Air Carrier, Turbine Powered & Large Aircraft Reports

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National Transportation Safety Board - Aircraft Accident/Incident Database
Accident Rpt# CEN15TA017
10/17/2014 1245
Regis# N852BP
Lordsburg, NM
Acft Mk/Mdl AIRBUS HELICOPTERS AS 350 B3
Acft SN 3592
Acft Dmg: SUBSTANTIAL
Rpt Status: Factual Prob Caus: Pending
Eng Mk/Mdl TURBOMECA ARRIEL 2B
Acft TT
Fatal
Flt Conducted Under: FAR PUBU
Opr Name: US CUSTOMS AND BORDER
PROTECTION
Opr dba:
6081
0
Ser Inj
Apt: N/a
0
Aircraft Fire: NONE
Events
1. Takeoff - Dynamic rollover
Narrative
On October 17, 2014, about 1245 mountain daylight time, an Airbus Helicopters AS350B3 helicopter, N852BP, rolled over on its right side during takeoff near
Lordsburg, New Mexico. The pilot sustained minor injuries and the helicopter sustained substantial damage. The helicopter was registered to the United States
Department of Homeland Security and operated by the United States Customs and Border Protection (CBP) under the provisions of 14 Code of Federal
Regulations Part 91 as a public use flight. Visual meteorological conditions prevailed and no flight plan was filed. The flight was originating at the time of the
accident.
According to the CBP investigators, the helicopter departed from a level terrain area, drifted backward and the aft right skid contacted a rock, which was
embedded in the ground and 10 inches of it was exposed. After contact, the helicopter hinged around a rock, the rotor blades impacted the ground and the
helicopter rolled over to the right side. The fuselage was substantially damaged during the accident sequence.
At 1256, the automated weather observation at the Bisbee Douglas International Airport, Douglas/Brisbee, Arizona, located 38 miles west of the accident site,
reported: calm wind, visibility 10 miles, clear sky, temperature 79ø Fahrenheit (F), dew point 39ø F, and altimeter setting 30.08 inches of mercury.
The CBP investigators examined the wreckage and determined that there were no anomalies with the helicopter that would have precluded normal operation.
Representatives from Airbus Helicopters downloaded and analyzed the data from the Vehicle Engine Multifunction Display (VEMD). All of the recorded failures
were associated with the ground impact. The parameters associated with the failures and times of the failures did not indicate any preimpact anomalies.
Printed: May 01, 2015
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an airsafety.com e-product
Prepared From Official Records of the NTSB By:
Air Data Research
9865 Tower View, Helotes, Texas 78023
210-695-2204 - [email protected] - www.airsafety.com
Copyright 1999, 2015, Air Data Research
All Rights Reserved
Accident Rpt# ERA15LA179
03/24/2015 1730 EDT Regis# N43CM
Taylors Island, MD
Apt: N/a
Acft Mk/Mdl BELL 206B
Acft SN 2025
Rpt Status: Prelim
Eng Mk/Mdl ALLISON 250 C208
Acft TT
Fatal
Flt Conducted Under: FAR 091
Opr Name: DC HELICOPTERS INC
Opr dba:
6678
0
Ser Inj
0
Prob Caus: Pending
Aircraft Fire: NONE
AW Cert: STN
Events
1. Enroute-cruise - Loss of engine power (total)
Narrative
On March 24, 2015, about 1730 eastern daylight time, a Bell 206B, N43CM, was substantially damaged during an autorotation near Taylor Island, Maryland.
The commercial pilot and one passenger were not injured. Visual meteorological conditions prevailed and no flight plan was filed. The helicopter was registered
to and operated by DC Helicopters Incorporated as a personal flight, which was conducted under the provisions of Title 14 Code of Federal Regulations Part
91.
According to the pilot, he was in cruise flight at 3,000 feet above ground level, when there was a sudden drop in altitude followed by a 90 degree rotation to the
left and a loud "bang." The pilot lowered the collective and noted that the torque gauge went below 10 percent. He then rolled the throttle to idle, entered an
autorotation and landed on a shore.
According to the Federal Aviation Administration inspector that examined the helicopter, the tail boom was buckled near the tail rotor gear box. Further
examination revealed that the tail rotor drive shaft was broken. The helicopter was recovered and is awaiting further examination.
