.356 and 5.358 (Reversionary Governor)
Fault Code Display (Cont)
STATUS MESSAGE FAULT DESCRIPTION MAINTENANCE MESSAGE CODE
FADEC FAULT cockpit lamp flashes 12 times
RESTART FAULT cockpit lamp flashes 1 time
Engine Overspeed OSFlag or OSEventLmpRG
Not Used
RESTART FAULT cockpit lamp flashes 2 times
RESTART FAULT cockpit lamp flashes 3 times
RESTART FAULT cockpit lamp flashes 4 times
RESTART FAULT cockpit lamp flashes 5 times
RESTART FAULT cockpit lamp flashes 6 times
RESTART FAULT cockpit lamp flashes 7 times
RESTART FAULT cockpit lamp flashes 8 times
This Status Message is not used.
Verify that message is indicated by lamp.
TMOP Sensor – Torque
Indication Failure
This Status Message is not used.
Verify that message is indicated by lamp.
This Status Message is not used.
Verify that message is indicated by lamp.
Failure in HMU (PLA
Potentiometer) Reading
PMA Power Supply
Failure to control HMU (Hot Start
Abort Solenoid)
This Status Message is not used.
Verify that message is indicated by lamp.
Failure to control Ignition Relay
Interface
Np Speed Indication Failure
QFltRG
Not Used
Not Used
PLA1Flt, PLA2Flt, PLA12Flt or
PLARfFlt
Al28Flt or Al28FltRG
StSFlt, StSIFlt, StSFltRG or StSIFltRG
Not Used
IgnFlt or IgnIFlt RESTART FAULT cockpit lamp flashes 9 times
RESTART FAULT cockpit lamp flashes 10 times
RESTART FAULT cockpit lamp flashes 11 times
RESTART FAULT cockpit lamp flashes 12 times
FADEC DEGRADED, FADEC
FAULT, and RESTART FAULT cockpit lamps on steady
Incorrect Auto/Manual Switch
Indication
Quiet Mode Switch
All faults have been displayed
Np1Flt, Np2Flt, NpDFlt, Np12Flt,
NpDFltRG or Np1FltRG
AMSwFlt or AMSwFltRG
QMSwFlt
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BHT-407-MM-9
76-40. FADEC FAULT CODES —
PROCEDURE TO DETERMINE LAST
ENGINE RUN FAULTS FROM
“CURRENT” FAULTS
The procedure to determine if the fault codes displayed are “Last Engine Run” faults or “Current” faults may be accomplished by performing the steps in paragraph 76-39 with the throttle in the idle position.
With the throttle positioned to idle, any FADEC fault code that is displayed will be a “current” fault.
In addition, if no FADEC related lights are displayed on the caution, warning, advisory panel with electrical power applied, the FADEC in AUTO mode and the throttle positioned to idle, no “current” faults exist.
76-41. CHECK RUN PROCEDURE
Following completion of any FADEC system maintenance, a successful check run procedure is to be carried out prior to flight. Do the procedures from the BHT-407-FM-1that follow:
NOTE
Following maintenance actions and prior to performing check run procedure, ensure no
“current” faults exist.
1.
Apply electrical power to the helicopter and position FADEC Mode switch to AUTO. Wait for completion of FADEC system self-test and position throttle to idle. If no FADEC related lights are illuminated on the caution, warning, advisory panel with the throttle positioned to idle, no “current” faults exist. If a fault is displayed, refer to paragraph 76-38 .
2.
Do the PREFLIGHT and PRESTART CHECK procedures.
3.
Do the ENGINE START procedure.
4.
Do the FADEC MANUAL CHECK procedure.
5.
Do the ENGINE RUNUP procedure.
6.
Do the ENGINE SHUTDOWN procedure, use the
Overspeed shutdown test.
76-00-00
Page 44
7.
Ensure FADEC DEGRADED light does not illuminate when N
G
speed decays through 9.5%. If a fault is displayed, refer to paragraph 76-38 for a determination of faults/exceedances and required troubleshooting steps.
76-42. FADEC DOWNLOAD INTERVALS
The FADEC system is designed to notify operators of faults and exceedances as they occur or at shutdown/ pre start-up. This is accomplished through light displays on the caution, warning, and advisory panel.
This design is sufficient to advise operators when
FADEC system maintenance actions are required.
Bell Helicopter's recommendation to advise operators to carry out periodic FADEC downloads is solely intended to provide a means for all involved to protect
their interests should a discrepancy arise in regards to
how or when a fault or exceedance was recorded.
Although a specific FADEC download interval is not required in accordance with Chapter 5 of the
BHT-407-MM, Bell Helicopter recommends that operators perform a download with the Maintenance
Terminal under the following conditions:
• Every 300 hours of operation or annually
• Prior to and following completion of training flights (i.e., operation in Manual Mode)
• Prior to operation and following return of the helicopter by a 2nd party (i.e., leasing, rental, etc.)
• Prior to and following completion of maintenance actions by a maintenance facility
The FADEC download should confirm the status of
Current Faults, Last Engine Run Faults, Accumulated
Faults, Time Stamped Faults (fault history), and
Exceedance Data (engine history) ( paragraph 76-38 ).
A download of possible FADEC ECU recorded N
G
MGT, and torque (Q) exceedances is also
, recommended ( paragraph 76-37 ). During the download procedure, a printout of the fault and exceedance data should be taken and kept on file. If faults or exceedances exist, determine if maintenance actions are required.
BHT-407-MM-9
To ensure the FADEC ECU is clear of all faults and exceedances prior to return of the helicopter to service, complete the following steps:
1.
Refer to Engine History Data Screen of the
Maintenance Terminal and write down the Engine Run
Time (EngRnTm) and Number of Engine Start
(NumStrt) values.
2.
Use the Engine History Data "Clear All" command of the Maintenance Terminal. This will clear all faults and exceedances from engine history.
3.
Use Engine History "Edit" feature of the
Maintenance Terminal to enter Engine Run Time
(EngRnTm) and Number of Engine Start (NumStrt) values recorded in step 1 .
4.
Use the Fault History Data "Clear All" command of the Maintenance Terminal. This will clear all faults and exceedances from fault history.
By performing these recommended procedures, operators will be sure that the ECU is free of faults and exceedances.
76-43. ENGINE START —
TROUBLESHOOTING
The following information is provided to assist in troubleshooting engine starting problems. This information should be used in conjunction with
Rolls-Royce CSL-6108.
Lightoff parameters and standard characteristics can be described as follows:
FADEC Software Versions 5.356 and 5.358 — Once
N
G
speed reaches 10% for ambient temperatures of
80°F (26.6°C) or below, or 12% for ambient temperatures above 80°F (26.6°C) the FADEC system will introduce fuel, detect the lightoff, and smoothly accelerate the engine to idle while limiting MGT if necessary.
FADEC Software Version 5.202 — Once N
G
speed reaches 10% for ambient temperatures of 20°F
(-6.7°C) or below, or 12% for ambient temperatures above 20°F (-6.7°C) the FADEC system will introduce fuel, detect the lightoff, and smoothly accelerate the engine to idle while limiting MGT if necessary.
• A normal lightoff can be described as a start attempt that lights off at fuel introduction or 2%
N
G
after fuel introduction.
• A delayed lightoff can be described as a start attempt that does not achieve lightoff until approximately 3 to 6% after fuel introduction.
• A failure to lightoff condition can be described as a start attempt that fails to lightoff regardless of N
G
speed.
The most common factors contributing to starting problems are:
• HMU pressurizing/shutoff valve leakage
• Improper fuel nozzle shimming
• Faulty fuel nozzle
• Fuel leakback into fuel cell
• Faulty igniter circuit
The following troubleshooting may be carried out in any order to isolate the problem.
1.
The HMU pressurizing/shutoff valve may be tested for leakage as follows:
a. Locate the fuel discharge fitting on HMU. This is the long-necked fitting located directly above the
Power Lever Angle (PLA) quadrant on the HMU. The outlet of this HMU fitting faces aft and feeds the lines that lead to the fuel nozzle. Once located, remove the metal line that attaches to the fitting.
b. Ensure the throttle is in cutoff and that the
FADEC mode switch is in AUTO mode.
c. Apply electrical power to the helicopter, turn
ON one or both fuel boost/transfer pump switches, and
OPEN the fuel shutoff valve. Confirm approximately
10 to 15 PSI on the fuel pressure gauge.
NOTE
Starting problems have been identified with pressurizing valves that leak as little as three drops per minute. If the engine experiences a failure to lightoff because of pressurizing/shutoff valve leakage, the
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Page 45
BHT-407-MM-9
second start attempt is usually successful because the system has been purged.
It should also be stated that fuel leakage could also occur back through the thermal relief features of the airframe mounted fuel shutoff valve and fuel pump check valves
( step 4 ). Experience has shown however, that leakage through the HMU pressurizing/ shutoff valve almost always occurs at a much lower pressure than that which is required to crack the thermal relief valves on the airframe side. This is why we use the
10 to 15 PSI output of the boost pumps to check the HMU pressurizing/shutoff valve
(i.e., this pressure is below the cracking pressure of airframe thermal relief valves).
d. Monitor the discharge fitting for leakage
(dripping) over a 5-minute period. If any leakage is detected, the HMU will need to be replaced to resolve the starting problem ( paragraph 76-45 ). If leakage is detected, count the number of drops per minute and provide this information on the removal tag.
2.
Information on fuel nozzle shimming follows:
a. Although most engines will have approximately 4 1/2 to 5 1/2 shims installed under the nozzle, the only way to positively confirm the actual installed depth of the nozzle is to carry out the measurement procedure in Chapter 73-10-03 of the
Rolls-Royce C47B Operation and Maintenance
Manual. Once the installed depth of the nozzle is determined, you can calculate how much additional shimming may be added to move the nozzle further aft.
b. If you are in a position where you don’t have
the tools to measure the installed depth, a general rule
of thumb is to add 1/2 shim or 1 full shim maximum, and give it a try. Regardless of the shimming added, it must be ensured that the nozzle has 3 full threads for engagement.
3.
Information on fuel nozzle replacement follows:
a. If nozzle shimming does not improve lightoff characteristics and no problems are found during any of the tests provided, nozzle replacement should be considered.
b. Prior to replacing the fuel nozzle you may want to carry out the fuel nozzle inspection portion of
76-00-00
Page 46 Rev. 25 29 SEP 2008
Chapter 73-10-03 of the Rolls-Royce C47B Operation and Maintenance Manual. This test allows you to look at the spray pattern of the nozzle.
4.
