AVIATION INVESTIGATION REPORT A11C0047 DOUBLE ENGINE POWER LOSS AND FORCED LANDING

AVIATION INVESTIGATION REPORT A11C0047 DOUBLE ENGINE POWER LOSS AND FORCED LANDING
AVIATION INVESTIGATION REPORT
A11C0047
DOUBLE ENGINE POWER LOSS AND FORCED LANDING
FUGRO AVIATION CANADA LIMITED
CASA C-212-CC40, C-FDKM
SASKATOON, SASKATCHEWAN
01 APRIL 2011
The Transportation Safety Board of Canada (TSB) investigated this occurrence for the purpose
of advancing transportation safety. It is not the function of the Board to assign fault or
determine civil or criminal liability.
Aviation Investigation Report
Double Engine Power Loss and Forced Landing
Fugro Aviation Canada Limited
CASA C-212-CC40, C-FDKM
Saskatoon, Saskatchewan
01 April 2011
Report Number A11C0047
Synopsis
At 1503 Central Standard Time, the Construcciones Aeronauticas SA (CASA) C-212-CC40
(registration C-FDKM, serial number 196) operated by Fugro Aviation Canada Ltd., departed
from Saskatoon/Diefenbaker International Airport, Saskatchewan, under visual flight rules for
a geophysical survey flight to the east of Saskatoon. On board were 2 pilots and a survey
equipment operator. At about 1814, the right engine lost power. The crew shut it down, carried
out checklist procedures, and commenced an approach for Runway 27. When the flight was 3.5
nautical miles from the runway on final approach, the left engine lost power. The crew carried
out a forced landing adjacent to Wanuskewin Road in Saskatoon. The aircraft impacted a
concrete roadway noise abatement wall and was destroyed. The survey equipment operator
sustained fatal injuries, the first officer sustained serious injuries, and the captain sustained
minor injuries. No ELT signal was received.
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Table of Contents
1.0 Factual Information ......................................................................... 2
1.1
1.2
1.3
1.4
1.5
1.6
1.7
1.8
1.9
1.10
1.11
1.12
1.13
1.14
1.15
History of the Flight......................................................................................................... 2
Injuries to Persons ............................................................................................................ 8
Damage to Aircraft .......................................................................................................... 8
Other Damage .................................................................................................................. 8
Personnel Information ..................................................................................................... 8
Aircraft Information ........................................................................................................ 9
Meteorological Information .......................................................................................... 16
Flight Recorders ............................................................................................................. 16
Wreckage and Impact Information ............................................................................. 16
Fire ................................................................................................................................... 24
Survival Aspects............................................................................................................. 24
Tests and Research ......................................................................................................... 24
Organizational and Management Information .......................................................... 24
Additional Information ................................................................................................. 24
Useful or Effective Investigation Techniques ............................................................ 25
2.0 Analysis ........................................................................................... 26
2.1
2.2
2.3
2.4
2.5
2.6
2.7
2.8
2.9
2.10
Right Engine Power Loss .............................................................................................. 26
One Engine Inoperative Performance ......................................................................... 26
Bird Recovery ................................................................................................................. 27
Fuel System ..................................................................................................................... 27
Low Fuel Level Warning .............................................................................................. 28
Left Engine Power Loss................................................................................................. 29
Flight Following ............................................................................................................. 29
Annunciator Bulb Analysis .......................................................................................... 30
Forced Landing .............................................................................................................. 30
Flight Recorders ............................................................................................................. 30
3.0 Conclusions..................................................................................... 31
3.1
3.2
3.3
Findings as to Causes and Contributing Factors ....................................................... 31
Findings as to Risk ......................................................................................................... 31
Other Findings................................................................................................................ 31
4.0 Safety Action................................................................................... 33
4.1
Action Taken................................................................................................................... 33
5.0 Appendices ..................................................................................... 35
Appendix A -Flight Path 1814 to 1825:36.................................................................................... 35
Appendix B – Engine Failure in Flight Procedure..................................................................... 36
Appendix C – Flight Path 1825:36 to 1830:02 ............................................................................. 37
Appendix D – Before Landing Check ......................................................................................... 38
Appendix E – Multiple Engine Failure and Emergency Landing Procedures ...................... 39
Appendix F –Fuel Level Warning Procedure ............................................................................ 40
Appendix G – Fuel Pressure Warning Procedure ..................................................................... 41
Appendix H – Left Wing Fuel System Testing Results............................................................. 42
Table of Photos
Photo 1. C-FDKM in flight with birds stowed. (Source: Fugro Aviation Canada Ltd, used with
permission).................................................................................................................................................. 3
Photo 2. C-FDKM in flight with birds deployed. (Source: Jacqueline Thiessen, used with
permission).................................................................................................................................................. 3
Photo 3. Crash site overview (Source: Saskatoon Police Service, used with permission) .......Error!
Bookmark not defined.
Photo 4. Failed gear on the torque sensor shaft ................................................................................... 18
Photo 5. Tooth from gear 896884-1 showing wear on the loaded face and flank ........................... 19
Photo 6. Gear 3103589-1 tooth wear on the loaded face and flank ................................................... 19
Photo 7. No. 2 ejector nozzle and foreign object debris...................................................................... 20
Photo 8. Foreign object debris as ingested into No. 2 ejector pump during bench testing ........... 21
Table of Figures
Figure 1. Simplified schematic of the left wing fuel system ............................................................. 12
Figure 2. Annunciator panel .................................................................................................................. 23
Figure 3. Annunciator panel showing illuminated segments ........................................................... 28
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1.0 Factual Information
1.1
History of the Flight
1.1.1
Pre-flight Preparation
The aircraft and crew conducted a survey flight on the morning of the occurrence, and the
aircraft operated normally throughout the flight. After that flight, an aircraft maintenance
engineer employed by Fugro Aviation Canada Limited (Fugro) fuelled the aircraft with 1703
litres of Jet A fuel.
Before the occurrence flight, the pilots conducted a pre-flight inspection of the aircraft during
which all annunciator panel lights were confirmed to be operational. The single red line (SRL) 1
and torque/temperature limiting (TTL) 2 systems were also confirmed to be operational. The
pilots calculated the aircraft weight and balance using a spreadsheet on a laptop computer.
The captain occupied the left seat and the first officer (FO) the right seat. Prior to engine start,
the crew reset the fuel totalizers 3 to zero. Both engine starts were normal. The flight departed
from Saskatoon at 1503. 4
1.1.2
Survey Flight
The area to be surveyed was within 30 nautical miles (nm) to the east of Saskatoon over
generally flat terrain at an elevation of 1700 to 1800 feet above sea level (asl). There were a
number of towers and electrical lines in the area.
Initial survey equipment calibration was done at 5500 feet asl, followed by descent to 400 feet
above ground level (agl) for the survey. The aircraft was not equipped with an autopilot and the
pilot flying maintained survey altitude using visual reference and the radar altimeter.
The first 3 hours of the flight were completely normal. The pilots transferred control hourly and,
during the occurrence, the FO was the pilot flying. The flight was northbound on the last northsouth line of the survey block with about 10 nm to complete.
The aircraft was flying straight and level in the survey configuration with flaps up and the 2
survey sensors, or birds, (Photo 1 and Photo 2) deployed behind and below the aircraft. Altitude
was 2300 feet asl, indicated airspeed was 130 knots (KIAS) with engine power stable at 99.7%
rpm and 60–65% torque. The left and right normal fuel booster pumps were on, the fuel
crossfeed was closed, and both left and right fuel pressure instruments were indicating normal
readings. No annunciators or warning lights were illuminated and there were no abnormal
1
2
3
4
To eliminate repeated reference to graphs to obtain the variable exhaust gas temperature
(EGT) limits, the SRL computer emits a conditioned signal to the EGT indicator so the
indicated EGT limit is always 650 °C regardless of conditions.
The TTL system protects against inadvertent exceedance of torque and EGT limits.
The aircraft was equipped with fuel totalizers which record the amount of fuel consumed by
each engine.
All times Central Standard Time (Coordinated Universal Time minus 6 hours).
-3-
engine instrument indications. None of the engine control levers had been moved recently. Fuel
flow was 340 pounds per hour (pph) on the left engine and 360 pph on the right engine.
Photo 1. C-FDKM in flight with birds stowed. (Source: Fugro Aviation Canada Ltd,
used with permission)
Photo 2. C-FDKM in flight with birds deployed. (Source: Jacqueline Thiessen, used
with permission)
1.1.3
Right Engine Power Loss
At about 1814:20 ( Appendix A), the aircraft experienced a shudder from the right engine just
before it smoothly spooled down. The propeller rotation slowed and the aircraft yawed. The
right engine torque fell below 20%. The crew confirmed that the right engine had lost power.
1.1.4
Response to Right Engine Power Loss
The captain set maximum power on the left engine (100% torque, 100% rpm and exhaust gas
temperature [EGT] greater than 600 °C and less than the limit of 650 °C) and verbally advised
the FO that maximum power was set. All engine indications for the left engine remained
normal.
The FO applied pressure to the left rudder pedal to counter the yaw and adjusted the rudder
trim to reduce, but not eliminate, the rudder pedal force required. The rudder trim was not set
to full travel. The FO requested the captain assist on the left rudder pedal and the captain did
-4-
so. This method was used for the remainder of the flight, and the rudder trim setting remained
unchanged for the remainder of the flight.
The FO banked the aircraft 5° left to help maintain directional control, and maintained the left
bank until the left engine lost power later in the flight. The ball on the turn and slip indicator
was steady at well left of centre.
The captain began to action the memory items for the engine failure in flight procedure
(Appendix B). When the captain called for the FO to confirm the right emergency shutdown
lever, the FO halted him before the lever was moved. The FO then turned on the left and right
auxiliary fuel booster pumps; both the left and right normal booster pumps were already on.