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Accident Rpt# CEN14LA390
07/26/2014 1430 CDT Regis# N240SJ
Bennet, NE
Apt: N/a
Acft Mk/Mdl BELL 206B-B
Acft SN 2354
Eng Mk/Mdl ROLLS-ROYC 250-C20B
Acft TT
Fatal
Opr Name: SAMARITAN AIR INC
Opr dba:
7216
0
Ser Inj
0
Aircraft Fire: NONE
Summary
The pilot reported that he refueled the helicopter from a stationary fuel tank that the operator had recently had filled.ÿAfter sumping the helicopter's fuel tank
and the airframe fuel filter and noting that the strained fuel was absent of debris and water,ÿthe pilot departed on the flight.ÿAbout 12 to 15 minutes into the
flight, the engine lost total power. The pilot made a low-altitude autorotation to a field, andÿthe helicopter subsequently slid about 10 to 15 ft and struck a berm.
No preaccident mechanical deficiencies were found that would have precluded normal operation of the engine.
The pilot said that he thought the power loss was due to fuel contamination, so he sumped fuel from the helicopter and the storage tank, and the fuel from each
source was dark brown with visible contaminants. Postaccident examination of a large sampleÿof fuel drained from the storage fuel tank confirmed that the fuel
was brown with visible contaminants. The tank's filter was also found to be contaminated, and sludge was observed in the bottom of the tank. Contaminated
fuelÿwas found in the helicopter's fuel tank and fuel filter bowl, and the fuel nozzle wasÿcovered with a hardened dark brown substance.
Review ofÿfueling records revealed that the fuel contractor had delivered a 1,000-gallon fuel storage tank labeled "diesel" to the fueling site and that the
contractor placedÿ920 gallons of Jet A fuel in the tank 14 days later (5 days before the accident). It could not be determined what type of fuel was in the tank
before it was filled with Jet A fuel.ÿ
Cause Narrative
THE NATIONAL TRANSPORTATION SAFETY BOARD DETERMINED THAT THE CAUSE OF THIS OCCURRENCE WAS: A total loss of engine power due
to fuel contamination, which resulted from the pilot's inadequate preflight inspection. Contributing to the accident was the operator's failure to properly maintain
the fuel storage tank.
ÿ
Events
1. Maneuvering-low-alt flying - Loss of engine power (total)
2. Emergency descent - Loss of engine power (total)
Findings - Cause/Factor
1. Aircraft-Fluids/misc hardware-Fluids-Fuel-Fluid condition - C
2. Organizational issues-Management-Resources-Adequacy of equipment-Operator - F
3. Organizational issues-Support/oversight/monitoring-Documentation/record keeping-Testing records-Operator
4. Personnel issues-Task performance-Inspection-Preflight inspection-Pilot - C
Narrative
On July 26, 2014, at 1430 central daylight time, N240SJ, a Bell 206B helicopter, sustained substantial damage during a forced landing to a soybean field after a
total loss of engine power near Bennet, Nebraska. The commercial pilot was not injured. The helicopter was registered to and operated by a private company.
Visual meteorological conditions prevailed and no flight plan was filed for the local, aerial application flight conducted under 14 Code of Federal Regulations
Part 137.In a telephone conversation, the pilot stated that he had flown 4 uneventful hours that day and had to stop to refuel. He said the operator had just
purchased fuel, which had been placed into a stationary fuel storage tank and this was the first time he was using fuel from that tank. The pilot pumped 25
gallons of fuel into the helicopter, then sumped the airframe fuel tank and fuel filter. The pilot said the strained fuel was absent of water and debris. He then
departed. About 12-15 minutes into the flight, when the helicopter was about 16 feet above the ground, at an airspeed of 80 miles per hour (mph), the
turbine-engine lost total power. The pilot made an autorotation to a soybean field and slid about 10-15 feet into a berm. Due to the rocking motion from hitting
the berm, the main rotor blade severed the tail boom.
The pilot said that he thought the engine "flame-out" was due to fuel contamination so he sumped the helicopter's airframe fuel filter and the stationary fuel
storage tank after the accident. The drained fuel from both sources was dark brown with visible contaminates.