Information on fuel leakback to the fuel cell follows:
This test will confirm if the check/thermal relief valves in the main fuel cell are allowing fuel to drain back into the fuel cell. This test will require the helicopter to sit for approximately 10 to 12 hours (overnight or the number of hours it takes between shutdown and start to experience your specific starting problem) and may be conducted as follows:
a. Remove the inlet line to the airframe fuel filter
(inboard line) and suspend it so that the opening in the fitting is facing up.
b. Open the fuel shutoff valve.
c. Fill the inlet line to the top of the fitting with fuel and loosely cover the fitting with a plastic bag to prevent contamination (air must be allowed to enter the bag).
d. Allow the helicopter to sit for approximately 10 to 12 hours (overnight) or the number of hours it takes for you to experience your specific starting problem.
e. The next morning, or following the number of hours it takes for you to experience your specific starting problem, inspect the line to see if any fuel has drained back. If fuel has drained back, refill the line with fuel using a syringe to measure the amount of fuel that is required to refill the line. If it only takes a few cc’s to refill the line (due to evaporation or the effects of expansion/contraction), drain back is not an issue and it will confirm that the check/thermal relief valves are not leaking. If the amount of fuel required to refill line is representative of a leak, the check valves should be cleaned, repaired or replaced (Chapter 28).
5.
The following will provide information to test the igniter circuit.
This test may be repeated as required to ensure the auto relight light is illuminating and that the igniter is firing 100% of the time. If this test proves successful, but you feel that the igniter may not be firing during an actual start attempt, it is suggested that someone (or two people) stand by the igniter during start and listen for actual firing. Although this method of testing may
BHT-407-MM-9
not be the most scientific, it is simple and accurate. It should also be stated that the firing of the igniter on the
250-C47B engine is much softer sounding than on
C20, C28, or C30 engines.
NOTE
Prior to checking the ignition system, ensure there is no fuel in the combustion section of the engine. If in doubt, conduct a dry motoring run (BHT-407-FM-1).
Following completion of the ignition system check, ensure the FADEC mode switch is repositioned to AUTO prior to conducting a start.
a. Position the FADEC mode switch into
MANUAL (innermost position) with the battery switch
OFF and external power not applied.
b. Pull the N
G
indicator circuit breaker.
c. Ensure the throttle is positioned to cutoff and the fuel valve is closed.
d. Position the battery switch to ON. Once the instrument check light goes out on the caution panel, the auto relight light will illuminate and the igniter will fire continuously.
e. To deactivate the igniter, position the battery switch to OFF.
f. To make sure the igniter will fire 100% of the time, repeat this test procedure as required. Following testing, ensure the FADEC mode switch is repositioned to AUTO and the N
G closed.
circuit breaker is
g. If the igniter always fires during this test procedure, but you feel that during actual start attempts that the igniter is not firing, it is recommended that the starter relay 1K1 be replaced (Chapter 96).
This is based on the fact that the above test procedure uses the auto relight circuitry to test the system and the normal start circuitry through the starter relay is bypassed.
6.
Information on testing results follow:
a. Based on the results of all the above-mentioned tests, a determination can be made on the possible corrective action or actions. Following any corrective action, monitor subsequent starts very closely to determine if the lightoff characteristics have been improved. Do not make more than one change to the system at any given time, unless you are absolutely sure that multiple problems exist.
b. If all testing fails to identify the problem (starts continue to fail or be delayed), it is recommended that the HMU be replaced. Although this should be accomplished as a last resort only, it is possible that the HMU may have an intermittent internal problem, which is contributing to the start problem.
c. Return helicopter to flyable condition.
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BHT-407-MM-9
HYDROMECHANICAL
76-44. HYDROMECHANICAL UNIT (HMU)
76-45.
HYDROMECHANICAL UNIT (HMU) —
REMOVAL
NOTE
Refer to Chapter 12 for information on Fuel
System — General Servicing Instructions.
1.
Disconnect electrical power from helicopter.
2.
Remove airframe drain line (1, Figure 76-10 ) and fuel line (2).
3.
Disconnect tube assembly (1, Figure 76-11 ) from lever (2) by removing nut (14), washer (12), penny washer (13), and bolt (11).
NOTE
HMU can be removed with starter generator installed.
4.
Refer to Rolls-Royce 250-C47B Operations and
Maintenance Manual, Publication CSP 21001 for remaining procedure to remove HMU.
76-46.
HYDROMECHANICAL UNIT (HMU) —
INSTALLATION
NOTE
Refer to Chapter 12 for information on Fuel
System — General Servicing Instructions.
1.
Refer to Rolls-Royce 250-C47B Operations and
Maintenance Manual, Publication CSP 21001 for installation instructions.
2.
Install airframe drain line (1, Figure 76-10 ) and fuel line (2).
3.
If previously removed, install toothed spacer (8,
Figure 76-11 ), lever (2), washer (10), and nut (9). Prior to torquing nut (9), orientate lever and toothed spacer
(8) to position noted during removal procedure. Torque nut
T
with 0.156 inch (3.96 mm) rigging pin installed through HMU lever (3) and into HMU rigging pin hole.
If position of lever and toothed spacer was not noted during the removal procedure, refer to appropriate rigging procedure per step 5 .
The following recommendations will aid in correctly orientating the lever (2) and
toothed spacer (8) during installation of
replacement HMU. Prior to removing lever
and toothed spacer from HMU, position
HMU lever (3) to allow installation of 0.156
inch (3.96 mm) rigging pin. Note orientation of lever to a reference point of your choice on HMU. Additionally, match-mark mated position of toothed spacer to lever with felt tipped marker or equivalent item. Ensure rigging pin is installed during removal of nut (9).
5.
If replacement HMU is to be installed, remove nut
(9), washer (10), lever (2), and toothed spacer (8).
6.
If replacement HMU is to be installed, remove all fittings for installation into replacement unit.
NOTE
Install bolt (11, Figure 76-11 , Detail C) with head of bolt outboard. Install large penny
washer (13) against head of bolt (11).
There is no washer installed between lever
(2) and rod end of tube assembly (1).
4.
Install tube assembly (1) to lever (2) by installing bolt (11), penny washer (13), washer (12), and nut
(14).
5.
Confirm throttle rigging. Refer to paragraph
76-47 .
6.
Perform air purge and HMU piston parking procedures in accordance with Rolls-Royce 250-C47B
Operation and Maintenance Manual, Publication CSP
21001 (Chapter 73-00-00).
7.
Perform Check Run Procedure ( paragraph
76-41 ).
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BHT-407-MM-9
2
1
1. Drain line
2. Fuel line into HMU
(from air frame fuel filter)
Figure 76-10. HMU — Removal/Installation
76-00-00
Page 50 Rev. 25 29 SEP 2008
407MM_76_0013
BHT-407-MM-9
7
9
10
SEE DETAIL
E
SEE DETAIL
C
8
SEE DETAIL
A
2
1
FWD
POSITION "X''
(THROTTLE IN
IDLE DETENT)
SEE DETAIL
D
5 10°
MIN.
B
5
4
STA
168.78
RIGGING PIN
HOLE (HMU LEVER)
MIN. STOP
6
A
SEE DETAIL
B
3
SURFACE OF
ENGINE PAN (REF)
RIGGING PIN
HOLE (HMU)
0.156 IN. (3.96 mm)
DO NOT
ADJUST
8
MAX. STOP
7
DETAIL
A
Figure 76-11. Engine Control Rigging (Sheet 1 of 3)
407MM_76_0014
29 SEP 2008 Rev. 25
76-00-00
Page 51
BHT-407-MM-9
RIGGING DIMENSIONS (NOMINAL)
6
DIMENSION
A
DIMENSION
B
DIMENSION
C
DIMENSION
D
DIMENSION
E
3.00 IN.
(76.2 mm)
0.46 IN.
(11.68 mm)
0.82 IN.
(20.83 mm)
4.28 IN.
(108.71 mm)
2.00 IN.
(50.8 mm)
FIREWALL
STA. 155.00
D
5
C
E
4
AFT
DETAIL
B
1
14
18
2
1
11
1
6
12
4
13
3
DETAIL
C
VIEW LOOKING FORWARD
18
17
16
4 16
2
15
18
17
4
16
16
15
5
DETAIL
D
Figure 76-11. Engine Control Rigging (Sheet 2 of 3)
76-00-00
Page 52 Rev. 25 29 SEP 2008
407MM_76_0015
BHT-407-MM-9
2
9
10
8
7
19
3
DETAIL
E
1.
2.
3.
4.
5.
6.
7.
Tube assembly
Lever
HMU lever
Throttle control cable
Tube assembly
Bellcrank
Hydromechanical unit (HMU)
8.
9.
10.
11.
12.
13.
14.
15.
16.
17.
18.
Serrated washer (Toothed spacer)
Nut
Washer
Bolt
Washer
Washer (Penny)
Nut
Bolt
Washer
Nut
Cotter pin
19. Fireshield
CORROSION PREVENTIVE COMPOUND (C-101)
LOCKWIRE (C-405)
80 TO 120 IN-LBS
(9.04 TO 12.43 Nm)
NOTES
1
2
3
4
Position bolt head outboard.
Position bolt head inboard.
Penny washer against bolt head (11).
One additional thick washer or thin washer may be added to accommodate cotter pin engagement.
Throttle closed 10° minimum.
5
6 Dimensions are for guidance only.
Exact dimensions may vary from helicopter to helicopter due to rigging requirements.
7
Rotation (indexing) of serrated washer
(8) will allow minor positioning adjustment of lever (2).
8
Roll pin installed in rod end. Do not adjust this end of tube assembly during rigging procedure.
407MM_76_0016
Figure 76-11. Engine Control Rigging (Sheet 3 of 3)
29 SEP 2008 Rev. 25
76-00-00
Page 53/54
BHT-407-MM-9
MECHANICAL
76-47. THROTTLE/FLY DETENT RIGGING
PROCEDURE
NOTE
The following information is only applicable to Model 407 Helicopters S/N 53390 and subsequent and those that have complied with ASB 407-99-31.
The throttle positions referred to in the following procedure reflect those stated in the BHT-407-FM-1. The positions are
Closed, Idle, FLY, and Full Open.
1.
Open the access door for the right side of the engine (Chapter 53).
2.
Disconnect the tube assembly (1, Figure 76-11 ) and tube assembly (5) from bellcrank (6).
3.
Install a rigging pin of 0.156 inch (3.96 mm) diameter through the HMU lever (3) into the rig pin hole on the HMU (7). The rig pin hole is located at the
35° power lever angle (PLA) marking on the HMU.
9.
Install tube assembly (1) to bellcrank (6) with attaching hardware (Detail D). To enable installation of tube assembly onto lever (2), position lever as required. Fine adjustments of lever can be made by indexing the serrated washer (8). Install lever with attaching hardware (Detail E). With rig pin installed, torque nut (9)
T
. Install tube assembly to lever with attaching hardware (Detail C).
10. Remove rigging pin and roll throttle to the Full
Open and Closed Positions. Verify contact is made against both upper and lower HMU stops (min stop and max stop). If applicable, this check must also be performed with copilot's throttle.
CAUTION
NOTE
HMU rigging pin must be installed for step 4 through step 9 .
4.
Rotate pilot's throttle grip to the idle position.
5.
Adjust the throttle control cable (4) to dimensions
“D” and “E”.
6.