After a couple of seconds, the FO confirmed the right emergency shutdown lever and the
captain moved it to the feather position. The propeller feathered, and the captain completed the
remaining memory items of the procedure, including “hydraulic pump on” and setting the
flaps to 25%. The flaps stayed at 25% for the remainder of the flight. After completing the
memory items, the captain reset the master caution light.
The FO concentrated on aircraft control because performance was marginal. Airspeed decreased
substantially to less than 95 KIAS.
As the FO had more experience on the aircraft than the captain, the pilots agreed to let the FO
continue as pilot flying. The captain would continue with the checklists and radio
communications with air traffic control (ATC).
The crew decided that, before recovering the sensor birds, the survey equipment operator
would remain seated until they had confirmed the aircraft would climb. If necessary, the birds
would be released. Either recovery or release would require the survey equipment operator to
get out of his seat and move about the cabin to operate the required equipment.
At 1816:30, the flight had climbed to 2500 asl, and the FO turned toward the northwest. This
permitted the flight to proceed closer to Saskatoon while avoiding a group of towers
immediately to the west.
The captain then referred to the quick reference handbook (QRH) for the engine failure in flight
procedure checklist, confirmed all memory items had been carried out, and completed the
remaining items on the checklist. Both the right normal and right auxiliary fuel booster pumps
were turned off; the left normal and left auxiliary fuel booster pumps remained on.
At 1817:44 the captain notified the Saskatoon control tower of the situation with a MAYDAY
call, advising they had lost power on one engine and were returning to Saskatoon. The aircraft
was equipped with an on-board SkyTrac system, which transmitted position data and other
aircraft information during the flight via a satellite link. The FO activated the emergency
MAYDAY function of the SkyTrac system.
After completing the checklist, the captain kept his left hand on the control column, right hand
on the power levers, and feet on the rudder pedals. He was ready to assist the FO or take
control if necessary, but was not applying any pressure to the controls other than the left rudder
pedal.
-5-
The crew did not attempt to restart the right engine; their priorities were aircraft controllability,
climbing to higher altitude, recovering the birds, and returning to Saskatoon.
After the engine failure in flight checklist had been completed, the indicated right fuel pressure
increased to and remained at 50 pounds per square inch (PSI). This was a result of high ambient
nacelle temperature causing expansion of the trapped fuel in the lines between the closed
firewall fuel valve and the fuel control unit. The fuel crossfeed was closed at the time and the
fuel pressure on the left engine remained normal.
1.1.5
Bird Retrieval
By 1818:07, the aircraft had climbed to 2600 feet asl and accelerated to about 99 KIAS, and the
pilots had instructed the survey equipment operator to recover the birds. The survey equipment
operator got up to do so. The recovery process took 4 to 5 minutes, during which time the flight
initially continued climbing to the northwest to avoid the towers to the west and then turned
westward.
The captain referred to the QRH a second time to verify that all steps in the engine failure in
flight procedure had been completed. During this period the MAYDAY function of the SkyTrac
system was unintentionally deactivated.
By about 1823:30, bird recovery was complete, and the survey equipment operator returned to
his seat and strapped in. The flight had climbed to 2900 feet asl, and airspeed was about 100
KIAS.
1.1.6
Preparation for Landing
As the flight continued toward the Saskatoon airport, rather than optimizing climb rate the
crew chose to slowly increase airspeed to provide an increased margin of safety above the
minimum control speed. At 1825:36 (Appendix C), the flight reached 3100 feet asl (about 1400
agl), the maximum altitude following the right engine power loss.
At about 1826, the crew completed the portion of the “Before landing check” above the line
(Appendix D). The right normal and right auxiliary fuel booster pumps were turned on during
the before landing check. Because the nose wheel steering tiller is on the left side of the cockpit,
the pilots discussed the need to transfer control after landing.
By 1827, airspeed was about 105 to 110 KIAS, and both pilots had visual contact with the
runway. The FO’s plan was to fly a high steep approach, and he had not yet reduced the left
engine power lever setting for the approach.
At 1827:11 the tower cleared the flight to land on runway 27.
The FO still had the aircraft banked 5° left with the rudder pedal at full deflection, and the ball
in the turn and slip indicator was displaced about halfway to the left.
The captain still had his left hand on the control column, right hand on the power levers, and
feet on the rudder pedals, but was not applying any pressure to the controls. The captain was
-6-
following through on the controls and was ready to assist or assume control if necessary. The
FO’s hands were on the control column.
Neither pilot observed the fuel level left tank or fuel pressure left engine annunciators
illuminate after the engine failure in flight procedure had been completed for the right engine
power loss. The FO did not look at the master caution light. The captain did not see the master
caution light illuminate after it had been reset following completion of the engine failure in
flight procedure for the right engine power loss.
The captain was frequently scanning the left engine instruments after the right engine had lost
power. He observed that:
• there were no abnormal indications;
• there were no annunciator panel lights associated with the left engine;
• the left fuel pressure light was not illuminated;
• the left fuel pressure indication was stable and not fluctuating;
• there were no low fuel warnings; and
• fuel flow was stable.
1.1.7
Left Engine Power Loss
At about 1828:25, airspeed was 105 KIAS or greater when the left engine smoothly lost power
with no surging. The captain was looking at the engine instruments at the time, and all
indications had been normal. Torque and rpm were 100% with no fluctuations; EGT was
slightly greater than 600 °C and less than the 650 °C limit; fuel pressure, oil temperature, oil
pressure, and fuel flow were all normal. The captain observed the torque indicator smoothly
and rapidly decreasing from 100% to 20% within a couple of seconds. The captain did not
observe any of the other engine instruments as the loss of power occurred.
Both pilots could feel the aircraft decelerate when the engine lost power. As the torque
decreased to 20%, the captain could feel the FO changing the position of the rudder pedals. Both
pilots could feel the aircraft yawing to the left, and it was clear to both that the left engine had
lost power. Airspeed at the time was about 110 KIAS, and the altitude was 3000 feet asl (about
1300 feet agl). The captain verbally advised the FO that the left engine had lost power.
1.1.8
Forced Landing
Both pilots immediately concluded the flight would not reach the runway, and the captain
pointed out a large street to the FO as a feasible forced landing site and the FO concurred.
Neither pilot could see any better alternative.
The captain immediately reached for the left emergency shutdown lever and called for the FO
to confirm the captain had the correct lever. The FO confirmed it and the captain moved the left
emergency shutdown lever to feather.
The FO turned the aircraft toward the street. At 1829:07, the captain advised the Saskatoon
tower controller of the second loss of power and indicated that they would be landing on the
street.
-7-
Because of the low altitude and limited time to select and fly to a suitable forced landing site,
the crew elected to leave the flaps at 25% and did not complete any other emergency checklist
actions or refer to the QRH (Appendix E).
The FO maintained control of the aircraft throughout the descent. Airspeed slowed initially to
100 KIAS and later to 90 KIAS. The goal was to control the descent and impact, with a minimum
airspeed of 90 KIAS to prevent a departure from controlled flight. The stall horn did not sound
although the FO did observe the stall warning light flashing. The captain continued to back up
the FO on the controls.
The pilots observed traffic on the road. They revised the intended touchdown site to the grass
beside the road, to the right of the traffic and light poles and the left of what they perceived as a
frangible fence.
Late in the descent the pilots identified that the fence was actually a noise abatement wall. The
FO concluded that the flight could not be extended beyond the noise abatement wall and
advised the other crew members. The aircraft landed astride the wall at 90 KIAS (Photo 3).
When the aircraft came to a stop, the captain turned off the batteries and the master switch and
pressed the cockpit engine fire extinguisher buttons.
Photo 3. Crash site overview (Source: Saskatoon Police Service, used with permission)
-8-
1.2
1.3
Injuries to Persons
Crew
Passengers
Others
Total
Fatal
1
–
–
1
Serious
1
–
–
1
Minor/None
1
–
–
1
Total
3
–
–
3
Damage to Aircraft
The aircraft was destroyed by impact forces.
1.4
Other Damage
Property damage was limited to the noise abatement wall and landscaping in the vicinity of the
wall, with minor environmental damage from hydraulic and battery fluid spillage.
1.5
Personnel Information
Records indicate that both pilots were certified and qualified for the flight in accordance with
existing regulations. There was nothing found to indicate that either pilot’s performance was
degraded by physiological factors. Both flight crew members work/rest schedules were such
that fatigue is not considered a factor in this occurrence.
The captain held an Airline Transport Pilot Licence, a glider pilot licence, and a category 1
medical certificate valid until 01 October 2011. The captain had accumulated approximately
7400 hours of total flight time prior to the occurrence with approximately 75 flight hours on the
CASA C-212.
The FO held an Airline Transport Pilot Licence, a glider pilot licence, and a category 1 medical
certificate valid until 01 July 2011. The FO had accumulated approximately 7800 hours of total
flight time prior to the occurrence with approximately 1800 flight hours on the CASA C-212.
The FO was certified and qualified as a CASA C-212 captain. Fugro practice is for pilots to
alternate seats and pilot-in-command duties weekly, and it was the FO’s week to fly from the
right seat as FO.
No simulator is available for the survey configuration of the CASA C-212, and the captain and
FO received all their training in the aircraft. Training for one engine inoperative with the birds
deployed cannot be done in the aircraft, because the birds cannot be deployed without a survey
equipment operator and simulated engine failures cannot be conducted with an equipment
operator on board.