A post accident examination of the helicopter and stationary fuel storage tank was conducted on August 5, 2014, by the Federal Aviation Administration (FAA),
Bell Helicopter and Rolls Royce. Examination of the stationary fuel storage tank's fueling receipt and company records revealed that on July 7, 2014, the
operator's fuel contractor delivered a 1,000 gallon Flameshield tank to the remote fueling site. This tank was marked "Diesel" on the side. On July 21, 2014, the
fuel contractor placed 920 gallons of Jet A fuel into the 1,000 gallon "Diesel" tank. The FAA drained a large fuel sample from the tank and it was dark brown in
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color with contaminates. The tank's filter was also contaminated and visible sludge was observed at the bottom of the tank. Contaminated fuel was also found
in the helicopter's fuel tank and fuel filter bowl, and the fuel nozzle was covered with a hardened dark brown substance. There were no other pre-mishap
mechanical deficiencies observed that would have precluded normal operation of the engine.
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Accident Rpt# WPR13FA411
09/16/2013 1535 PDT Regis# N204UH
Acft Mk/Mdl BELL UH 1B
Acft SN 62-2034
Eng Mk/Mdl LYCOMING T5313B
Opr Name: R&R CONNER
Detroit, OR
Apt: N/a
Fatal
1
Ser Inj
Opr dba:
0
Aircraft Fire: NONE
Events
1. Maneuvering-low-alt flying - Part(s) separation from AC
Narrative
HISTORY OF FLIGHT
On September 16, 2013, about 1535 Pacific daylight time, a Garlick UH-1B, N204UH, experienced a tailboom separation while logging in heavily wooded terrain
about 3 miles east of Detroit, Oregon. The pilot, who was the sole occupant on board, was fatally injured. The helicopter sustained substantial damage to the
tailboom, main rotor system, and fuselage. The helicopter was registered to Gitmo Holdings LLC, Stevensville, Montana, and operated by R&R Conner under
the provisions of 14 Code of Federal Regulations Part 133 as an external load logging flight. Visual meteorological conditions prevailed for the flight, and no
flight plan had been filed. The flight originated at about 1500.
Witnesses reported that when the helicopter was just above the trees, they either observed or heard the load of logs release early and impact the ground hard.
After looking up, they observed the helicopter's fuselage separate from the tailboom; both descending through the trees. The fuselage impacted the ground
inverted and the tailboom came to rest about 140 feet away.
A maintenance worker reported that shortly before the flight, the pilot had landed and shut down the helicopter for about a 45 minute lunch break. The pilot
looked over the helicopter and said he was very happy with it; he said it was running really well.
PERSONNEL INFORMATION
The pilot, age 53, held a commercial pilot certificate in helicopter, airplane single-, and multi-engine land, issued on April 27, 2010. The pilot also held an
instrument rating in both helicopter and airplane. The pilot held a second-class medical certificate issued on February 12, 2013, with the limitations that he is
not valid for any class after, and he must wear corrective lenses. According to the pilot's US Forest Service Helicopter Pilot Qualifications and Approval
Records dated July 17, 2013, he reported having 19,000 total helicopter hours, 14,000 of which were in the accident helicopter make and model.
AIRCRAFT INFORMATION
The Garlick helicopter, serial number 62-2034, was manufactured by Bell Helicopter as serial number 554 in 1962. It was powered by a T53-L13BA engine. The
maintenance logbook records were found within the helicopter. The records did not contain dates or aircraft total time, therefore, the most recent maintenance
was unable to be determined. The documents did reveal that several component inspections were not completed within the manufacturer's recommended time.
During the postaccident examination, the hobbs meter was located and read 6,061.3 hours.
According to the Federal Aviation Administration (FAA), the previous owner relinquished the aircraft's airworthiness certificate to avoid punitive action by the
FAA, who had been trying to revoke the airworthiness certificate due to the owner's poor maintenance of the aircraft. In 2010, a new airworthiness certificate
was issued for the helicopter to the accident pilot.
A different mechanic reported that the helicopter had recently sat unused for about one month between jobs. The helicopter was put back in service the day
before the accident occurred. The mechanic mentioned that the pilot had previously indicated the helicopter felt like it "shuffled" during translational lift;
however, the mechanic suspected the transmission mounts were starting to wear and would need to be changed at a later date.
METEOROLOGICAL INFORMATION
The nearest weather reporting station was about 38 miles to the northwest at McNary Field Airport in Salem, Oregon at an elevation of 214 feet. At 1556, the
weather was reported as wind from 130 degrees as 3 knots, visibility 10 statute miles, broken clouds at 4,900 and overcast clouds at 5,500 feet above ground
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level (agl), temperature 21 degrees C, dewpoint 13 degrees C, and an altimeter setting of 29.94 inches of mercury. In the remarks section it stated that rain
started at 1537 hours and ended at 1552 hours.