Adjust tube assembly (5) to dimension “C”. Install tube assembly to bellcrank (6) with attaching hardware
(Detail D).
7.
Measure dimension “A”. Bellcrank (6) position “X”
(throttle in idle position) is to be within ±0.05 inch
(0.127 mm) of dimension “A”. If required, adjust rod end of tube assembly (5) to obtain dimension “A”
±0.05 inch (0.127 mm).
8.
Adjust the tube assembly (1) thread length to dimension “B”.
D O N O T A D J U S T M I N I M U M A N D
MAXIMUM STOPS ON HMU.
11. If adjustment is required to ensure contact at both stops, adjust tube assembly (1) dimension “B”, or rotate lever (2) by indexing the serrated washer (8) between the lever and HMU (7). When an adjustment is made to achieve contact at either HMU stop, the idle position must be reverified using the rigging pin and step 10 must be repeated.
12. If it is not possible to contact both HMU stops, rigging may be accomplished, per step 11 procedure, to achieve a maximum 0.020 inch (0.51 mm) gap at the maximum stop (lower stop). If it is still not possible to contact the minimum stop (upper stop) rigging may be accomplished, per step 11 procedure, to achieve a maximum 0.020 inch (0.51 mm) gap at the minimum stop. The Idle position must be verified with the rigging pin.
13. An additional method to confirm throttle rigging
(Power Lever Angle – PLA) at the Closed, Idle, and
Full Open Positions may be accomplished with the
Maintenance Terminal. Refer to the FADEC
Maintenance Terminal User's Guide, located within the
Rolls-Royce 250-C47B Operation and Maintenance
Manual, for operational information. From the Main
Menu, select Real Time Data. From the Real Time
Data Menu, select Analog Parameters. The Analog
Parameters screen will provide information on PLA
29 SEP 2008 Rev. 25
76-00-00
Page 55
BHT-407-MM-9
position. With the throttle positioned to Closed, Idle, and Full Open, the corresponding PLA signals must be as indicated in Table 76-4 .
14. Roll the throttle to the Closed position. Ensure the bellcrank (6) and the tube assembly (5) do not lock over their pivot point. Ensure an acute angle of 10° minimum between station 168.78 and the centerline of arm of bellcrank that attaches to tube (1)
( Figure 76-11 ).
15. Once rigged configuration is obtained, confirm proper installation and safteying of all hardware.
Perform throttle linkage clearance checks in engine compartment. Ensure throttle turns smoothly through complete operating range. If required, refer to paragraph 76-48 for Throttle/FLY Detent Friction
Check procedure or to paragraph 76-49 for Throttle/
FLY Detent Friction Adjustment.
NOTE
Throttle/FLY Detent Friction values must be in accordance with paragraph 76-48 prior to proceeding to step 16 .
16. Rig the detented FLY throttle position as follows:
NOTE
The following rigging steps are very important to ensure the helicopter will operate at 100% N
P
/N
R
in Auto Mode, with the throttle positioned to FLY. The steps
also configure the system to provide a N
G speed of approximately 90% in Manual
Mode, with the throttle positioned to FLY.
The Maintenance Terminal is necessary to complete the following rigging steps. Refer to the FADEC Maintenance Terminal User's
Guide, located within the Rolls-Royce
250-C47B Operation and Maintenance
Manual, for operational information.
a. Loosen four setscrews (3, Figure 76-12 ).
b. Connect Maintenance Terminal and apply helicopter electrical power. Select Real Time Data –
Analog Parameters.
c. Monitor Power Lever Angle (PLA) on Analog
Parameters screen. Position throttle grip assembly (1) to Full Open and then roll throttle down until a PLA value of 70° is obtained.
d. Maintain the 70° PLA value with throttle position and turn ferrule/bezel (2) until the ball plunger
(6) engages the groove of FLY detent (7).
e. Tighten the four setscrews (3) sufficiently to keep ferrule/bezel (2) in position.
f. Turn the throttle grip assembly (1) from Full
Open until the ball plunger (6) is engaged in the groove of FLY detent (7). Make sure that the PLA value is 69.5 to 70.5°, when the throttle is positioned from Full Open into the FLY detent. If necessary, readjust per step 16 until the 69.5 to 70.5 PLA value is obtained.
g. If repainting of the throttle position line (5) is required, position throttle until ball plunger (6) engages in groove of FLY detent (7). Paint white line 0.06 inch
(1.52 mm) wide by 0.25 inch (6.35 mm) long, aligned with the FLY position of throttle.
17. Perform Check Run Procedure ( paragraph
76-41 ).
THROTTLE POSITION
Closed
Idle
FLY
Full Open
76-00-00
Page 56 Rev. 25 29 SEP 2008
PLA
Table 76-4. Throttle Rigging Parameters
MAXIMUM DISTANCE BETWEEN HMU LEVER
AND HMU STOP (MINIMUM/MAXIMUM)
0.020 inch (0.51 mm) (Minimum Stop) -2.5 to +2.5°
34 to 36°
69.5 to 70.5°
97.5 to 102.5° 0.020 inch (0.51 mm) (Maximum Stop)
2
5
2
A
0.06 IN.
(1.52 mm)
A
3
4
0.25 IN.
(6.35 mm)
BHT-407-MM-9
1
FERRULE SETSCREW
POSITIONS (4 PLACES)
1
THROTTLE GRIP ASSEMBLY -
RECESSED AREAS FOR FERRULE
SETSCREWS (4 PLACES)
1
SECTION
A-A
DETAILS OMITTED FOR CLARITY
407MM_76_0017
Figure 76-12. Rigging Fly Detent (Sheet 1 of 2)
29 SEP 2008 Rev. 25
76-00-00
Page 57
BHT-407-MM-9
FLOAT
ARM
LDG
LTS
BOTH
F
W
D
FLOAT
INFLATE
OFF
START
DISENG
4
C
B B
C
1
SECTION
B-B
DETAILS OMITTED FOR CLARITY
2
6
3
1. Throttle grip assembly
2. Ferrule/Bezel
3. Setscrew (qty. 4)
4. Collective switch box
5. Throttle position line
6. Ball plunger
7. FLY Detent
8. Detent block
8
SECTION
C-C
1
NOTES
There is only one position, between ferrule and throttle grip, where all four setscrews will fully engage.
2 Mark with 299-947-096 epoxy paint, color to be white No. 37925 per FED-STD-595 (C-207).
Figure 76-12. Rigging Fly Detent (Sheet 2 of 2)
76-00-00
Page 58 Rev. 25 29 SEP 2008
7
407MM_76_0018
BHT-407-MM-9
76-48. THROTTLE/FLY DETENT FRICTION
CHECK
NOTE
The following information is only applicable to Model 407 Helicopters S/N 53390 and subsequent and those that have complied with ASB 407-99-31.
Copilot throttle friction will be higher than pilot throttle friction due to gearing in back of pilot collective assembly. With dual controls installed, pilot’s throttle friction may be set to a minimum of 7.8 inch-pounds
(0.88 Nm) to obtain a maximum resultant copilot throttle friction of 14.9 inch-pounds
(1.68 Nm) after break-away torque.
±0.5 pounds (6.3 to 6.8 ±0.23 kg) when the plunger is pulled through the detent.
4.
If throttle or FLY detent friction requires adjustment, refer to paragraph 76-49 .
76-49. THROTTLE/FLY DETENT FRICTION
ADJUSTMENT
NOTE
The following information is only applicable to Model 407 Helicopters S/N 53390 and subsequent and those that have complied with ASB 407-99-31.
Copilot throttle friction will be higher than pilot throttle friction due to gearing in back of pilot collective assembly. With dual controls installed, pilot’s throttle friction may be set to a minimum of 7.8 inch-pounds
(0.88 Nm) to obtain a maximum resultant
copilot throttle friction of 14.9 inch-pounds
(1.68 Nm) after break-away torque.
1.
MATERIALS REQUIRED
Refer to BHT-ALL-SPM for specifications.
NUMBER NOMENCLATURE
C-480 Cord
Wrap a suitable length of cord (C-480), or equivalent around the pilot's throttle grip and attach a fish scale to the cord.
NOTE
Depress the idle release button during the following step. Friction will increase as ball plunger (6, Figure 76-12 ) rides on surface of detent block (8).
2.
Pull the fish scale to turn the pilot's throttle grip from the Closed position to the FLY detent position.
Verify that the fish scale indicates a maximum of 10
±0.5 pounds (4.5 ±0.23 kg) after the initial breakaway force. The operation should be smooth throughout the range.
3.
Turn the pilot's throttle to put the ball plunger (6,
Figure 76-12 ) near the groove of the FLY detent (7).
Pull the fish scale to move the ball plunger in and out of the groove. Verify the fish scale value is 14 to 15
MATERIALS REQUIRED
Refer to BHT-ALL-SPM for specifications.
NUMBER NOMENCLATURE
C-480 Cord
1.
Adjust ball plunger (6, Figure 76-12 ) so that it does not contact face of detent block (8). If ball plunger has lost its thread locking capability, replace with new unit.
2.
With the use of a 3/16 inch allen key, remove the throttle friction adjustment setscrew (1, Figure 76-13 ).
Remove spring washers (2) and plug (3). Inspect washers and plug for condition. If washers are flat, cracked, or broken, replace with serviceable washers.
Inspect plug for wear. Face of plug is to be flat. If groove is worn into face of plug, replace with serviceable unit.
3.
Examine the setscrew (1) to see if there is a teflon locking element.
29 SEP 2008 Rev. 25
76-00-00
Page 59
BHT-407-MM-9
A
A
THROTTLE GRIP (REF)
3
2 2
1 1
1. Setscrew
2. Spring washers
3. Plug
SECTION
A-A
1
NOTES
Do not adjust the setscrew in too far. Damage to the spring washers will occur which will require removal of the spring tension washers to restack or replace them.
2
Stack washers as shown to create spring assembly.
Figure 76-13. Collective Throttle Friction
76-00-00
Page 60 Rev. 25 29 SEP 2008
USE 3/16" ALLEN
WRENCH TO REMOVE/
INSTALL
407MM_76_0019
BHT-407-MM-9
NOTE
Do not adjust the setscrew in too far.
Damage to the spring washers will occur, which will require removal of spring tension washers to restack or replace them.
4.
If the setscrew (1) has a locking element, install the removed components in the reverse order of step 2 and go to step 6 .
5.
If the setscrew has no locking element, it is recommended to replace the setscrew (1) with a new self-locking setscrew (P/N NAS1081-6B8).
6.
Tighten setscrew (1) until you have a slight friction when the throttle grip is turned.
NOTE
If dual controls installed, pilot throttle friction may be set to a minimum of 7.8
inch-pounds to obtain a maximum resultant copilot throttle friction of 14.9 inch-pounds after break-away torque.