Despite these factors, Fugro does provide pilots with training in the CASA C-212 for engine
failure in flight during survey operations, albeit without the birds deployed. To simulate the
-9-
additional drag of the birds, aircraft performance is reduced by conducting the training at 5000
to 8000 feet agl at the maximum aircraft weight permitted for ambient conditions. The training
exercise is initiated with a simulated engine failure, with the crew carrying out the engine
failure in flight procedure and climbing 500 to 1000 feet above the altitude at which the exercise
began. Applicable checklists are then completed, and the exercise concludes with a return to the
airport for a landing with one engine simulated inoperative. Training records show that both
the captain and the FO received this training.
While multiple engine failure cannot be practiced in the aircraft, the situation is discussed
during training.
1.6
Aircraft Information
1.6.1
General
Manufacturer
Type and model
Year of manufacture
Serial number
Special Certificate of Airworthiness – Restricted, issue date
Total airframe time
Engine type (number of)
Propeller/rotor type (number of)
Maximum allowable take-off weight
Recommended fuel type(s)
Fuel type used
Construcciones Aeronauticas SA
C-212-CC40
1981
196
2004-06-01
21292.7 hrs
Garrett TPE331-10R-511C (2)
Hartzell HCB4MN-5AL (2)
7700 kg
Jet A, Jet A-1, JP-8, Jet B, JP-5
P4, Jet A, Jet A1sss
Jet A
The aircraft (Photo 1) was authorized for use to conduct aerial geophysical surveys under the
provisions of a Special Certificate of Airworthiness – Restricted. The restricted classification was
assigned because the aircraft type certificate was in the restricted category and also because the
aircraft had been modified in such a manner that it no longer complied with the original
certificate. See section 1.6.3 regarding the modifications.
Records indicate that the aircraft was certified, equipped, and maintained in accordance with
existing regulations and approved procedures. The aircraft underwent an inspection, which
primarily focused on the airframe with a few engine-related requirements, as well as an
inspection of the right engine on 03 Dec 2010, approximately 98 flight hours prior to the
occurrence.
On the morning of the occurrence, the aircraft was due for other inspections, including one of
the left engine, within 8.4 hours. These inspections had been largely completed and the
completion of the remaining inspection items and final signoff had been extended 10 hours in
accordance with the Fugro Maintenance Control Manual. Maintenance personnel had recorded
the completed inspection items on the applicable tally sheets. Items that were complete
included replacement of the wing fuel system ejector pump filters and functional tests of fuel
system valves and booster pumps.
-10-
1.6.2
Engines and Propellers
The engines had an approved overhaul period of 7000 hours. A spectrographic oil analysis
program (SOAP) was an integral part of the operator’s maintenance program. The SOAP
required that oil samples be taken every 200 hours and sent to an independent lab for analysis.
Both engines were equipped with magnetic chip plugs in the reduction gear housing. When the
contact of a chip plug is shorted to ground by metal particles the corresponding chip detector
segment of the annunciator panel will illuminate.
The engines on C-FDKM were not equipped with an auto-ignition system. Such a system is not
normally installed on these engines on the CASA C-212-CC40.
The aircraft was equipped with propellers which rotate counter-clockwise as viewed by the
crew. Because of the counter-clockwise rotation, the right engine is the critical engine. 5 “The
critical engine is the engine, the failure of which would most adversely affect the performance
or handling qualities of an aircraft”. 6
1.6.3
Aircraft Modifications
In 1989, the aircraft was modified in accordance with supplemental-type approval SA86-9
“Installation of Electro-Magnetic (EM) Survey System and Associated Stability Finlets”. The EM
survey system included a loop comprising 6 heavy-gauge cables that surrounded the aircraft.
The cables were secured at structural points at each wing tip, at tubular extensions on the nose,
and at an aft fuselage extension. The main batteries were relocated from the left landing gear
sponson to a more forward location in the main cabin for centre of gravity reasons. The
auxiliary battery remained in its original location in the sponson. Finlets were installed midspan on the upper and lower surfaces of each horizontal stabilizer.
As a result of SA86-9, the performance characteristics of the modified aircraft differed from the
unmodified aircraft as manufactured by CASA. See section 1.6.7 for information regarding
aircraft performance.
The unmodified CASA C-212 is required to have rudder control surface movement range of
+27.5° ± 1° to -27.5° ± 1°. 7 After modification, in accordance with SA86-9, the rudder control
surface movement range is required to be +20° to -20°. 8
During development of SA86-9, the contractor and Transport Canada determined the CASA C212, with the EM system installed, would not meet the certification requirements for directional
5
6
7
8
A propeller develops more thrust on the downward travelling side of the propeller disc due to
that blade’s higher angle of attack. In the case of a propeller rotating counter-clockwise, this
results in a yawing moment to the right. With 2 propellers rotating counter-clockwise, the
thrust line of the left engine is further from the aircraft centreline than the right engine.
Consequently, the right engine is referred to as the “critical engine” because loss of this engine
will result in the greatest yawing moment.
Canadian Aviation Regulations, Part 1 – General Provisions, Subpart 1 – Interpretation.
CASA C-212-200 Maintenance Manual 28-20-10, pp. 601–604.
Fugro Aviation Canada Limited (FACL) Maintenance manual supplement FACL-6238-02,
issue 2 (20 July 1986), p. 27.2.
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stability. Consequently, stability finlets and a rudder limitation were installed to bring
performance within the directional stability requirements.
Transport Canada type certificate A-147 prescribes the conditions and limitations under which
the modified CASA C-212-CC40 meets the standards of airworthiness required by the Canadian
Aviation Regulations. 9 This type certificate was originally approved 29 August 1986. Regarding
the basis of certification, the certificate states that:
•
•
special performance conditions apply with respect to the Magnetic Loop/Finlets
modification contained in letter 5010-10-375 (ABP/A) to Field Aviation dated May 8,
1986. Ref. Supplemental Type Approval SA86-9;
all aircraft must include the incorporation of supplemental type approval SA86-9 for the
magnetic loop and finlets and must be operated in accordance with the appropriate DOT
approved Flight Manual Supplement.
SA86-9 also included the installation of a geophysical sensor bird, a winch, and a stowage rack.
The bird could be deployed behind and below the aircraft, and, when recovered, was stowed in
the rack located on the exterior surface of the aircraft cargo door. In 1997, a further modification
added a second sensor bird and winch system in accordance with LSTA O-LSA-96-0139/D. In
2004, a new bird rack was installed in accordance with LSTA C-LSA4-108D.
Other modifications included improvements to the survey equipment operator seat that added
a shoulder harness in 2004 and a headrest in 2006.
At the time of the occurrence, the emergency procedure to release the birds was for the survey
equipment operator to use a pair of wire clippers to manually cut the cables. The pilots had no
means of jettisonning the birds from the cockpit.
9
Transport Canada Type Certificate Data Sheet A-147, Issue 7, 28 May 1997.
-12-
1.6.4
Fuel System
Each wing, outboard of
the engine nacelle,
houses 2 integral fuel
tanks located between
the forward and rear
spars. The lower surface
of the wing has no
dihedral, so the bottoms
of the fuel tanks are flat
in the span-wise
direction.
The left and right wing
fuel systems are identical
but independent. A
crossfeed line with a
mechanical valve
controlled from the
cockpit permits fuel from
either wing to be
supplied to the engine on
the opposite wing
(Figure 1). Throughout
the flight, the fuel
crossfeed valve remained
closed. Normal operating
procedures and the
Figure 1. Simplified schematic of the left wing fuel system (source:
CASA C-212-200 Maintenance Manual, chapter 28-20-10, p. 20)
engine failure in flight
procedure (Appendix B) did not require it to be opened.
The capacity of each inboard and outboard tank is 719 and 299 litres respectively. The aircraft’s
total fuel capacity is 2036 litres, of which 1999 litres is usable.
Usable fuel is fuel available for aircraft propulsion. Unusable fuel is fuel remaining in the
system and in the tanks when the fuel quantity indicators read zero. 10
The rear inboard corner of each inboard tank is closed off by a wing rib and a forward baffle to
form a collector tank. Fuel is delivered from each collector tank to its respective engine. During
normal operation, the collector tanks are kept full, and serve to prevent fuel starvation due to
movement and sloshing of fuel in the main tanks. Each collector tank has a capacity of
165 pounds of Jet A fuel (89.7 litres at 1.8397 pounds per litre) with wings level. With the wings
banked, capacity will decrease due to spillage into the inboard tank through stringer cut-outs in
the collector tank walls on the low side. At 5° bank angle, collector tank capacity is 157 pounds
of Jet A fuel (85.3 litres at 1.8397 pounds per litre).
10
CASA C-212 Airplane Flight Manual, page A.1-6.
-13-
Each collector tank houses a low fuel level switch that illuminates a corresponding (left or right)
fuel level warning light on the annunciator panel if the fuel quantity in the collector tank
decreases to about 100 pounds. Because of the location of the level switch in the middle of the
collector tank, this warning level does not vary significantly with side to side movement of the
fuel in the collector tank.
In normal cruise flight, the fuel level warning light illuminates at a fuel quantity sufficient for
about 17 minutes of engine operation. At full power, that duration decreases to about 12
minutes. The low fuel level warning system is functionally tested every 1800 hours.
Illumination of the fuel level warning light alerts the crew to impending fuel starvation, and
requires the crew to complete the fuel level warning procedure (Appendix F). The first step in
this procedure is to move the fuel feeding lever to the “crossfeed” position, thereby opening the
crossfeed valve.
Two submersed electric fuel booster pumps are installed in each collector tank and deliver fuel
through check valves and a series of shut-off valves to provide a positive inlet pressure to the
engine-driven fuel pump. At least one booster pump is required to be operating at all times. The
booster pumps also provide primary fuel flow to operate 3 ejector pumps in each wing.