WRECKAGE AND IMPACT INFORMATION
On scene examination by a FAA Inspector revealed that the helicopter came to rest on the opposite side of a northwest/southeast orientated dirt road from the
log landing site. The terrain was hilly, heavily wooded, and remote. The trees around the accident site sustained limited damage; one tree was topped and
others sustained vertical scrapes down the trunks.
The wreckage debris path extended almost parallel to the dirt road; the helicopter came to rest in four major pieces the fuselage/transmission, engine, main
rotor blades, and tailboom. The fuselage and transmission were found upside down at the southeastern most point of the wreckage path. The engine was found
in the same general vicinity as the fuselage. The main rotor head and blades were separated from the main rotor shaft, and were located about 120 feet
northwest of the main fuselage. One of the main rotor blades was embedded into the ground and extended the second blade into the air at about a 45 degree
angle. The tailboom was separated from the fuselage and was located 140 feet northwest of the main rotor blades. The tail rotor gearbox, assembly, and tail
rotor blades were still attached to the vertical fin. One tail rotor blade remained mostly undamaged; the second tail rotor blade sustained a 45 degree bend
away from the vertical fin.
MEDICAL AND PATHOLOGICAL INFORMATION
An autopsy was performed on the pilot on September 17, 2013 by the Office of the State Medical Examiner, Clackamas, Oregon. The pilot's cause of death
was blunt force head trauma.
The FAA Civil Aerospace Medical Institute performed forensic toxicology on specimens from the pilot with negative results for carbon monoxide and ethanol.
Rosuvastatin, which is used to treat high cholesterol and related conditions, was detected in the blood and liver.
TESTS AND RESEARCH
A post accident examination of the airframe and engine occurred in Dallas, Oregon on January 29, 2014.
Airframe
The cabin sustained significant damage. The windscreen, chin bubble, instrument panel, and roof were all found separated from the structure. The aft fuselage
was mostly intact with the transmission still attached. The tailboom had separated from the aft fuselage at its attachment points; the skin along the sides of the
tailboom had a "wave" like appearance. The tailboom attachment points were removed from the airframe for further examination.
Control continuity was established throughout the airframe with the exception of a segment of the tail rotor drive shaft that extended from the transmission,
which was not located.
The main rotor shaft was fracture separated just below the main rotor head. The fracture surface was indicative of overload. At the fracture point, the main rotor
shaft was oblong with impact damage on two opposing sides. Damage was also noted on the main rotor blade hub, indicative of a mast bump event.
Engine
The engine was found separated from the helicopter. The exhaust and airframe inlet were removed. Organic debris was noted in the engine inlet, and metal
spray was found on the second stage power turbine nozzle vanes. Tear and batter damage was noted to the first stage axial compressor blades and the inlet
guide vanes. Rotation of the power turbine produced corresponding rotation to the engine output shaft and overspeed governor drive gearbox; the engine
rotated smoothly. The chip detector was examined and no debris was noted.
ADDITIONAL INFORMATION
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The tailboom attachment points were removed from the airframe and sent to the National Transportation Safety Board Laboratory for further examination.
The material research engineer reported that the fracture surfaces of the top right and left attachment point fittings exhibited relatively flat morphologies, with no
indications of local material deformation or out of plane fracture. Conversely, the two bottom fittings exhibited darker and rough tortuous fracture surfaces,
consistent with overstress failure.
Right Top Fitting
Both mating surfaces of the right top fitting were examined and crack arrest marks, indicative of progressive crack growth, was evident over almost the entire
fracture surface. The direction of these arrest marks indicated the cracks initiated near and emanated from a rivet hole within the fitting. The larger crack grew
through almost the entire fitting cross section; the smaller crack progressed toward the opposite direction.
The fitting aft fracture surface was further analyzed and the fracture surface exhibited striations, which are consistent with fatigue failure. The area around the
rivet hole possessed two fatigue crack initiate sites. The larger crack initiate site was on the outside surface of the fitting, and the smaller crack initiate site
occurred at a corner adjacent to the rivet holes.