To adjust the throttle friction grip setscrew (1) to get a fish scale value of 10 ±0.5 pounds (4.5 ±0.23
kg), after the initial breakaway force, do the procedure that follows:
a. Wrap a suitable length of cord (C-480), or equivalent around the pilot's throttle grip and attach a fish scale to the cord.
b. Depress the Idle detent button and pull the fish scale to rotate the throttle from the Closed position to the Full open position.
c. Adjust setscrew (1) to get the required friction value.
8.
Adjust the ball plunger (6, Figure 76-12 ) to a depth that lets it lightly contact the face of detent block (8).
9.
To adjust ball plunger (6) to get a peak rotational value of between 4 to 5 pounds (1.8 to 2.3 kg) in excess of throttle friction adjustment in step 7 , when the ball plunger rides through groove of the FLY detent
(7), do the procedure that follows:
a. Turn the throttle grip assembly (1) to put the ball plunger (6) near the groove of the FLY detent (7).
b. Wrap a suitable length of cord (C-480), or equivalent around the pilot's throttle grip assembly (1) and attach a fish scale to the cord.
c. Pull the fish scale to move the ball plunger (6) in and out of the groove. Adjust ball plunger to make sure the fish scale value is 14 to 15 ±0.5 pounds (6.3
to 6.8 ±0.23 kg) when the plunger is pulled through groove of the FLY detent (7).
10. Following the throttle and FLY detent friction adjustments, make sure the operation of the throttle is smooth and that a positive indication is felt when the ball plunger (6) is engaged in the FLY detent (7). If applicable, also use the copilot's throttle to confirm a positive indication is felt in the FLY Detent.
11. Perform Check Run Procedure ( paragraph
76-41 ).
76-50. THROTTLE CONTROL CABLE
76-51.
THROTTLE CONTROL CABLE —
REMOVAL
1.
Access the throttle control cable (68,
Figure 76-14 ).
NOTE
Prior to removing cable, note installed bends and radiuses. Also, note clamping positions on cable and installed adjustment of external threaded areas of cable. This
information may be useful during
installation of cable.
There can be two thin washers
(AN960JD10L) in place of one AN960PD10 washer (2).
2.
Remove the nut (1) and the washer (2) from the bolt (4).
3.
Remove the bolt (4) from the spacer (5), the rod end (6), and the throttle clevis (3).
4.
Loosen the jam nut (7) and remove the rod end
(6) from the adapter (8).
29 SEP 2008 Rev. 25
76-00-00
Page 61
BHT-407-MM-9
SEE DETAIL
D
70
69
66
67
66
67
66
67
25
26
27
66
67
71
SEE DETAIL
B
SEE DETAIL
A
72
68
SEE DETAIL
E
24
26
23
DETAIL
A
22
21
15
16
17
19
20
16
17
18
12
13
14
1
2
1
11
10
SEE DETAIL
C
9
8
7
6
4
3
5
DETAIL
B
Figure 76-14. Throttle Control Cable (Sheet 1 of 5)
407MM_76_0020
76-00-00
Page 62 Rev. 25 29 SEP 2008
BHT-407-MM-9
1.250 IN.
(31.75 mm)
APPROXIMATELY
15
12
13
14
11 10
9
8
7
18
DETAIL
C
6
2.0 IN.
(50.8 mm)
APPROXIMATELY
68
69 69
70 71
DETAIL
D
407MM_76_0021
Figure 76-14. Throttle Control Cable (Sheet 2 of 5)
29 SEP 2008 Rev. 25
76-00-00
Page 63
BHT-407-MM-9
SEE DETAIL
F
SEE DETAIL
G
SEE DETAIL
H
64
65
SEE DETAIL
J
DETAIL
F
63
62
SEE DETAIL
K
SEE DETAIL
L
SEE DETAIL
N
SEE DETAIL
P
SEE DETAIL
M
48
50
68
72
DETAIL
E
47
51
49
60
61
58
57
59
DETAIL
G
Figure 76-14. Throttle Control Cable (Sheet 3 of 5)
76-00-00
Page 64 Rev. 25 29 SEP 2008
DETAIL
H
407MM_76_0030
BHT-407-MM-9
Figure 76-14. Throttle Control Cable (Sheet 4 of 5)
29 SEP 2008 Rev. 25
76-00-00
Page 65
BHT-407-MM-9
8.
9.
10.
11.
12.
13.
14.
15.
1.
2.
3.
4.
5.
6.
7.
16.
17.
18.
19.
20.
21.
22.
23.
24.
Nut
Washer
Collective throttle clevis
Bolt
Spacer
Rod end
Jam nut
Adapter
Jam nut
Ball joint
Nut
Seal nut assembly (13, 14)
Sleeve
Nut
Jam nut
Screw
Washer
Bracket
Screw
Washer
Spacer
Clamp
Nut
Adapter
Washer
Washer
Clamp
Screw
Bracket
Spacer
Nut
Washer
Screw
Clamp
Nut
Washer
Clamp
Washers
Screw
Screw
Washer
Spacer
Clamp
Nut
Washer
Screw
Washer
Washer
44.
45.
46.
47.
48.
37.
38.
39.
40.
41.
42.
43.
25.
26.
27.
28.
29.
30.
31.
32.
33.
34.
35.
36.
CORROSION PREVENTIVE COMPOUND (C-101)
LOCKWIRE (C-405)
1
NOTE
Two thin washers may be substituted for one regular washer.
Install one thin washer under bolt head and install one thin washer under nut.
61.
62.
63.
64.
65.
66.
67.
68.
69.
70.
71.
72.
55.
56.
57.
58.
59.
60.
49.
50.
51.
52.
53.
54.
Spacer
Screw
Clamp
Nut
Washer
Spacer
Screw
Clamp
Nut
Washer
Screw
Clamp
Clamp
Screw
Washer
Spacer
Clamp
Sleeve
Tube
Throttle control cable
Jam nut (qty. 4)
Jam nut
Ball joint
Grommet
Figure 76-14. Throttle Control Cable (Sheet 5 of 5)
76-00-00
Page 66 Rev. 25 29 SEP 2008
407MM_76_0023
BHT-407-MM-9
5.
Loosen the jam nut (9) and remove the adapter
(8) from the ball joint (10).
6.
Remove the seal nut assembly (12) from the throttle control cable (68).
7.
Remove the inboard jam nut (15) from the throttle control cable (68).
8.
Remove the screw (19), washer (20), and the spacer (21) from the clamp (22). Remove the clamp
(22) from the throttle control cable (68).
9.
Remove the screw (27), the washer (26), and the nut (23) from the clamp (25). Remove the clamp (25) from the throttle control cable (68).
10. Remove the screw (28), the washer (29), and the spacer (30) from the clamp (31). Remove the clamp
(31) from the throttle control cable (68).
11. Remove the nut (32), the washer (33), the screw
(34), and washer (35) from the clamp (39) and the bracket (41).
12. Remove the two nuts (43), the two washers (44), and the two screws (45) from the two clamps (46).
Remove the two clamps (46) from the throttle control cable (68).
13. Remove the two nuts (47), the two washers (48), the two spacers (49) from the two screws (50).
Remove the two clamps (51) from the throttle control cable (68).
14. Remove the nut (52), the washer (53), the spacer
(54), and the screw (55) from the clamp (56).
15. Remove the nut (57), the washer (58), and the screw (59) from the clamp (60), and the clamp (61).
Remove the clamp (60) from the throttle control cable (68).
16. Remove the screw (62), the washer (63), and the spacer (64) from the clamp (65). Remove the clamp
(65) from the throttle control cable (68).
17. Loosen the four nuts (69) and the nut (70) and remove the throttle control cable (68) from the ball joint (71).
18. Remove the throttle control cable (68) from the helicopter.
76-52.
THROTTLE CONTROL CABLE —
INSPECTION
1.
Examine the throttle control cable for smooth operation.
2.
Examine the throttle control cable for signs of damage to the threaded ends.
3.
Examine the throttle control cable insulation for signs of damage and cracks.
76-53.
THROTTLE CONTROL CABLE —
INSTALLATION
C-405
MATERIALS REQUIRED
Refer to BHT-ALL-SPM for specifications.
NUMBER NOMENCLATURE
C-101 Corrosion Preventive
Compound
Lockwire
NOTE
Do not kink or apply sharp radius bends during installation of throttle control cable.
1.
Prior to installing throttle control cable (68,
Figure 76-14 ), ensure operation of internal cable is smooth throughout its full range of travel.
2.
To protect cable in roof beam passageways, install sleeves (66) and tubes (67) onto throttle control cable (68). Refer to Figure 76-14 and previously removed throttle control cable for installed locations of sleeves and tubes. Ensure grommet (72) is installed in seat structure panel.
3.
Making sure not to kink or excessively bend throttle control cable (68), route cable through airframe structure to match all clamping positions. Ensure cable is routed so that interference will not occur when flight controls are moved through their full range of travel.
Prior to installing throttle control cable into mounting hole in engine pan, install two jam nuts (69) on
29 SEP 2008 Rev. 25
76-00-00
Page 67
BHT-407-MM-9
threaded portion of throttle control cable (68). Position jam nuts so that 2.0 inches (50.8 mm) of thread will extend on aft side of engine pan ( Figure 76-14 ,
Detail D).
4.
Ensuring throttle control cable is not twisted, install two jam nuts (69) on throttle control cable (68) threads extending into engine pan. Tighten jam nuts to secure aft end of throttle control cable.
5.
Remove throttle control cable (68) from bracket
(18). Install one jam nut (15) on threaded portion of throttle control cable. Position jam nut so that 1.250
inches (31.75 mm) of thread extends outboard from bracket (18) ( Figure 76-14 , Detail C). Install throttle control cable through bracket and install inboard jam nut. Ensuring throttle control cable is not twisted, tighten jam nuts to secure forward end of throttle control cable. Ensure operation of internal cable is
smooth throughout its full range of travel.
6.
Loosely install throttle control cable (68) into all clamp assemblies per step a through step i and as shown in Figure 76-14 . Where applicable, position cable in clamps to ensure smooth bend radiuses.
Ensure that clamped position of cable will provide clearance when flight controls are moved through their full range and that clearance exists with plumbing and wire bundles.
a. Install the clamp (65) on throttle control cable
(68) and to the structure and electrical clamps with
screw (62), washer (63), and spacer (64).
b. Install clamp (60) on throttle control cable (68) and to the clamp (61) with screw (59), washer (58), and nut (57). Ensure shank of bolt faces away from control tubes.
c. Install two clamps (51) on throttle control cable
(68) and to the structure with two screws (50),
washers (48), two spacers (49), and two nuts (47).
d. Install clamp (56) on throttle control cable (68) and to the structure with screw (55), washer (53), spacer (54), and nut (52).
e. Install clamps (46) on throttle control cable
(68) and to the structure with screw (45), washers
(44), and nuts (43).
76-00-00
Page 68 Rev. 25 29 SEP 2008
f. Install clamp (39) on throttle control cable (68) and to bracket (41) with screw (34), washers (33), and nut (32).
g. Install clamp (31) on throttle control cable (68) and to the structure with screw (28), washer (29), and spacer (30).
h. Install clamp (25) on throttle control cable (68) and to adapter (24) with screw (27), washers (26), and nut (23).
i. Install clamp (22) on throttle control cable (68) and to the structure with screw (19), washer (20), and spacer (21).