A fuel pressure transmitter and a fuel low pressure switch, mounted on the forward face of each
engine firewall, are installed in the fuel lines between the electric firewall valves and the
engines. Each fuel pressure transmitter has a corresponding fuel pressure instrument in the
cockpit. Each fuel pressure switch will illuminate a corresponding segment of the annunciator
panel if fuel pressure drops below 3.5 pounds per square inch (psi). The fuel pressure
annunciator will extinguish if the fuel pressure subsequently increases to above 7.0 psi.
Illumination of the fuel pressure warning light alerts the crew to impending fuel starvation, and
requires the crew to complete the fuel pressure warning procedure (Appendix G).
1.6.5
Ejector Pumps
Ejector pumps have no moving parts; they operate using the Venturi principle and incorporate
an unscreened inlet pipe situated close to the bottom surface of the fuel tank. “Primary fuel
pressure from the booster pump(s) is accelerated through the jet nozzle into the mixing section
where its energy is imparted to the secondary fuel which enters the inlet causing the secondary
fuel to accelerate. After mixing the combined flow is partially decelerated in the diffuser and
static pressure is recovered.” 11 Before travelling to the ejectors the primary fuel flows through a
filter element that is replaced on the 1A inspection every 100 hours.
Ejector pump No. 1 is mounted forward of, and pumps fuel directly into, the collector tank.
Ejector pump No. 2 is mounted at the outer end of the inboard tank, and also pumps fuel
directly into the collector tank. Ejector pump No. 3 is mounted at the inner end of the outboard
tank, and pumps fuel into the inboard tank.
Ejector pump No. 1 is larger and has a greater secondary flow capacity than ejector pumps nos.
2 and 3. The flow rates of ejector pump nos. 2 and 3 are the same. Fuel also flows via gravity
through “clack” valves from the outboard tank into the inboard tank, and also from the inboard
tank into the collector tank.
11
CASA C-212-200 Maintenance Manual 28-20-10, p. 17, paragraph K.
-14-
Fuel system components including the ejectors are subject to a “visual inspection for security
and condition” every 4 years. Removal and disassembly is not specified. Performing the “Basic
Fuel Distribution System – General Check” 12 may detect problems with the No. 3 ejector pump.
This check does not have a specific interval and is performed as an “Indication Check” or
“Booster Pump Operational Check”. However, there is no specific test to check the delivery rate
of the No. 2 ejector pump.
1.6.6
Fuel Distribution
Normal fuel consumption of each engine during cruise was approximately 350 pounds per hour
(pph). Calculations and test cell data showed that, at 3000 feet asl and 2 °C and operating at
100% torque and 100% rpm, fuel flow for one engine would have been 500 pph. Additional data
provided by the engine manufacturer indicates that the minimum fuel flow to maintain 100%
rpm at minimal torque would have been 193 pph, although the engine would not be producing
an appreciable amount of thrust in this condition. With fuel flow reduced to 171 pph, propeller
rpm would begin to decay. 13
Takeoff fuel was 3569 pounds of usable fuel. With a standard 50 pound fuel consumption for
start, taxi, and takeoff, the ramp fuel before engine start would have been 3619 pounds.
The investigation could not determine whether the fuel was evenly distributed between the
wings. Consequently, it is assumed that this was the case and that the usable fuel in each wing
at engine start was 1809.5 pounds.
The totalizer recorded that the left engine consumed 1271 pounds of fuel from engine start to
the time the left engine lost power. The difference between fuel at engine start and when the left
engine lost power, 538.5 pounds, is calculated to be the remaining fuel in the left wing at the
crash site.
The time from when the right engine lost power until the left engine lost power was 14 minutes
5 seconds. At 500 pph, the fuel consumed by the left engine during this period would be
117.5 pounds, which, when added to the 538.5 pounds fuel remaining in the left wing resulted
in a fuel load of 656 pounds of usable fuel in the left wing at the time of the right engine power
loss.
The totalizer recorded that the right engine consumed 1100 pounds of fuel from engine start to
the time the right engine lost power. The difference between fuel at engine start and when the
right engine lost power, 709.5 pounds, is calculated to be the remaining fuel in the right wing at
the crash site.
A sample of fuel obtained from the inboard left wing fuel tank was tested and determined to
meet the requirements of the ASTM D1655 and CGSB 3.23-2009 Jet A-1 specifications for
density, flash point, freezing point and distillation properties. Visual examination of the fuel
sample found it was clear and free of undissolved water. Density of the fuel sample was
measured at 825.6 kg/m3 at 15 °C, which was calculated to be equal to 1.8397 pounds per litre at
12
13
CASA C-212-200 Maintenance Manual 28-20-10, pp. 601–604.
Estimated by engine manufacturer based on typical engine performance data.
-15-
2 °C. Based on this density, the volume of the 656 pounds usable fuel in the left wing, when the
right engine lost power, was calculated to be 356.5 litres.
1.6.7
Aircraft Performance
Takeoff weight for the occurrence flight was 15 671 pounds; the maximum take-off weight is
16 976 pounds. The centre of gravity was at moment 235.48, within the limit of 228.83 to 240.91.
As described in section 1.6.3, the aircraft was substantially modified for geophysical surveying.
The modifications resulted in substantial changes in aircraft performance. Performance
requirements drafted specifically for the project included climb requirements for 100 feet per
minute (fpm) 1000 feet above the take-off surface in the take-off configuration with one engine
inoperative, and 50 fpm in the enroute configuration with one engine inoperative.
Minimum control speed in the air (Vmca) is the minimum flight speed at which the aircraft is
controllable, with a maximum 5° bank angle, when one engine is made suddenly inoperative
and the other engine is at takeoff power. For the as-manufactured configuration of the CASA C212-CC40, Vmca is 85 knots calibrated airspeed (KCAS). With the aircraft modified to the EM
survey configuration, supplemental type approval SA86-9 specifies Vmca is 90 KIAS.
Following incorporation of modification SA86-9, C-FDKM was subjected to flight testing to
verify it met the standard specified in Transport Canada type certificate A-147. The flight testing
was conducted with a cradle for one bird mounted on the exterior of the rear cargo door, but
with no bird installed. The flight testing determined the modification had no significant effect
on stall speeds published in the CASA airplane flight manual (AFM), and also that aircraft
controllability was satisfactory with one engine inoperative at the SA86-9 specified Vmca of 90
KIAS.
The multiple engine failure procedure (Appendix E) includes a chart listing the best glide speed
for various weights, with flaps to be kept up until absolutely necessary. The aircraft weight at
the time of the forced landing was 13 350 pounds, for which the best glide speed is 98 KIAS.
The stall speed (Vs) for the as-manufactured CASA C-212-CC40 at 13 350 pounds with wings
level is 70 KCAS with flaps at 25% and 76 KCAS with flaps up. With 30° bank, Vs increases to
76 KCAS with flaps at 25% and 82 KCAS with flaps up. The AFM supplement for the
SA86-9 EM modification specifies that the stall speed chart in the CASA AFM is still applicable.
The SA86-9 AFM supplement includes a chart for one engine inoperative net rate of climb with
the inoperative propeller feathered, flaps at 25%, operating engine at maximum continuous
power, and 99 KIAS. With the EM loop on the aircraft, at 2 °C, 2500 feet pressure altitude, and
13 350 pounds, the chart shows the rate of climb to be 300 fpm. The LSTA O-LSA-96-0139/D
AFM supplement does not specify any performance decrement, despite the addition of a second
externally mounted sensor bird which would increase drag even in the stowed position.
Flap 25% is specified for one engine inoperative flight, but flaps up is specified for multiple
engine inoperative flight. During certification testing of the modified aircraft, one engine
inoperative climb performance was better with 25% flap at 99 KIAS (V2) than with flaps up at
112 KIAS. Therefore, the certificate applicant chose to use the 25% flap configuration for the one
engine inoperative climb.
-16-
Aircraft performance with the birds deployed was not considered at the time modification
SA86-9 and Canadian type certificate A-147 were approved because the expectation was that, in
the event of an emergency, the birds would be expeditiously recovered or jettisoned. When
deployed, the birds create additional drag which reduces climb rate. There is no information
available for one engine inoperative climb performance with either one or both birds deployed.
Analysis of radar data showed the climb rate immediately following the right engine power loss
averaged 79 fpm from 1814:20 until 1818:07, when the bird recovery commenced. Bird recovery
was complete by about 1823:30, after which the climb rate averaged 95 fpm.
1.7
Meteorological Information
The 1800 aviation routine weather report for Saskatoon was as follows: wind 260° true at
12 knots, visibility 15 statute miles with a few high clouds, temperature 3 °C, dew point -1 °C,
altimeter setting 29.75 inches of mercury. At survey altitude, outside air temperature was about
2 °C with a westerly wind at about 20 knots.
1.8
Flight Recorders
Existing regulations did not require the aircraft to be equipped with flight data or cockpit voice
recorders, and no recorders were installed.
1.9
Wreckage and Impact Information
1.9.1
Wreckage Examination
The aircraft was facing north straddling the concrete noise abatement wall on the east side of
Wanuskewin Drive. The left wing tip was resting in the northbound curb traffic lane. The
cockpit was on the west side of the wall with the wing centre section, aft fuselage and
empennage resting on top of the wall. The right wing was on the east side of the wall resting on
the ground close to a residential fence line. The bottom of the fuselage and cabin floor had been
torn from the aircraft from just aft of the nose gear to the rear cargo door.
The aircraft impacted the wall in a nose level pitch attitude with a slight right bank. The aircraft
came to rest in about 98 feet after first contact with the wall. The majority of the impact energy
was dissipated by the destruction of the aircraft’s lower fuselage, floor structure and the wing
centre section.
The emergency locator transmitter (ELT) was located in the wreckage with the antenna severed;
no ELT signal had been received. The left main landing gear leg was located on the median of
the road approximately 30 feet ahead and 60 feet to the left of the aircraft.