7.
Once all clamps are loosely installed to throttle control cable (68), ensure operation of internal cable is smooth throughout its full range of travel, all bend radiuses are smooth, and that clamped position of cable provides clearance when flight controls are moved through their full range. Once this is confirmed, all clamps are to be permanently installed by tightening attaching hardware.
NOTE
Do not overtighten seal nut (14). Ease of internal control cable movement is to be confirmed after installation of seal nut assembly (12).
8.
Install seal nut assembly (12) on throttle control cable (68). Seal nut assembly contains sleeve (13) and nut (14).
9.
Install nut (11) onto throttle control cable (68).
Install ball joint (10) onto throttle control cable and secure with nut).
10. Install jam nut (9) on ball joint (10). Install adapter
(8) on ball joint and secure with jam nut.
11. Install jam nut (7) on rod end (6). Install rod end in adapter (8) and secure with jam nut.
NOTE
As a substitute for one washer (2) under nut (1) two thin washers may be used.
Place one thin washer under head of bolt
(4) and one under nut.
BHT-407-MM-9
12. Position rod end (6) on aft side of collective throttle clevis (3). With spacer (5) installed in collective throttle clevis, install bolt (4) through rod end (6) and collective throttle clevis. Secure with nut (1) and washer (2). Ensure head of bolt (4) is installed against aft side of rod end. Safety nut with cotter pin.
13. Install jam nut (70) onto aft end of throttle control cable (68). Install ball joint (71) to throttle control cable and secure with jam nut.
14. Confirm throttle rigging and friction settings are correct. Refer to the appropriate paragraphs as follows:
•
Paragraph 76-47 , Throttle/Fly Detent Rigging
Procedure
•
Paragraph 76-49 , Throttle/Fly Detent Friction
Adjustment
15. Confirm proper installation of all hardware and install 0.032 inch (0.812 mm) lockwire (C-405) in all locations shown in Figure 76-14 . Apply Grade 1 corrosion preventive compound (C-101), in locations shown in Figure 76-14 .
16. Perform Check Run Procedure ( paragraph
76-41 ).
29 SEP 2008 Rev. 25
76-00-00
Page 69/70
BHT-407-MM-9
ELECTRICAL
76-54. ELECTRICAL ENGINE CONTROLS —
GENERAL
This section of the chapter will cover removal, inspection, and installation of airframe mounted electrical items that are part of the FADEC system. For information on FADEC related airframe electrical circuits, which integrate with the FADEC, refer to
Chapters 95, 96, and 98.
For information on FADEC related cockpit switches
(i.e., FADEC auto/manual switch) refer to Chapter 96.
For additional information on FADEC Maintenance
Guidelines, refer to Rolls-Royce 250-C47 Commercial
Service Letter CSL-6069.
76-55.
ELECTRONIC CONTROL UNIT (ECU)
76-56.
Electronic Control Unit (ECU) — Removal
S/N 53000 Through 53749 Pre TB
407-07-75
A check for FADEC ECU recorded N
G
, MGT, and torque (Q) exceedances is also recommended
( paragraph 76-37 ).
Refer to Engine History Data Screen of Maintenance
Terminal and save or hand copy Engine Run Time
(EngRnTm) and Number of Engine Start (NumStrt) values. These values can be input into replacement
ECU following installation.
1.
Disconnect the helicopter electrical power.
2.
Gain access to the ECU by removing the forward transmission cowling (Chapter 53).
NOTE
Make sure that you do not cause damage to the contacts during removal of the ECU connectors. Put protective covers on the
ECU and harness connectors immediately after removal.
3.
Disconnect the FADEC harness electrical connectors (1, 2, Figure 76-15 ) from the ECU.
NOTE
Bell Helicopter recommends that the ECU be checked for faults and exceedances prior to removal. This may be accomplished with use of the Maintenance Terminal.
Refer to the FADEC Maintenance Terminal
User's Guide, located within the
Rolls-Royce 250-C47B Operation and
Maintenance Manual, for operational information.
The Fault History option of the Maintenance Terminal
Main Menu is to be used to check for Current, Last
Engine Run, Accumulated and Time Stamped Faults.
The Engine History option is to be used to check the
Engine History Data screen for exceedances. If faults or exceedances exist, determine appropriate maintenance action. Refer to paragraph 76-38
(Determination of Faults/Exceedances and Required
Troubleshooting Steps).
4.
Remove the screw (4), lockwasher (5), bonding strap (7), and washer (6) from the ECU (3).
5.
Remove the bolts (8, 9, 10, and 11) and washers
(13) from the top of the ECU mounting pads.
6.
Remove the ECU (3).
7.
Capture the washers (13) located on top of the roof shell inserts for the ECU (3).
8.
Remove spacers (12) from the rubber isolation dampers with finger pressure ( Figure 76-15 , Detail A).
29 SEP 2008 Rev. 25
76-00-00
Page 71
BHT-407-MM-9
10
11
8
9
6
13
5 12
9
11
9
10
3
1
14
1 2 3 4
SEE DETAIL
A
9
13
10
8
A
10
A
4
5
8
6
7
7
2 2 3 4
10
11
8
9
12
13
3
6
13
DETAIL
A
(TYPICAL)
HELICOPTER STRUCTURE
HELICOPTER MOUNTED
NUTPLATE
SECTION
A-A
(TYPICAL)
407MM_76_0024
Figure 76-15. Electronic Control Unit (ECU) — Removal/Installation S/N 53000 Through 53749
Pre TB 407-07-75 (Sheet 1 of 2)
76-00-00
Page 72 Rev. 25 29 SEP 2008
BHT-407-MM-9
5.
6.
7.
8.
9.
1.
2.
3.
4.
10.
11.
12.
13.
14.
ECU-to-airframe electrical connector
ECU-to-engine electrical connector
ECU
Screw (MS35206-261)
Lockwasher (MS35338-43)
Plain washer (NAS1149D0332J)
Bonding strap
Bolt (NAS6203-12)
Bolt (NAS6203-12)
Bolt (NAS6203-12)
Bolt (NAS6203-18)
Spacer (NAS43DD3-34N)
Washer (AN970-3)
Decal (31-053-18CFHP)
30 TO 40 IN-LBS
(3.4 TO 4.5 Nm)
1
NOTES
Disconnect the electrical power from the helicopter when you remove or install the ECU.
2 Make sure that you do not cause damage to the connector contacts when you remove or install.
3
4
5
6
Put protective covers on the ECU and the harness connectors immediately after the removal.
Make sure that the red band on each ECU connector is not visible after you install the harness connectors.
Install one spacer in each of the ECU mounting pad isolation dampers.
Install one washer on the top of and three washers under each ECU mount.
7
8
Pre S/N 53200 or Pre TB 407-98-9, bonding strap
MS25083-2BB6, S/N 53200 and subsequent or Post
TB 407-98-9, bonding strap 961114-1.
Use only screw MS35206-261. Use of any other type screw will damage the threads of the ECU casing.
9
10
Inboard mounting pad isolation damper coler (RED).
Outboard mounting pad isolation damper color (GREEN).
407MM_76_0025
Figure 76-15. Electronic Control Unit (ECU) — Removal/Installation S/N 53000 through 53749
Pre TB 407-07-75 (Sheet 2 of 2)
29 SEP 2008 Rev. 25
76-00-00
Page 73
BHT-407-MM-9
76-57.
Electronic Control Unit (ECU) — Removal
S/N 53000 Through 53749 Post TB
407-07-75 and S/N 53750 and Subsequent
NOTE
Bell Helicopter recommends that the ECU be checked for faults and exceedances prior to removal. This may be accomplished with use of the Maintenance Terminal.
Refer to the FADEC Maintenance Terminal
User’s Guide, located within the
Rolls-Royce 250-C47B Operation and
Maintenance Manual, for operational information.
The Fault History option of the Maintenance Terminal
Main Menu is to be used to check for Current, Last
Engine Run, Accumulated and Time Stamped Faults.
The Engine History option is to be used to check the
Engine History Data screen for exceedances. If faults or exceedances exist, determine appropriate maintenance action. Refer to paragraph 76-38
(Determination of Faults/Exceedances and Required
Troubleshooting Steps).
A check for FADEC ECU recorded N
G
, MGT, and torque (Q) exceedances is also recommended
( paragraph 76-37 ).
Refer to Engine History Data Screen of Maintenance
Terminal and save or hand copy Engine Run Time
(EngRnTm) and Number of Engine Start (NumStrt) values. These values can be input into replacement
ECU following installation.
1.
Disconnect the helicopter electrical power.
2.
Gain access to the ECU by removing the forward transmission cowling (Chapter 53).
NOTE
Make sure that you do not cause damage to the contacts during removal of the ECU connectors. Put protective covers on the
ECU harness connectors immediately after removal.
76-00-00
Page 74 Rev. 25 29 SEP 2008
3.
Disconnect the FADEC harness electrical connectors (1 and 2, Figure 76-16 ) from the ECU.
4.
Remove the screw (4), lockwasher (5), bonding strap (7), and washer (6) from the ECU (3).
5.
Remove the bolts (8, 9, 10, and 11) and washers
(12) from the top of the ECU mounting pads.
6.
Remove the ECU (3).
7.
Capture the washers (12) located on top of the roof shell inserts for the ECU (3).
76-58.
Electronic Control Unit (ECU) —
1.
and corrosion.
2.
NOTE
damage (e.g., bent pins) and corrosion.
3.
dampers for condition and security.
4.
blockage.
5.
Ensure the “NO STEP” decal (Bell Helicopter P/N
31-053-18CFHP) is installed on top outer case of
ECU.
For additional information on care and inspection of the FADEC system, refer to
Rolls-Royce 250-C47 Series CSL-6069.
Examine the ECU outer case for signs of damage
Examine the ECU electrical receptacles for
Examine the ECU mounting pad isolation
Examine the ECU pressure sensing (P1) port for
6.
Ensure ECU mounting fasteners on roof shell are checked for condition and security.
7.
Inspect ECU bonding strap for condition and security.
BHT-407-MM-9
8
9
10
11
5
12
8
11
8
10
3
1
14
1
2 3 4
SEE DETAIL
A
9
8
8
B
8
B
4
5
7
6
7
6
2
2 3 4
8
9
10
11
12
5
12
12 3
DETAIL
A
(TYPICAL)
HELICOPTER
STRUCTURE
HELICOPTER
MOUNTED
NUTPLATE
SECTION
B-B
(TYPICAL)
407MM_76_0031
Figure 76-16. Electronic Control Unit (ECU) — Removal/Installation S/N 53000 Through 53749 Post TB
407-07-75 and S/N 53750 and Subsequent (Sheet 1 of 2)
29 SEP 2008 Rev. 25
76-00-00
Page 75
BHT-407-MM-9
7.
8.
9.
4.
5.
6.
10.
11.
12.
13.
1.
2.
3.
ECU-to-airframe electrical connector
ECU-to-engine electrical connector
ECU
Screw (MS35206-261)
Lockwasher (MS35338-43)
Plain washer (NAS1149D0332J)
Bonding strap
Bolt (NAS6203-14)
Bolt (NAS6203-14)
Bolt (NAS6203-14)
Bolt (NAS6203-20)
Washer (AN970-3)
Decal (31-053-18CFHP)
30 TO 40 IN-LBS
(3.4 TO 4.5 Nm)
1
2
3
4
5
6
7
8
NOTES
Disconnect the electrical power from the helicopter when you remove or install the ECU.
Make sure that you do not cause damage to the connector contacts when you remove or install.
Put protective covers on the ECU and the harness connectors immediately after the removal.
Make sure that the red band on each ECU connector is not visible after you install the harness connectors.
Install one washer on the top of and three washers under each ECU mount.
Pre S/N 53200 or Pre TB 407-98-9, bonding strap
MS25083-2BB6, S/N 53200 and subsequent or Post
TB 407-98-9, bonding strap 961114-1.
Use only screw MS35206-261. Use of any other type screw will damage the threads of the ECU casing.
Mounting pad isolation dampers contain integral one piece spacer assembly.
407MM_76_0032
Figure 76-16. Electronic Control Unit (ECU) — Removal/Installation S/N 53000 Through 53749 Post TB
407-07-75 and S/N 53750 and Subsequent (Sheet 2 of 2)
76-00-00
Page 76 Rev. 25 29 SEP 2008
BHT-407-MM-9
76-59.
Electronic Control Unit (ECU) —
Installation S/N 53000 Through 53749 Pre
TB 407-07-75
NOTE
ECU has removable NAS43DD3-34N spacers. Install spacers in accordance with step 1 .
1.
Install the four spacers (12, Figure 76-15 , Detail
A) into the ECU mounting pads.
2.
Install a stack-up of three washers (13) over each of the four ECU mounting positions.
3.
Position the ECU (3) mounting pads over the washers (13).
4.
Install one washer (13) on top of each of the four
ECU mounting pads.
NOTE
ECU mounting bolts (8, 9, and 10) have a
-12 grip length. ECU mounting bolt (11) has
a -18 grip length.
5.
Install ECU mounting bolts (8, 9, 10, and 11).
Torque bolts
T
.
6.
Install washer (6), bonding strap (7), lockwasher
(5), and screw (4) into the ECU.
7.
Only remove protective covers on the ECU and FADEC harness connectors just prior to installation. Make sure that you do not cause damage to the contacts during installation of the ECU connectors. Make sure that the red band on each ECU electrical receptacle is not visible after installation of the FADEC harness connectors.
Install the FADEC harness electrical connectors
(1 and 2) to the ECU (3).
NOTE
NOTE
Bell Helicopter recommends that the ECU be checked for faults and exceedances following installation (prior to the first engine start). This will ensure the ECU is free of faults and exceedances prior to return to service. This may be accomplished with use of the Maintenance
Terminal. Refer to the FADEC Maintenance
Terminal User's Guide, located within the
Rolls-Royce 250-C47B Operation and
Maintenance Manual, for operational information.
The Fault History option of the Maintenance Terminal
Main Menu can be used to check for Current, Last
Engine Run, Accumulated and Time Stamped Faults.
The Engine History option can be used to check the
Engine History Data screen for exceedances. Any last
Engine Run Faults, Accumulated Faults, or
Exceedances should be cleared prior to first engine start. The Fault History Data "Clear All" command of the Maintenance Terminal may be used to ensure
Fault History is clear of faults or exceedances. To ensure N
G
, MGT, and Torque (Q) exceedances do not exist in ECU memory, use Engine History Data “Clear
All” command of the Maintenance Terminal. If active
“current” faults exist, determine appropriate maintenance action ( paragraph 76-38 ).
Use Engine History “Edit” feature of Maintenance
Terminal to enter Engine Run Time (EngRnTm) and
Number of Engine Starts (NumStrt) that were saved or hand copied during ECU removal.
8.
Apply electrical power to the helicopter and position FADEC Mode switch to AUTO. Ensure no active Current Faults exist. This can be accomplished by waiting for completion of FADEC system self-test and positioning throttle to idle. If no FADEC related lights are illuminated on the caution, warning, advisory panel with the throttle positioned to idle, no “current” faults exist. If a fault is displayed, refer to paragraph
76-38 .
9.
Install required cowlings (Chapter 53).
10. Perform Check Run Procedure ( paragraph
76-41 ).
29 SEP 2008 Rev. 25
76-00-00
Page 77
BHT-407-MM-9
76-60.
Electronic Control Unit (ECU) —
Installation S/N 53000 Through 53749 Post
TB 407-07-75 and 53750 and Subsequent.
1.
Install a stack-up of three washers (12,
Figure 76-16 , Detail A) over each of the four ECU mounting positions.
2.
Position the ECU (3) mounting pads over the washers (12).
3.
Install one washer (12) on top of each of the four
ECU mounting pads.
NOTE
ECU mounting bolts (8,9, and 10) have a
-14 grip length. ECU mounting bolt (11) has a -20 grip length.
4.
Install ECU mounting bolts (8, 9, 10, and 11).
Torque bolts
T
.
5.
Install washer (6), bonding strap (7), lockwasher
(5), and screw (4) into the ECU.
NOTE
Only remove protective covers on the ECU and FADEC harness connectors just prior to installation. Make sure that you do not cause damage to the contacts during installation of the ECU connectors. Make sure that the red band on each ECU electrical receptacle is not visible after installation of the FADEC harness connectors.
6.
Install the FADEC harness electrical connectors
(1 and 2) to the ECU (3).
NOTE
Bell Helicopter recommends that the ECU be checked for faults and exceedances following installation (prior to the first engine start). This will ensure the ECU is free of faults and exceedances prior to return to service. This may be
accomplished with use of the Maintenance
Terminal. Refer to the FADEC Maintenance
Terminal User’s Guide, located within the
Rolls-Royce 250-C47B Operation and
76-00-00
Page 78 Rev. 25 29 SEP 2008
Maintenance Manual, for operational information.
The Fault History option of the Maintenance Terminal
Main Menu can be used to check for Current, Last
Engine Run, Accumulated and Time Stamped Faults.
The Engine History option can be used to check the
Engine History Data screen for exceedances. Any last
Engine Run Faults, Accumulated Faults, or
Exceedances should be cleared prior to first engine start. The Fault History Data "Clear All" command of the Maintenance Terminal may be used to ensure
Fault History is clear of faults or exceedances. To ensure N
G
, MGT, and Torque (Q) exceedances do not exist in ECU memory, use Engine History Data “Clear
All” command of the Maintenance Terminal. If active
“current” faults exist, determine appropriate maintenance action ( paragraph 76-38 ).
Use Engine History “Edit” feature of Maintenance
Terminal to enter Engine Run Time (EngRnTm) and
Number of Engine Starts (NumStrt) that were saved or hand copied during ECU removal.
7.
Apply electrical power to the helicopter and position FADEC Mode switch to AUTO. Ensure no active Current Faults exist. This can be accomplished by waiting for completion of FADEC system self-test and positioning throttle to idle. If no FADEC related lights are illuminated on the caution, warning, advisory panel with the throttle positioned to idle, no “current” faults exist. If a fault is displayed, refer to paragraph
76-38 .
8.
Install required cowlings (Chapter 53).
9.
Perform Check Run Procedure ( paragraph
76-41 ).
76-61.
COLLECTIVE PITCH TRANSDUCER (CPT)
The Collective Pitch (CP) transducer is installed under the copilot seat. The transducer tells the FADEC system the rate of movement of the collective control stick. The transducer is an electrical potentiometer that is connected to the airframe at one end and connected to a clamp assembly installed on the collective jackshaft. When the collective control stick is
lifted or lowered, the transducer increases or
decreases in length, which changes the resistance output to the FADEC system. The output signal allows the FADEC to provide anticipation logic, which helps to reduce main rotor RPM droop and overshoot.
BHT-407-MM-9
76-62.
Collective Pitch Transducer (CPT) —
Removal
1.
Disconnect helicopter electrical power.
2.
Remove the copilot seat and seatback
(Chapter 25).
3.
Remove the metal copilot seat panel assembly.
4.
Disconnect the CP transducer electrical connector (1, Figure 76-17 ).
5.
Remove the nut (2) and the spacer (3) from the screw (4).
6.
Remove the screw (4) and the washer (5) from the CP transducer (6) and the support (7).
7.
Remove the nut (8) and the washer (9) from the screw (10).
8.
Remove the screw (10) and the spacer (11) from the clamp assembly (12) and the CP transducer (6).
9.
Remove the CP transducer (6).
Collective Pitch Transducer (CPT) —
Inspection
NOTE
For additional information on care and inspection of the FADEC system, refer to
Rolls-Royce 250-C47B Series CSL-6069.
1.
Examine the CP transducer for signs of damage, pitting, and corrosion.
2.
Examine the CP transducer, mounting support, and clamp assembly for condition, security, and eccentric bolt holes.
3.
Examine the electrical leads for signs of chafing
and damaged insulation. Examine the electrical
connector for condition.
4.
Examine spherical bearings on CP transducer for condition and security.
76-64.
Collective Pitch Transducer (CPT) —
Installation/Rigging
CAUTION
M A K E S U R E T H E C O L L E C T I V E
CONTROL SYSTEM IS RIGGED BEFORE
YOU RIG THE CP TRANSDUCER. REFER
TO CHAPTER 67.
NOTE
It is recommended that the CP transducer workaid be used to avoid damage to the CP transducer ( Figure 76-17 , Detail B).
1.
Adjust the collective pitch (CP) transducer (6,
Figure 76-17 , Detail A) to obtain 6.18 inch (156.9 mm) dimension between centers of the grounded bearing and adjustable rod end bearing with the transducer movable rod at the mid stroke position.
2.
Remove forward transmission cowling
(Chapter 53).
MATERIALS REQUIRED
Refer to BHT-ALL-SPM for specifications.
NUMBER NOMENCLATURE
C-101 Corrosion Preventive
Compound
NOTE
As an alternate procedure to step 3 and step 10 , a hydraulic cart may be connected to the helicopter (Chapter 29).
3.
Remove attaching hardware between collective link assembly (17) and collective lever (18). Position collective link assembly clear of collective lever.
4.
Raise collective stick until up-stop is contacted.
Hold in position with collective friction ( Figure 76-17 ,
Detail C).