The vertical and horizontal stabilizers and all control surfaces were intact. The left and right
wings received light damage. The right flap had been driven up by contact with the wall and
the left flap was partially deployed. Due to the disruption of the flap linkages in the wing centre
section it was not possible to ascertain what the flap position was prior to contact with the wall.
-17-
The propellers were in the feathered position and showed no signs of rotation at impact. One
blade of the right propeller had fractured on contact with the wall.
No fuel leaks were evident. All fuel caps were in place. Fuel was visible in each inboard tank.
Because each wing sloped downward from the wall to the ground, fuel was slowly migrating to
the outboard tanks. After fuel samples were taken from each wing tank, fuel was pumped into a
recovery tanker to facilitate wreckage removal.
All 4 engine fire extinguisher guards were up, although the fire extinguishers had not
discharged. All 4 fuel booster pump switches were in the on position. The left fuel valve switch
was in the open position, and the right fuel valve switch was in the closed position.
Examination revealed that the position of the electric firewall fuel valves matched the cockpit
switch selections. The main batteries that were mounted in a box behind the cockpit bulkhead
on the right side cabin floor were found broken apart in the wreckage. The auxiliary battery was
found broken apart approximately 30 feet in front of the aircraft.
The wreckage was recovered to a hangar for closer examination. Fuel could be drawn by
suction from the right engine firewall fitting. Similar attempts on the left firewall were
unsuccessful. Examination of the left wing inboard fuel tank revealed that the internal
plumbing connections at the booster pump check valves were lockwired, but were not
completely tight. After cutting the lockwire and tightening these fittings approximately one
wrench flat each, it was possible to draw fuel by suction from the left engine firewall. All fuel
valves, controls, and fittings were examined and found to be functional. The fuel system pipes
and hoses were examined internally from the fuel tank outlet to the engine and were found to
be free from obstruction. The ejector pump filters were removed from each wing and were
found to be installed correctly and free of contaminants.
1.9.2
Engine Examination
Both Garrett (Honeywell) TPE331-10R-511C engines were removed from the aircraft for further
examination.
1.9.3
Right Engine
The right engine (serial number P37186) had accumulated approximately 6470 hours since
overhaul. During engine removal, the chip plug was removed from the reduction gear housing
and was found to be contaminated with metal particles. Initial continuity tests of the chip plug
showed that the contact was not electrically bridged to ground. The engine oil filter was
removed and no flakes or chips of metal were evident.
The last SOAP sample was taken in December 2010, 100 flight hours prior to the accident, and
no abnormal readings were identified.
Examination of the engine logbooks revealed an event in January 2009, approximately 840
operating hours prior to the accident, in which the right engine was intentionally shut down in
flight to facilitate flight testing after a modification to the geophysical survey equipment. The
outside air temperature was approximately -20º C. During the restart attempt the engine was
slow to rotate and would not accelerate above 18% rpm. The start attempt was terminated and
the flight was concluded with one engine inoperative. A manual air start should be performed
-18-
before the oil temperature cools to 4º C. 14 The oil temperature during the start attempt was not
recorded.
1.9.4
Right Engine Torque Sensor
Removal of the reduction gear housing from the
engine revealed that the intermediate spur gear
(part number 896884-1) had failed and was no
longer keyed to the torque sensor shaft (Photo 4).
The intermediate spur gear transmits power, via
the torque sensor shaft, from the high speed
main shaft to a gear train which drives the high
pressure fuel pump, fuel control unit, and oil
pressure and scavenge pumps.
The torque sensor had accumulated
approximately 6470 hours since overhaul in 1997.
The torque sensor assembly does not have a
designated overhaul period and is normally
overhauled with the engine. The torque sensor
Photo 4. Failed gear on the torque sensor shaft
was removed and visually examined by the
manufacturer under the supervision of TSB investigators. The TSB Laboratory conducted a
detailed examination of the torque sensor, as well as gear fragments, metal shavings and chips
recovered from the reduction gear housing. See TSB Laboratory report LP 055/2011 –
Metallurgical Examination of Torque Sensor.
The laboratory analysis of the failed intermediate spur gear revealed fatigue cracks in the roots
of most of the gear teeth. The gear had separated into several fragments as a result of fatigue
cracking followed by overstress fractures. The woodruff key (part number MS35756-1) that
indexed the gear to the torque sensor shaft had fractured in a fatigue mode followed by
overstress fracture. The case hardening of the intermediate spur gear and one other spur gear
(part number 3103589-1) in the gear train was slightly below the specified level on the tooth
flanks and significantly lower in the tooth root areas. Both gears exhibited significant wear on
the loaded faces and flanks of the gear teeth (Photo 5 and Photo 6).
14
CASA C-212 Airplane Flight Manual, p. 2-11.
-19-
Photo 5. Tooth from gear 896884-1 showing wear
on the loaded face and flank
Photo 6. Gear 3103589-1 tooth wear on the loaded
face and flank
The overhaul manual for the torque sensor states that “dimensional, magnetic particle, or
fluorescent penetrant inspection procedures are required only when a visual inspection
indicates wear or damage that is detrimental to the function of the parts.” 15 A review of records
provided by the overhaul agency indicated that the failed gear had not been subjected to
magnetic particle or fluorescent penetrant inspection.
1.9.5
Left Engine
The left engine (serial number P37046) had accumulated approximately 1805 hours since
overhaul. Visual examination did not reveal any faults that would explain a power loss. Engine
testing, using a test cell, was conducted by the manufacturer under TSB supervision. The engine
performed normally as an assembly. The fuel control unit and fuel pump were then removed
and tested individually with no faults detected. The solenoid fuel shutoff valve was removed
and functionally tested with no faults detected. Subsequent disassembly of the valve revealed
no defects.
1.9.6
Left Wing Fuel System Examination
The left wing outboard of the engine nacelle was moved to the TSB regional wreckage
examination facility in Winnipeg, Manitoba. The interior of the fuel tanks was examined and
several discrepancies were identified.
The plastic tubing that delivered fuel from the No. 2 ejector pump was pinched and chafed at
several points where the tubing passed through wing rib stringer cut-outs. The plastic tubing
transitioned to an aluminum tube (part number 212-54227.3) at wing station 4300. Three
aluminum tabs were welded to the tube to facilitate securing the assembly to the wing structure
and the collector tank. Two of these tabs were broken, including the uppermost tab which was
bolted to the forward face of the collector tank. The head of the bolt had chafed a small hole in
the tube. The tube was also heavily chafed at the point where it entered the collector tank. The
nature of these discrepancies indicate that they had developed over many hours of operation
and predated the accident.
15
Allied Signal Aerospace Company Overhaul Manual, Garrett PN 3101726, 72-00-17, p. 302.
-20-
During the in-situ examination of the No. 2
ejector pump, a piece of flexible foreign object
debris (FOD) fell out of the pump intake port.
The piece of L-shaped FOD was approximately
2 mm thick, 4 mm wide and 13 mm long.
Disassembly and examination of the No. 2
ejector pump nozzle revealed a stain
approximately 4 mm wide (Photo 7).
Analysis of the FOD by the TSB Laboratory
revealed that it was composed largely of a
structure of polymeric materials. For
comparison purposes, samples of sealant were
removed from various locations in the fuel tank.
Unlike the FOD, these samples were
determined to be homogenous, particulatefilled polysulphide compounds.
1.9.7
Left Wing Fuel System Testing
Photo 7. No. 2 ejector nozzle and foreign object
debris
The testing protocol was developed during initial sessions in August and November 2011. The
final round of testing was carried out in January 2012. The primary aim of the testing was to
determine the ability of the fuel system to deliver fuel to the firewall in uncoordinated flight
with various amounts of left bank.
A bank into the operating engine with uncompensated yaw results in uncoordinated flight. This
will cause the fuel to move outboard in the low wing, away from the collector tank clack valves
and the No. 1 ejector pump inlet. This fuel movement will be exacerbated because the bottoms
of the fuel tanks are flat.
The testing also examined
•
•
•
•
•
the effect of the chafed hole in the No. 2 ejector pump delivery pipe;
the effect of the loose booster pump plumbing fittings;
the ability of the No. 2 ejector pump to ingest the FOD into the pump nozzle;
the effect of the FOD on the No. 2 ejector pump delivery rate;
the functionality of the low fuel level and low fuel pressure warning switches and lights.
The wing was placed on a trestle to enable the wing to be tilted to various angles to simulate
uncoordinated flight with 3°, 5°, and 7° of left bank. Royco 950 calibration fluid was used for the
testing because of its low volatility and similar density to Jet A fuel.
The volume of usable fuel in the left wing at the time of the right engine power loss was
determined to be approximately 356.5 litres. The volume of unusable fuel in the left wing was
determined to be 33.5 litres by putting a known amount of fluid in the inboard fuel tank with
the wing level and determining how much remained after pumping out as much fluid as
possible with the booster pumps. The unusable volume of 33.5 litres was added to the 356.5
-21-
litres of usable fuel resulting in a total volume of 390 litres. This quantity of Royco 950
calibration fluid was then pumped into the wing.
The density of the Royco 950 calibration fluid used for testing was 1.68 pounds per litre.
Therefore, 390 litres of Royco 950 weighed 655.2 pounds.
An empty drum on a digital scale was used to permit monitoring of the fluid flow rate
throughout each test run, facilitate returning the fluid to the wing after each test run, and to
keep the starting fluid level of each test run consistent. The ejector pump filter was replaced
prior to commencing the testing.
Each test commenced with the wing level and the collector tank full. Fluid flow from the wing
was controlled at a target rate of 456.6 pph (271.8 litres per hour) which is equivalent in volume
to 500 pph of Jet A. The wing was then tilted to the applicable angle, the booster pump outlet
flow was directed into the empty drum, and the timing commenced. The times at which the
following events occurred were recorded:
•
•
•
•
•
•
Exposure of No. 1 ejector pump inlet
Illumination of low fuel level warning lamp and master caution lamp
Audible booster pump cavitation
Fuel pressure fluctuation reached 15 PSI
Low fuel pressure warning lamp flicker
Low fuel pressure warning lamp steady illumination and illumination of master
caution lamp.