29 SEP 2008 Rev. 25
76-00-00
Page 79
BHT-407-MM-9
12
13
8
9
11
SEE DETAIL
A
10
6.18 IN.
(156.9 mm)
4
5
6
3
2
1
ADJUST MIDSTROKE POSITION
TO 6.18 IN. (156.9 mm)
DETAIL
A
7
0.38 IN.
(9.65 mm)
0.50 IN.
(12.7 mm)
0.193 TO 0.198 IN.
(4.90 TO 5.03 mm)
2 HOLES
6.77 IN.
(171.95 mm)
0.25 IN.
(6.35 mm)
WORK AID
7.60 IN.
(193.0 mm)
DETAIL
B
COLLECTIVE PITCH TRANSDUCER RIGGING WORK AID
0.125 IN.
(3.17 mm)
STOCK
Figure 76-17. Collective Pitch Transducer (CPT) — Removal/Installation (Sheet 1 of 3)
407MM_76_0026
76-00-00
Page 80 Rev. 25 29 SEP 2008
BHT-407-MM-9
Figure 76-17. Collective Pitch Transducer (CPT) — Removal/Installation (Sheet 2 of 3)
29 SEP 2008 Rev. 25
76-00-00
Page 81
BHT-407-MM-9
CLAMP
LBL
8.05
SCREW
7.47 IN.
(189.7 mm)
WL
28.62
STA
73.12
WL
28.62
6.77 IN.
(171.95 mm)
WL
21.03
2
SUPPORT
DETAIL
E
STA
71.03
DETAIL
F
SHOWN COLLECTIVE FULL UP
1. Electrical connector
2. Nut
3. Spacer
4. Screw
5. Washer
6. Collective pitch transducer
7. Support
8. Nut
9. Washer
10. Screw
11. Spacer
12. Clamp assembly
13. Collective jackshaft
14. Bolt
15. Washer
16. Nut
17. Collective link assembly
18. Collective lever
CORROSION PREVENTIVE COMPOUND (C-101)
95 TO 110 IN-LBS
(10.7 TO 12.4 Nm)
1
NOTES
When reinstalling collective link assembly to collective lever, torque nut. See Detail D.
2
Transducer mounting faces of support (7) and clamp (12) are to be in-line.
Figure 76-17. Collective Pitch Transducer (CPT) — Removal/Installation (Sheet 3 of 3)
407MM_76_0028
76-00-00
Page 82 Rev. 25 29 SEP 2008
BHT-407-MM-9
(
7.
9.
screw (4).
NOTE
If adjustments are required in the following step 5 or step 6 , reposition clamp (12) as required. Tighten screws on clamp maintaining equal gaps between clamp halves. Do not tighten screws more than necessary to hold adjustment.
5.
Confirm clamp assembly (12) is installed on collective jackshaft (13) at left butt line (LBL) 8.05
Figure 76-17 , View E).
6.
With collective full-up, confirm dimension between mounting hole of support (7) and clamp (12) is 6.77 inch (171.95 mm). Once dimension has been obtained, remove workaid if used. Refer to
Figure 76-17 , Detail B for recommended workaid.
Position the CP transducer (6,
NOTE
the support (7) and the CP assembly (6).
Figure 76-17
between the support (7) and the clamp assembly (12).
The head of the screw (4) must point inboard when installed.
8.
Install the washer (5) and the screw (4) through
Install the spacer (3) and the nut (2) on the
10. Prior to installing the CP transducer (6) to the clamp (12), make sure the CP transducer rod end will fit the clamp mounting hole position with collective full
up and collective full down without causing restriction
to collective travel or damage to CP transducer.
NOTE
The head of the screw (10) must point outboard when installed.
11. Install the spacer (11) and the screw (10) through the CP transducer (6) and the clamp assembly (12).
12. Install the washer (9) and the nut (8).
13. Apply Grade 1 corrosion preventive compound
(C-101), as shown in Figure 76-17 .
)
14. Connect the CP transducer electrical connector
(1). Fold and secure the excess connector harness wiring with lacing cord or plastic cable ties.
NOTE
If a collective pitch (CP) transducer functional test is to be carried out per paragraph 76-65 , it should be conducted prior to accomplishment of step 15 .
15. Reinstall collective link assembly (17) to collective lever (18) with bolt (14), washers (15), and nut (16). Torque nut
T
and apply Grade 1 corrosion preventive compound (C-101) per Figure 76-17 ,
Detail D.
16. Install forward transmission cowling (Chapter 53).
17. Install the metal copilot seat panel assembly.
18. Install copilot seat and seat back (Chapter 25).
19. Apply electrical power to the helicopter and position FADEC Mode switch to AUTO. Ensure no
and positioning throttle to idle. If no FADEC related lights are illuminated on the caution, warning, advisory panel with the throttle positioned to idle, no “current” faults exist. If a fault is displayed, refer to paragraph
76-38 .
20. Perform Check Run Procedure ( paragraph
76-41 ).
76-65.
Collective Pitch Transducer (CPT) —
Functional Test
NOTE
The Maintenance Terminal will be required to complete the Functional Test.
1.
Remove forward transmission cowling
(Chapter 53).
As an alternate procedure to step 2 and step 5
, a hydraulic cart may be connected
to the helicopter (Chapter 29).
29 SEP 2008 Rev. 25
76-00-00
Page 83
BHT-407-MM-9
2.
Remove attaching hardware between collective link assembly (17) and collective lever (18). Position collective link assembly clear of collective lever
( Figure 76-17 , Details C and D).
3.
Install the Maintenance Terminal. Refer to the
Maintenance Terminal Users Guide for operating instructions.
a. Connect the helicopter electrical power.
b. Select Real Time Data from the Main Menu.
c. Select Analog Parameters from the Real Time
Data Menu.
d. View Analog Parameters for Collective Pitch
(CP) reading.
NOTE
If step e and step f of this test are carried out by disconnecting the CP transducer and moving the shaft of transducer by hand, it is quite possible that a fault will be detected by the FADEC. This will be displayed on the caution, warning, advisory
panel as a FADEC DEGRADE. A defaulted
CP value will then be displayed on the
Maintenance Terminal regardless of shaft position. Cycle helicopter electrical power to remove the fault and defaulted CP value.
Conduct step e and step f with CP transducer installed.
e. With collective full down (down stop contacted, Figure 76-17 , View C), CP is to be 0 to 5%.
RESULT:
• If collective pitch is 0 to 5% with collective full down proceed to step f .
CORRECTIVE ACTION:
• If collective pitch is not 0 to 5% with collective full down, confirm rigging per paragraph
76-64 . If rigging is acceptable, ensure no shorts or opens exist in wiring between CP transducer and ECU. If wiring is acceptable,
consider replacing CP transducer.
f. With collective full up (up-stop contacted,
Figure 76-17 , View C), CP is to be 95 to 100%.
76-00-00
Page 84 Rev. 25 29 SEP 2008
RESULT:
• If collective pitch is 95 to 100% with collective full up proceed to step g .
CORRECTIVE ACTION:
• If collective pitch is not 95 to 100% with collective full up, confirm rigging per paragraph 76-64 . If rigging is acceptable, ensure no shorts or opens exist in wiring between CP transducer and ECU. If wiring is acceptable consider replacing CP transducer.
g. Disconnect the electrical power from the helicopter.
4.
Remove the Maintenance Terminal.
5.
Install attaching hardware between collective link
assembly (17) and collective lever (18) (
Figure 76-17 ,
Details C and D).
6.
Install forward transmission cowling (Chapter 53).
76-66.
COMPRESSOR INLET TEMPERATURE
(CIT) SENSOR
1.
The purpose of the Compressor Inlet
Temperature (CIT) sensor is to provide the FADEC with information on compressor air inlet temperature.
The CIT sensor is mounted on the upper left-hand side of the forward engine firewall. When installed, the actual temperature probe of the sensor is located on the forward side of the engine firewall. In this position, the probe of the CIT sensor is protected from foreign object damage and from ice buildup. The sensor has
two temperature sensing coils for redundancy.
76-67.
Compressor Inlet Temperature (CIT)
Sensor — Removal
1.
Remove electrical power from the helicopter.
NOTE
Install protective plastic cap on connector end of CIT sensor following removal.
2.
Remove electrical connector (1, Figure 76-18 ) from CIT sensor (2).
BHT-407-MM-9
3.
Remove CIT sensor attaching screws (3) and washers (4).
76-68.
Compressor Inlet Temperature (CIT)
Sensor — Inspection
faults exist. If a fault is displayed, refer to paragraph
76-38 .
4.
Perform Check Run Procedure ( paragraph
76-41 ).
76-70.
Compressor Inlet Temperature (CIT)
Sensor — Functional Test
NOTE
For additional information on care and inspection of the FADEC system, refer to
Rolls-Royce 250-C47B Series CSL-6069.
1.
Inspect CIT sensor and its electrical contacts
(pins) for visible damage, moisture, and corrosion.
2.
Inspect CIT sensor mating connector and its
contacts (sockets) for damage, moisture, and
corrosion.
3.
Inspect condition of CIT mounting fasteners on
firewall.
76-69.
1.
(3,
Compressor Inlet Temperature (CIT)
Sensor — Installation
Install CIT sensor to firewall with attaching screws
Figure 76-18 ) and washers (4).
The Maintenance Terminal (Windows version) will be required to complete the
Functional Test.
1.
Install the Maintenance Terminal. Refer to the
Maintenance Terminal Users Guide for operating instructions.
NOTE
NOTE
Low voltage may affect reading of CIT sensor. For best results apply 28 volts.
2.
Connect the helicopter electrical power.
3.
Using the Maintenance Terminal, select Real
Time Date from the Main Menu.
4.
Select Analog Parameters from the Real Time
Data Menu.
NOTE
Remove protective plastic cap from connector end of CIT sensor just prior to installation of mating connector. When installing mating connector to CIT sensor, ensure contacts are not damaged and that connector is tightened until red line on CIT sensor is no longer visible.
2.
Install electrical connector (1) to CIT sensor (2).
3.
Apply electrical power to the helicopter and position FADEC Mode switch to AUTO. Ensure no active Current Faults exist. This can be accomplished by waiting for completion of FADEC system self-test and positioning throttle to idle. If no FADEC related lights are illuminated on the caution, warning, advisory panel with the throttle positioned to idle, no “current”
NOTE
To ensure accurate temperature readings, helicopter must be positioned in an area where the effects of heat soaking do not exist. If helicopter is moved from outside to hanger for purposes of this test, ensure sufficient time is allowed for temperature of
CIT sensor to equal hanger temperature.
5.
View Analog Parameter for Compressor Inlet
Temperature (T1). Compare value shown with a known accurate source. CIT temperatures should be within ±1.8°F (±1°C). Operating range of CIT sensor is
-65°F to 212°F (-54°C to 100°C).
29 SEP 2008 Rev. 25
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BHT-407-MM-9
WL
92.22
LBL
11.50
5
4
3
2
1
4
3
1. Electrical connector
2. CIT sensor
3. Screw
4. Washer
5. Forward engine firewall
Figure 76-18. Compressor Inlet Temperature (CIT) Sensor — Removal/Installation
407MM_76_0029
76-00-00
Page 86 Rev. 25 29 SEP 2008
BHT-407-MM-9
RESULT:
• If temperatures are within ±1.8°F (±1°C), proceed to step 6 .
CORRECTIVE ACTION:
• If temperatures are not within ±1.8°F (±1°C), do not immediately assume the CIT sensor is faulty. As the CIT sensor contains two independent coils and the wiring to the ECU is separate for each coil, it is possible that the temperature source that the CIT was being compared to is inaccurate. If the FADEC detects a difference in the rate or range of the temperature signals between the two CIT circuits, a fault will be declared. To further check the CIT circuit, refer to the Fault
Isolation Manual in Chapter 73-25-04 of the
Rolls-Royce 250-C47B Operation and
Maintenance Manual, Publication CSP 21001.
The section “CIT (T1) TEMPERATURE
SENSOR CIRCUIT FAULT” will provide resistance values to check both the CIT sensor and its associated wiring to the ECU.
6.
Disconnect electrical power from the helicopter.
7.
Remove the Maintenance Terminal.
29 SEP 2008 Rev. 25
76-00-00
Page 87/88
BHT-407-MM-10
MAINTENANCE MANUAL
VOLUME 10
INSTRUMENTS/ELECTRICAL
NOTICE
The instructions set forth in this manual, as supplemented or modified by Alert Service Bulletins (ASB) or other directions issued by Bell Helicopter Textron Inc. and Airworthiness Directives (AD) issued by the applicable regulatory agencies, shall be strictly followed.
COPYRIGHT NOTICE
COPYRIGHT 2008
BELL ® HELICOPTER TEXTRON INC.
AND BELL HELICOPTER TEXTRON
CANADA LTD.
ALL RIGHTS RESERVED
22 FEBRUARY 1996
REVISION 25 — 29 SEPTEMBER 2008
BHT-407-MM-10
PROPRIETARY RIGHTS NOTICE
These data are proprietary to Bell Helicopter Textron Inc. Disclosure, reproduction, or use of these data for any purpose other than helicopter operation or maintenance is forbidden without prior written authorization from
Bell Helicopter Textron Inc.
PN
Additional copies of this publication may be obtained by contacting:
Commercial Publication Distribution Center
Bell Helicopter Textron Inc.
P. O. Box 482
Fort Worth, Texas 76101-0482
Rev. 24 2 OCT 2007
BHT-407-MM-10
can be reset at any time by using a computer and the software as described in Paragraph 95-60 of this chapter.
The life of the battery is estimated to be 10 years. If the battery goes dead, only the date function is affected. The indicator will continue to operate and record exceedances without the date stamp.
95-55. PROPULSION INSTRUMENTS
BUILT-INTEST (BIT) — DESCRIPTION
All of the propulsion instruments have a built-in test
(BIT) capability.
The two possible BIT functions are the power-on BIT and the commanded BIT.
95-56. Power-on BIT
The power-on BIT starts when power is applied to the instrument. The power-on BIT does an integrity check of the electronic components of the indicator. On all indicators during this BIT, except for the dual tachometer, all of the Liquid Crystal Displays (LCDs) on the analog single bar display come on and show the maximum scale. On the torque, MGT, and NG indicators, the digital display flashes and the letter E comes on. Any failure found during the BIT for these specific indicators is written to their respective non-volatile memory (NVM). Refer to paragraph 95-58 for the procedure to retrieve the failure codes for each indicator. On the dual tachometer indicator, the individual rotor (R) and turbine (T) pointers are driven to indicate their respective upper red line limits, if no faults are detected. If faults are detected, the pointer movement does not occur. The BIT causes these indications for approximately 6 to 8 seconds. After that time, the indicators go back to their normal indication.
95-57. Commanded BIT
The commanded BIT starts when the LCD CHECK switch (1S15) is pushed, and performs a check of all of the LCDs. On all indicators during this BIT, all of the
LCDs on the analog single bar display come on and show the maximum scale. In addition, the rotor (R) and turbine (T) pointers on the dual tachometer are driven to their respective upper limits.
95-58. PROPULSION INSTRUMENTS —
EXCEEDANCE DATA RETRIEVAL/
VIEWING/ERASING/INDICATOR CLOCK
SET
The torque indicator, MGT indicator, and NG indicator are connected to an RS-485 maintenance bus. The
RS-485 maintenance bus gives access to the memory of these instruments to retrieve and erase exceedance data. The RS-485 maintenance bus also gives access to the real time clock of the three instruments, to set the time of the clock.This is done with the use of a laptop computer and software. The laptop computer is connected to the MAINTENANCE PORT -
INSTRUMENTS receptacle of the RS-485 maintenance bus with a special cable assembly.
To download the latest propulsion instruments maintenance software and and associated user guide, please go to www.bellcustomer.com and select the link for “Maintenance Software”.
95-59. Propulsion Instruments — Exceedance
Data Retrieval
NUMBER
SPECIAL TOOLS REQUIRED
IBM ® or compatible
DOS (any version)
NOMENCLATURE
Computer (PC)
Disk operating system
(DOS)
Cable assembly 407-275-001-103
407-275-001-107
Software EPM Plus
(Requires approximately 60
Kbytes of diskspace.)
To download the latest propulsion instruments maintenance software and and associated user guide, please go to www.bellcustomer.com and select the link for “Maintenance Software”.
1.
Connect the cable assembly to the serial port of the laptop computer.
2.
At the left hand side of the pedestal, remove the cap from the ENG INSTR connector.
3.
Connect the end of the cable assembly to the
ENG INSTR connector.
4.
On the overhead panel, close the circuit breakers that follow:
• ENGINE INSTR TRQ (1CB25)
• ENGINE INSTR MGT (1CB10)
• ENGINE INSTR NG (1CB22)
29 SEP 2008 Rev. 25
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BHT-407-MM-10
5.
Connect 28 VDC external power to the helicopter.
6.
Set the switch of the laptop computer to the ON position.
7.
At the PC DOS prompt (C:), type 'time' and push
ENTER.
RESULT:
-
The current time of the computer clock is shown.
8.
If the current time is correct, push ENTER.
9.
If the current time is not correct, type the correct time (for example: 09:03:00) and push ENTER.
13. Push ENTER.
RESULT:
-
The menu display that follows comes on:
'ENTER LETTER CORRESPONDING TO
DESIRED FUNCTION Enter choice:'.
14. Type 'a' and push ENTER (to select which indicators are active).
RESULT:
-
The display 'SELECT INDICATORS TO BE
ACTIVE ON THE 485 DOWNLOAD BUS' comes on with a list of indicators.
15. Use the up and down arrows of the computer keyboard to move the cursor to a specific indicator.
NOTE
The procedure that follows assumes that the download software EPMPLUS is located in the drive A: of the computer. If the software is located elsewhere, then change to that drive and sub-directory prior to typing EPMPLUS.
10. Insert the software diskette into the drive A of the laptop computer.
11. At the PC DOS prompt (C:), type 'a:' and push
ENTER.
12. At the 'A:' prompt, type 'EPMPLUS' and push
ENTER.
RESULT:
-
The license agreement message comes on the screen of the computer.
NOTE
Do not add other indicators than MGT, Ng, and torque with an 'X' selection. The other indicators do not have data that can be retrieved and they are not connected to the
485 download bus. If you select 'X' these indicators, the program will show an error message since it will not find the requested data in the related instrument(s).
16. Select an indicator by typing an "X" in the square brackets. Select the desired indicators from MGT, Ng, and torque only.
17. Deselect any other indicators by hitting the computer keyboard spacebar to blank the space inside the square brackets.
18. Press ENTER to save the indicators selection and exit this screen.
NOTE
If this is the first time this software has been used, the first screen to display will be
'SELECT INDICATORS TO BE ACTIVE ON
THE 485 DOWNLOAD BUS'. Otherwise, the screen displayed is 'ENTER LETTER
CORRESPONDING TO DESIRED
FUNCTION Enter choice:'.
RESULT:
-
The main menu display that follows comes on:
'ENTER LETTER CORRESPONDING TO
DESIRED FUNCTION'.
19. Type 'C' and push ENTER (to select a COM port).
95-00-00
Page 74 Rev. 4
BHT-407-MM-10
Table 95-16. NG Scale Indication Vs Input Signal
INDICATION
(%)
5
10
20
105
105.6
106
107.0
110
120
130
70
80
90
100
30
50
60
63
INPUT
FREQUENCY
(HZ)
893
1785
3570
18 743
18 850
18 921
19 099
19 635
21 420
23 205
5355
8925
10 710
11 246
12 495
14 280
16 065
17 850
NOMINAL
SIGNAL
VOLTAGE
(PEAK TO
PEAK)
0.6
1.2
2.4
12.6
12.6
12.7
13.0
13.2
14.4
15.6
8.4
9.6
10.8
12.0
3.6
6.0
7.2
7.6
The input signal from the N1 pickup is supplied to the signal conditioner. The output of the signal conditioner is sent to the analog to digital (A/D) converter where the analog signal is converted to a digital signal. The digital signal is then sent to the processor. The processor controls the operation of the indicator from the operation parameters of the software program in its memory.
The 28 VDC bus supplies the power input to the electronic components of the indicator. The 28 VDC is supplied through the ENGINE INSTR NG (1CB24) circuit breaker on the overhead panel. Three fluorescent green lights are mounted behind the dial face of the NG indicator. The LCDs of the NG indicator do not normally allow the light to pass through them.
With input to the NG indicator, the NG indicator turns each LCD segment clear. This allow the green light to shine through illuminating each segment. The green lights are controlled in intensity depending on the input from the 5 VDC lighting circuit.
When the FADEC/ECU is operating in manual mode, the engine out warning system in the FADEC/ECU is inoperative. When the FADEC/ECU switch is placed in the manual mode, an alternate ground circuit is provided for the engine out warning horn through the
FADEC/ECU switch and the NG indicator. Pin 13 of the
NG indicator is normally closed to ground. This contact will open when the N1 pickup signal is 9818 Hz or above (55% NG or above). Therefore, with the FADEC/
ECU in the manual mode, if the NG indicator drops below 55 %, the engine out horn will operate.
95-88.
NG Indication System — Operational Test
NOTE
To view exceedances recorded in the NVM with a computer, refer to paragraph 95-58.
1.
Close the ENG INSTR NG (1CB24) circuit breaker.
2.
Connect 28 VDC external power to the helicopter or set the BATT ON OFF switch to ON.
RESULT:
• All LCDs come on.
• All digits and decimals are displayed.
• After 6 to 8 seconds, the indicator ends the power-on BIT and return to the normal indication of 0%.
CORRECTIVE ACTION:
• If the LCDs do not come on, refer to Trouble
No. 1, Figure 95-44.
• If all the digits and decimals do not show, refer to Trouble No. 2, Figure 95-45.
• If indicator does not return to the normal indication after 6 to 8 seconds, replace the indicator.
29 SEP 2008 Rev. 25
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BHT-407-MM-10
95-00-00
Page 112 Rev. 4
Figure 95-43. NG indication system - Simplified schematic
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