Tests were conducted with the wing in 4 conditions:
A.
B.
C.
D.
No. 2 ejector pump unobstructed
FOD manually positioned in the No. 2 ejector pump nozzle
FOD ingested by the No. 2 ejector pump while in the wing
FOD ingested by the No. 2 ejector pump while in the bench test rig.
When tested while installed in the wing, the No. 2 ejector pump (with the FOD removed) was
found to flow a total of 415 pph to the collector tank. Of this, 116 pph was primary flow from
the booster pump(s) needed to operate the ejector pump, resulting in a net fuel delivery to the
collector tank of 299 pph. The primary fuel from the booster pump(s) to operate the No. 3
ejector pump was drawn from the collector tank and discharged into the inboard tank at a rate
of approximately 138 pph. Consequently, when the No. 1
ejector pump inlet was not submersed, the resulting net
flow to the collector tank was 161 pph (95.8 litres/hr) of
Royco 950 equivalent in volume to 176 pph of Jet A.
The No. 2 ejector pump (with the FOD removed) was
removed from the wing and tested as described in the
CASA C-212 Component Maintenance Manual 28-13-04 and
was found to meet the specified delivery requirements.
During the last series of bench tests, the FOD was
positioned near the intake and was ingested into the pump.
During the last ingestion test, the output flow decreased
Photo 8. Foreign object debris as
ingested into No. 2 ejector pump
during bench testing
-22-
substantially as a result of the FOD lodging deep in the ejector nozzle (Photo 8). The pump was
reinstalled in the wing in this condition (Appendix H) and 4 additional test runs were
conducted.
All testing in January 2012 was performed with the leaks present at the booster pump check
valves as found at the crash site. Earlier testing had shown that there was no significant
decrease in the duration of fuel supply as a result of the loose fittings. Visual examination
confirmed that the leaks were minor and that the fluid that leaked from the fittings tended to
drip close to the No. 1 ejector pump inlet and was returned to the collector tank.
The chafed hole observed in the No. 2 ejector pump delivery pipe leaked fluid at a rate of 1.29
litres per hour. This leak rate over 14 minutes 5 seconds would not have had any material
impact on engine operation.
During preparation for a test on 17 August 2011, the low level sensor did not operate during the
filling of the collector tank. The low level warning lamp remained on for approximately 8
minutes, well after the collector tank was full. The top wing skin above the low level sensor was
rapped sharply and the lamp extinguished.
From the series of tests (Appendix H), the following observations were made:
•
•
•
•
•
A bank angle of 3° provided fluid flow to the engine of 500 pph for over 50 minutes.
At a bank angle of 3°, the No. 1 ejector intake became exposed with
238 (204.5 usable) litres of fluid remaining in the wing.
A bank angle of 5° provided fluid flow at 500 pph for approximately 33 minutes.
A bank angle of 7° provided fluid flow at 500 pph for approximately 25 minutes.
At a wing angle greater than 5.5°, the No. 1 ejector inlet became exposed with
390 (356.6 usable) litres of fluid remaining in the wing.
With the FOD in the No. 2 ejector pump, the following observations were made:
•
•
•
•
•
When testing at a 5° wing angle, the No. 1 ejector pump inlet, even though exposed,
continued to suck up some fluid that remained in the bay below the inlet until late into
the test.
When the wing was tilted to 5.8° no fuel remained in the bay below the No. 1 ejector
pump inlet.
A wing angle of 5° provided fluid flow to the engine of 500 pph for approximately 27 to
28 minutes, regardless of the FOD’s location in the nozzle.
A wing angle of 7° reduced fluid flow endurance at 500 pph to approximately 13 to 25
minutes, depending on the location of the FOD in the ejector pump nozzle.
In condition D, usable fluid remaining in the wing after depletion of the collector tank
varied from 233 litres (runs 18 and 20) to 286 litres (run 17).
Additional observations:
•
The No. 1 ejector pump has sufficient delivery capacity to maintain the collector tank in
an overflow condition while supplying 500 pph fuel flow to the engine.
-23-
•
•
•
•
1.9.8
The No. 2 and No. 3 ejector pump system does not have sufficient capacity to prevent
eventual depletion of the collector tank once the No. 1 ejector pump inlet becomes
exposed.
The primary flow from the No. 3 ejector pump does not return to the collector tank.
During the occurrence flight, the low level warning likely illuminated approximately
3 to 4 minutes after the right engine lost power (about 10 minutes prior to collector tank
depletion).
During the testing, illumination of the low fuel pressure light and master caution
annunciator was coincident with a substantial decrease in fluid flow.
Engine Flameouts
The engine ignition system is in operation only during the starting sequence. Once started,
combustion is continuous and self-sustaining as long as the engine is supplied with the proper
fuel-to-air ratio.
Flameout is a term used to describe the condition in which combustion in a gas turbine engine
unintentionally stops. This may be due to unusual attitudes, a malfunctioning fuel control
system, blocked fuel supply, air introduction into the fuel delivery system, turbulence, icing,
fuel starvation or fuel exhaustion.
The TPE-331-10R-511C engine, as installed on the CASA C-212, is not equipped with an
auto-ignition system, so if a flameout occurs, the engine will not automatically restart if the
fuel-to-air ratio is restored to normal operating parameters.
1.9.9
Annunciator Panel
The 30-segment annunciator panel was mounted in the glare-shield below the front windscreen.
The segments are arranged so as to group systems together, arranged in alternating columns for
the left and right engines (Figure 2). Each annunciator panel segment contains 2 lamp bulbs.
Both bulbs in the fuel level–left tank warning segment were found to be serviceable when
examined in Saskatoon.
Figure 2. Annunciator panel
The annunciator panel incorporates a 2-position switch to alter the brightness of the panel
between dim and bright to suit ambient conditions. The switch was damaged during the crash
and it was not possible to determine the switch selection position or serviceability.
One lamp in each of the 4 following segments was unserviceable: right generator, right fuel
level, and left SRL/start. Both lamps in the right fuel pressure segment were unserviceable.
The panel was sent to the TSB Laboratory for lamp filament analysis. When the bulbs from the
fuel level left tank segment were examined, the filament in one lamp was found to be fractured.
This likely occurred during transit. When the power is removed from a PN 327 lamp, it takes
approximately 50 milliseconds to cool. Analysis of the crash revealed that the deceleration from
-24-
90 knots in 98 feet occurred in approximately 1.25 seconds. Since hot lamp filaments stretch, and
the occurrence lamps had time to cool, it could not be confirmed that they were on at the time of
impact. The main batteries and auxiliary battery were destroyed early in the crash sequence.
A master caution annunciator was located on each pilot’s instrument panel below the glareshield. Approximately 5 seconds after the illumination of a segment in the annunciator panel
both master caution annunciators illuminate but do not flash. Either pilot may press the
annunciator to acknowledge the warning and extinguish both master caution annunciators.
The annunciator panel, master caution annunciator system, interconnecting components and
wiring were examined and no reason was identified for the display of any warnings to have
been prevented.
1.10
Fire
There was no in-flight or post-crash fire.
1.11
Survival Aspects
The survey equipment operator was found in the seat with the seatbelt still fastened and
attached to the seat. The seat and operator were on the ground outside of the fuselage on the left
side of the noise abatement wall. The seat legs had been fastened to a track in the floor on the
right side of the cabin centre section. The structural integrity of this area of the aircraft was
completely disrupted by the collision with the noise abatement wall, and the resulting forces
were not survivable.
1.12
Tests and Research
The following TSB Laboratory analysis was completed:
LP 039/2011 – Fuel Analysis
LP 042/2011 – Lamps & Fuel Data Analysis
LP 055/2011 – Metallurgical Examination of Torque Sensor
LP 151/2011 – Wire Examination
LP 013/2012 – Metallurgical Analysis of FOD
The reports are available from the TSB upon request.
1.13
Organizational and Management Information
The flight was operating under Subpart 702 - Aerial Work of the Canadian Aviation
Regulations.
1.14
Additional Information
Police and firefighters arrived at the crash site within 2 minutes, followed shortly afterward by
paramedics in ambulances, who transported the survivors to hospital for medical attention. The
flight crew and first responders reported an odour of fuel present on the left side of the aircraft.
There was no post-impact fire.
-25-
1.15
Useful or Effective Investigation Techniques
Not applicable for this occurrence.
-26-
2.0 Analysis
Weather at the time of the occurrence and air traffic control involvement were not factors in the
occurrence. The analysis will examine the reasons for the engine power losses and the crew
responses to them.
2.1
Right Engine Power Loss
TSB Laboratory analysis indicated that the case hardening of the gear tooth flanks and roots of
2 spur gears in the torque sensor gear train was below the manufacturer’s specification
requirements and likely led to the wear of the loaded faces and flanks of the gear teeth. The
combined wear of the 2 gears likely caused an abnormal vibration that produced excessive
cyclic loading and eventual fatigue cracking in the tooth roots of the intermediate gear. The
intermediate spur gear subsequently separated into several fragments and caused the loss of
power transmission to the high-pressure engine-driven fuel pump. The immediate result would
have been fuel starvation of the engine, flameout and the loss of power.
If no detrimental wear or damage had been observed when the torque sensor was overhauled,
there would have been no requirement for non-destructive inspection. Since it is not known
how fast the gear tooth wear or fatigue cracks propagated, it is not certain that routine
application of non-destructive testing during the 1997 overhaul would have revealed any
defects.
In January 2009 a failed air start attempt was experienced at an outside air temperature of
approximately -20 º C. Because it could not be determined whether the engine oil had cooled
below the 4° C minimum temperature allowed prior to the restart attempt, and the engine had
since operated 840 hours without further incident, the investigation could not determine what
impact, if any, this event had on the failure mode of the intermediate spur gear.
The shudder felt was almost certainly associated with the gear failure. The engine-driven fuel
pump was no longer operating so any attempt to restart the engine would have been
unsuccessful.
When the right engine chip plug was removed it did not initially display electrical continuity so
the chip warning light did not illuminate. The metal chips observed on the plug likely fell onto
the plug during the recovery of the aircraft wreckage.
2.2
One Engine Inoperative Performance
The investigation determined that the aircraft was flown with at least an average 5° left bank
after the first engine power loss. However, the bank angle would have been greater than 5° as
the flight made several left turns toward Saskatoon, and was less than 5° during some shallow
right turns to position the aircraft on final approach. The variability of the bank angles was not
considered to have significantly affected the aircraft’s performance. However, variable left bank
angles with left yaw could produce lateral accelerations that tend to move the fuel outboard in
the left wing fuel tank.
-27-
Aircraft climb performance was marginal following the right engine loss of power. Factors
which likely impaired climb performance are as follows:
•
•
•
2.3
Both birds were deployed at the time and remained deployed for 3 to 4 minutes after the
right engine power loss.
Loss of thrust from the right engine combined with drag from the deployed birds
resulted in the airspeed deteriorating below V2 (99 KIAS).
After the initial deceleration, the aircraft slowly accelerated.
Bird Recovery
For about 3 minutes after the right engine lost power, the aircraft operated in a high drag
configuration with both birds deployed. The reason for this was that the pilots were initially
uncertain whether the aircraft would maintain altitude, and recovering or jettisoning the birds
required the survey equipment operator to leave his seat to do so. This resulted in reduced
climb performance with one engine inoperative. Because the terrain was flat, this did not pose a
hazard. Over terrain with more vertical relief, the reduced climb performance could present a
controlled flight into terrain hazard.
2.4
Fuel System
Because the bottom of the wing has no dihedral, the fuel system depends on the ejector pumps
to move fuel from outboard to inboard and into the collector tank. This dependence is increased
in uncoordinated flight, such as single-engine flight after an engine power loss.
The usable fuel level in the left wing at the time of the right engine failure was approximately
656 pounds. At that fuel level, the No. 1 ejector pump inlet would have been exposed if the
bank angle in uncoordinated flight exceeded 5.8 degrees. In this condition, the No. 2 ejector
pump would have been the only means of fuel transfer to the collector tank.
Without foreign object debris (FOD) present, the depletion rate of the collector tank was
approximately 324 pph. This is the difference between the engine fuel consumption of 500 pph
at full power and the net fuel transfer rate of 176 pph to the collector tank from the No. 2 and
No. 3 ejector pumps. With no input from the ejector pumps, the collector tank should have been
able to supply fuel for almost 19 minutes.
With the FOD present, and depending on the position of the FOD in the No. 2 ejector pump
nozzle, testing demonstrated that the ability of the collector tank to supply fuel was
approximately 13:21 to 16:03 minutes. The length of time between the right engine power loss
and the left engine power loss was 14:05 minutes.
During normal coordinated flight the No. 1 ejector pump, because of its larger capacity, can
maintain the collector tank in an overflow condition; consequently, there would be no
indication to the crew that the performance of the No. 2 ejector pump had been compromised.
Analysis of the FOD material revealed that it was dissimilar to the type of sealant material
utilized in the construction, repair or maintenance of the fuel tank, but it could not be
determined how or when the FOD was introduced into the fuel tank. The staining of the No. 2
ejector pump nozzle indicates that it had been present for some time prior to the accident flight.
-28-
The origin of the materials in the FOD was not determined. The inlets of the ejector pumps are
unscreened; consequently, FOD of a compatible size and shape present in the fuel tank could be
ingested into the nozzle area of the ejector pump.
The booster pump check valve fittings were not completely tight and there was a small hole
chafed in the No. 2 ejector pump delivery pipe. These defects did not appreciably diminish the
ability of the collector tank to supply fuel to the engine. However, if the engine had been
running with the boost pumps inoperative, the loose check valve fittings would have allowed
air to be drawn into the fuel system. The pinching of the No. 2 ejector pump plastic delivery
tube at the wing stringer cut-outs did not affect the fuel delivery rate, but if the associated
chafing had resulted in the tubing being pierced, fuel transfer would have been limited. Finally,
the cracking of the welded tabs on the No. 2 ejector pump delivery pipe and subsequent chafing
where the pipe entered the collector tank would have limited the transfer of fuel if it had been
allowed to progress.
2.5
Low Fuel Level Warning
During TSB testing of the fuel system, the low level float sensor operated correctly and
illuminated the “fuel level–left tank” annunciator every time fluid was being pumped out of the
collector tank. The sensor failed to operate on only one occasion, and this occurred while the
collector tank was being filled. The tests were conducted with the wing stationary in a test jig. In
flight, the system would have been subjected to engine, propeller, and airframe vibration which
would tend to prevent the float sensor from sticking in position.
It is possible that the float stuck during the occurrence flight as the collector tank became
depleted and prevented the illumination of the low level warning lamp. However, this is
considered to be unlikely. On balance, it is likely that the low fuel level warning system was
functioning normally during the occurrence flight.
The left low fuel level warning lamp and the master caution lamp likely illuminated
approximately 3 to 4 minutes after the right engine power loss.
Figure 3. Annunciator panel showing illuminated segments
-29-
There are several reasons why the warning was not perceived by the crew:
•
•
•
•
•
•
The systems-oriented arrangement of the annunciator panel placed the “fuel level left
tank” segment adjacent to segments that were already illuminated as a result of the right
engine power loss and shutdown (Figure 3). Consequently, the illumination of a new
warning light adjacent to those already indicating status of the shut-down engine is less
likely to be perceived or associated with the operating engine.
There may have been no master caution warning associated with the low fuel level if it
occurred before the crew acknowledged and cancelled the master caution warning
triggered by the right engine power loss.
The increasing intensity of sunlight through the cockpit windows as the aircraft turned
to the west may have made the warning hard to see.
The crew was also communicating with Saskatoon air traffic control and attempting to
visually locate the runway several nautical miles ahead.
The left low fuel level warning lamp likely illuminated approximately 3 to 4 minutes
after the right engine power loss, at a time of relatively high cockpit workload.
The master caution annunciator may not have captured the crew’s attention due to the
absence of a flashing function.
The pilots did not execute the fuel level warning checklist because they did not perceive the
illumination of the fuel level left tank warning light. Consequently, the fuel crossfeed valve
remained closed and fuel from only the left wing was being supplied to the left engine.
2.6
Left Engine Power Loss
The left fuel pressure was normal following completion of the engine failure in flight procedure
for the right engine power loss. Therefore, the crew did not inadvertently turn off the left boost
pumps during the right engine power loss procedure. When the before-landing checklist was
performed, the right booster pumps were turned back on. At this time there would have been 2
green lights illuminated on the overhead booster pump control panel indicating that both
auxiliary pumps were selected “on”. Until the captain reached up for the emergency handle
after the loss of power had occurred, his hands were nowhere near the overhead panel.
Therefore, he could not have inadvertently closed the fuel valve on the left engine.
The left engine power loss was likely caused by fuel starvation resulting from operating at
various bank angles and left yaw with a compromised fuel system. This flight attitude exposed
the No. 1 ejector pump, and the reduced output of the contaminated No. 2 ejector pump led to
the depletion of the collector tank.
As the booster pump output pressure and flow decreased, the engine became fuel starved and
experienced a loss of power likely followed by a flameout. Testing of the fuel system
demonstrated that the illumination of the low fuel pressure warning light was probably
coincident with the left engine power loss.
2.7
Flight Following
The SkyTrac system provided timely position information and would likely have assisted
search and rescue personnel if position data had been required. The ELT was damaged and did
-30-
not activate or transmit. Neither system was required because the flight was on air traffic
control radar and its position was known.
2.8
Annunciator Bulb Analysis
Lamp filament analysis performed by the TSB Laboratory indicated that no annunciator lamps
could be confirmed to have been illuminated at the time of impact. However, the main batteries
and auxiliary battery were destroyed early in the crash sequence. Because of the cooling rate of
the lamp filaments and the aircraft’s deceleration time, it is likely that the lamps had time to
cool so the lamp filaments did not distort sufficiently to be confirmed “on” by laboratory
analysis.
2.9
Forced Landing
When the left engine lost power, the aircraft was approximately 3.4 nm from the threshold of
runway 27. The crew immediately determined that it was not possible to extend the glide to the
airport.
The crew had limited altitude and time to prepare for and execute the forced landing. Although
the multiple engine failure procedure specified the flaps should be retracted, the crew elected to
leave them at the current setting of 25%. Had the flaps been retracted, they would have needed
to be re-extended a short time later to prepare for the forced landing, and any improvement in
glide performance resulting from flap retraction would not have been sufficient for the aircraft
to reach the runway or a better landing site than the one chosen
The site that was chosen for the forced landing offered the most likelihood of success with the
least risk to persons on the ground. The crew landed the aircraft under control while avoiding a
busy highway to the left and residential buildings on the right. The concrete noise abatement
wall ran parallel to the roadway and initially would have been difficult to see from the air. The
crew received immediate assistance from bystanders and were aided by a quick response from
Saskatoon emergency services personnel.
2.10
Flight Recorders
The aircraft was not required to be equipped with certified flight recorders. When cockpit voice
and flight data recordings are not available to an investigation, this may preclude the
identification and communication of safety deficiencies to advance transportation safety.
-31-
3.0 Conclusions
3.1
Findings as to Causes and Contributing Factors
1.
The right engine lost power when the intermediate spur gear on the torque sensor
shaft failed. This resulted in loss of drive to the high-pressure engine-driven pump,
fuel starvation, and immediate engine stoppage.
2.
The ability of the left-hand No. 2 ejector pump to deliver fuel to the collector tank was
compromised by foreign object debris (FOD) in the ejector pump nozzle.
3.
When the fuel level in the left collector tank decreased, the left fuel level warning
light likely illuminated but was not noticed by the crew.
4.
The pilots did not execute the fuel level warning checklist because they did not
perceive the illumination of the fuel level left tank warning light. Consequently, the
fuel crossfeed valve remained closed and fuel from only the left wing was being
supplied to the left engine.
5.
The left engine flamed out as a result of depletion of the collector tank and fuel
starvation, and the crew had to make a forced landing resulting in an impact with a
concrete noise abatement wall.
3.2
Findings as to Risk
1.
Depending on the combination of fuel level and bank angle in single-engine
uncoordinated flight, the ejector pump system may not have the delivery capacity,
when the No. 1 ejector inlet is exposed, to prevent eventual depletion of the collector
tank when the engine is operated at full power. Depletion of the collector tank will
result in engine power loss.
2.
The master caution annunciator does not flash; this leads to a risk that the the crew
may not notice the illumination of an annunciator panel segment, in turn increasing
the risk of them not taking action to correct the condition which activated the master
caution.
3.
When cockpit voice and flight data recordings are not available to an investigation,
this may preclude the identification and communication of safety deficiencies to
advance transportation safety.
4.
Because the inlets of the ejector pumps are unscreened, there is a risk that FOD in the
fuel tank may become lodged in an ejector nozzle and result in a decrease in the fuel
delivery rate to the collector tank.
3.3
Other Findings
1.
The crew’s decision not to recover or jettison the birds immediately resulted in
operation for an extended period with minimal climb performance.
-32-
2.
The composition and origin of the FOD, as well as how or when it had been
introduced into the fuel tank, could not be determined.
3.
The SkyTrac system provided timely position information that would have assisted
search and rescue personnel if position data had been required.
4.
Saskatoon police, firefighters, and paramedics responded rapidly to the accident and
provided effective assistance to the survivors.
-33-
4.0 Safety Action
4.1
Action Taken
4.1.1
Fugro Aviation Canada Limited
Fugro grounded its remaining CASA C-212 aircraft immediately following the occurrence.
Before recommencing operations on 30 June 2011, the company
•
revised its CASA C-212 one engine inoperative emergency procedures to include
supplying the operating engine with fuel from both the left and right tanks by opening
the crossfeed valve; and
•
modified the aircraft with a remote-controlled cable cutter on the bird tow cables. This
cutter permits the pilots to jettison the birds from the cockpit, eliminating the
requirement for the survey equipment operator to leave his seat, and allows the pilots to
quickly improve the climb performance of the aircraft in the event of a loss of engine
power.
In October 2011, the aircraft was modified with the installation of a continuous ignition system
for the engines.
Fugro has also increased the frequency and expanded the scope of some maintenance
inspections of the CASA C-212 fuel system, including cleaning of the ejector pump nozzles.
4.1.2
Transport Canada
On 14 April 2011, Transport Canada conducted an inspection of Fugro’s operational control and
maintenance release processes exercised for the occurrence flight. The process inspection was
conducted in accordance with Civil Aviation Staff Instruction SI-SUR-001, Issue 4. The inspection
determined all processes reviewed met applicable regulatory requirements and were being
followed as described in approved company manuals.
4.1.3
Honeywell Aerospace
Honeywell Aerospace has initiated a revision to the component maintenance manual for the
torque sensor. The revision will require a magnetic particle inspection of the spur gear in
addition to visual inspection.
4.1.4
Airbus Military
Airbus Military has initiated a revision to the CASA C-212 Airplane Flight Manual procedure
for engine failure in flight. The revision adds a caution that the following steps must not be
actioned if a fuel leak on the affected side exists, and adds 2 subsequent procedural steps to turn
the normal fuel booster pump for the failed engine “on” and to move the fuel feeding lever to
the “crossfeed” position.
-34-
This report concludes the Transportation Safety Board’s investigation into this occurrence. Consequently,
the Board authorized the release of this report on 12 December 2012. It was officially released on
08 January 2013.
Visit the Transportation Safety Board’s website (www.bst-tsb.gc.ca) for information about the
Transportation Safety Board and its products and services. You will also find the Watchlist, which
identifies the transportation safety issues that pose the greatest risk to Canadians. In each case, the TSB
has found that actions taken to date are inadequate, and that industry and regulators need to take
additional concrete measures to eliminate the risks.
-35-
5.0 Appendices
Appendix A -Flight Path 1814 to 1825:36
-36-
Appendix B – Engine Failure in Flight Procedure
-37-
Appendix C – Flight Path 1825:36 to 1830:02
-38-
Appendix D – Before Landing Check
-39-
Appendix E – Multiple Engine Failure and Emergency Landing Procedures
-40-
Appendix F –Fuel Level Warning Procedure
-41-
Appendix G – Fuel Pressure Warning Procedure
-42-
Appendix H – Left Wing Fuel System Testing Results
Table H1. Condition A: Number 2 pump clean
Test
run
number
Date of
test run
Recorded
fuel
pressure
Recorded
primary
pressure
Pounds per square
inch
21
14
1
3-Jan-12
2
3-Jan-12
21
14
3
11
12
13
14
4-Jan-12
4-Jan-12
4-Jan-12
5-Jan-12
5-Jan-12
21
21
21
21
21
14
14
14
14
14
Flow rate
(pounds per
minutes)
Wing
angle
(degrees)
Exposure
of #1
ejector
Illumination
of low-level
warning
indicator
Audible
pump
cavitation
Pressure
fluctuation
Flickering
of lowpressure
lamp
Illumination of
low-pressure
caution lamp
Time of onset from beginning of test (minutes:seconds)
15.4/2:00
84.2/not
recorded
30.4/4:00
143.6/19:00
76.6/10:00
77.4/10:00
100/13:00
3
33:24
37:14
n/a
n/a
n/a
n/a
5
9:11
13:25
30:49
30:54
n/a
31:11
6.8
7
7
5
5
1:02
00:56
00:55
9:29
9:30
5:17
5:46
5:42
13:58
14:04
22:34
24:02
23:39
32:05
32:20
22:56
24:49
24:29
32:42
32:43
n/a
25:06
24:45
32:58
33:01
23:12
25:19
24:59
33:12
33:17
-43-
Table H2. Condition B: Number 2 pump with foreign object debris positioned manually
Test
run
number
Date of
test run
Recorded
fuel
pressure
Recorded
primary
pressure
Flow rate
(pounds per
minutes)
Wing
angle
(degrees)
Exposure
of #1
ejector
Pounds per square
inch
Illumination
of low-level
warning
indicator
Audible
pump
cavitation
Pressure
fluctuation
Flickering
of lowpressure
lamp
Illumination of
low-pressure
caution lamp
Time of onset from beginning of test (minutes:seconds)
4
4-Jan-12
21
15
23.4/3:00
3
29:40
36:23
49:35
49:40
n/a
50:31
5
4-Jan-12
21
15
152.4/20:00
5
8:05
12:16
25:25
25:33
25:48
27:45
6
4-Jan-12
21
15
61/8:00
7
0:58
3:20
12:36
12:59
13:09
13:21
7
4-Jan-12
21
15
38.2/5:00
5
8:21
12:46
25:36
26:01
26:19
28:22
8
4-Jan-12
21.5
15
75.8/10:00
5
8:21
12:41
25:12
25:47
26:06
27:54
9
4-Jan-12
21
15
74.2/10:00
7
0:58
3:21
12:33
13:01
13:09
13:23
10
4-Jan-12
21
15
61.2/8:00
7
0:58
3:26
12:42
13:20
13:29
13:34
Pressure
fluctuation
Flickering
of lowpressure
lamp
Illumination of
low-pressure
caution lamp
Table H3. Condition C: Number 2 pump with foreign object debris ingested in wing
Test
run
number
Date of
test run
Recorded
fuel
pressure
Recorded
primary
pressure
Pounds per square
inch
Flow rate
(pounds
per
minutes)
Wing
angle
(degrees)
Exposure
of #1
ejector
Illumination
of low-level
warning
indicator
Audible
pump
cavitation
Time of onset from beginning of test (minutes:seconds)
15
5-Jan-12
21
14
30.4/4:00
7
0:53
5:34
23:21
24:15
24:30
24:46
16
5-Jan-12
21
14
14.8/2:00
7
0:59
5:43
23:39
24:29
24:45
24:59
-44-
Table H4. Condition D: Number 2 pump with foreign object debris ingested during bench test
Test
run
number
Date of
test run
Recorded
fuel
pressure
Recorded
primary
pressure
Pounds per square
inch
Flow rate
(pounds
per
minutes)
Wing
angle
(degrees)
Exposure
of #1
ejector
Illumination
of low-level
warning
indicator
Audible
pump
cavitation
Pressure
fluctuation
Flickering
of lowpressure
lamp
Illumination of
low-pressure
caution lamp
Time of onset from beginning of test (minutes:seconds)
17
6-Jan-12
21
14
30/4:00
7
0:56
3:47
14:31
15:12
15:25
15:38
18
6-Jan-12
21
14
30.2/4:00
5
8:47
12:43
25:18
25:59
26:21
27:12
19
6-Jan-12
21
14
14.5/2:00
7
0:53
3:54
14:54
15:39
15:52
16:03
20
6-Jan-12
21
14
45.4/6:00
5
8:47
12:46
25:24
26:06
26:26
27:12
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