Left Top Fitting
The aft fracture surface was relatively flat, orientated approximately perpendicular to the length of the fitting. After cleaning the fracture surface, crack arrest
marks were observed over most of the fracture surface. The fracture surface consisted of two progressive cracks that initiated on the concave surface and
grew in both directions, with fatigue striations throughout. The cracks grew through approximately 75% of the fitting cross-section, the remaining 15%
succumbed to overstress.
Bottom Fittings
The fracture surfaces of the bottom fittings exhibited features consistent with overstress failure. The fracture surfaces displayed a dull luster and tortuous
surface appearance. Neither of the bottom fittings exhibited indications of fatigue and/or other failure modes.
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Incident Rpt# CEN11IA379
06/06/2011 2132 CDT Regis# N866AS
Milwaukee, WI
Apt: General Mitchell Intl MKE
Acft Mk/Mdl BOMBARDIER INC CL-600-2B19
Acft SN 7517
Acft Dmg: MINOR
Eng Mk/Mdl GE CF34-3B1
Acft TT
Fatal
Opr Name: SKYWEST AIRLINES INC
Opr dba:
24969
0
Ser Inj
0
Aircraft Fire: NONE
Events
2. Landing - Landing gear not configured
Narrative
HISTORY OF FLIGHT
On June 6, 2011, about 2132 central daylight time, N866AS, a Bombardier CL-600-2B19, operated as Skywest Airlines flight 4443, landed with the right main
landing gear retracted on runway 19R at the General Mitchell International Airport (MKE), Milwaukee, Wisconsin. The 2 pilots, 1 flight attendant, and 41
passengers reported no injuries. The airplane sustained minor damage. The scheduled domestic passenger flight was conducted under 14 Code of Federal
Regulations Part 121. Visual meteorological conditions prevailed and an activated instrument flight rules flight plan was on file. The flight departed
Cincinnati/Northern Kentucky International Airport (CVG), near Covington, Kentucky, about 1951, and was destined for MKE. All airplane occupants evacuated
the airplane via the main cabin door.
During an interview with the operator's safety staff personnel, the flight crew reported that visual meteorological conditions and calm wind prevailed during the
approach. The airplane was on final approach when the gear did not come down, the flight crew informed the tower they needed to go around and flew east
over Lake Michigan where they proceeded to complete the quick reference handbook (QRH). They indicated that the gear operated normally on all previous
flights.
When the landing gear selector lever was first placed in the down position, the flight crew first noticed that the right main gear did not indicate down and locked.
They then received the triple chime warning, gear disagree, and proceeded to go around. Upon running the QRH for "Gear Down Disagree" they would receive
different indications from right main to both main landing gear unsafe, and they got the nose gear door open warning and oral messages. They also got a No. 3
hydraulic system high temperature caution message. The flight crew stated that about 28 seconds elapsed between the time that the landing gear selector lever
was positioned in the down position to the "Gear Disagree" message posting.
The flight crew stated that the engine indication and crew alerting system (EICAS) primary page landing gear indications with the landing gear selector lever in
the up position, initially, was that all lights were out. When the landing gear selector lever was selected down the first time, the flight crew received nose green,
left main green and right main unsafe indications.
During the go around, the landing gear selector lever was put in the up position and they received the gear up normal indication. The first officer remembered
that the right main would indicate a red unsafe indication very quickly, while the other landing gear would show transit, and then green safe indications. During
the third or fourth landing gear selector lever selections, both main landing gears showed unsafe indications, which also happened very quickly.
The flight crew stated that the QRH procedures were followed during the attempted manual landing gear extension. They turned the No. 3 hydraulic system off
as directed by the QRH and the system pressure subsequently indicated zero.
The flight crew pulled the landing gear alternate release T-handle as the QRH directed. The T-handle had slipped back a few inches from the fully extended
position, which they had achieved during the manual extension.
The EICAS page exhibited red hash marks for the right main landing gear the entire time on the approaches. During one approach, the indications showed both
main landing gears were unsafe and a low approach was performed to confirm which of gear were extended. The tower reported that the nose and left main
landing gear were down. While returning for the last approach and landing, the left main landing gear subsequently indicated green safe and the right main
landing gear still indicated unsafe.
During the alternate landing gear extension, the first officer initiated the QRH. He reported his seat height typically is in a very low seat position and he did not
reposition its height. He tilted the seat-back forward. He then moved the seat to the aft position. He pulled the T-handle from his seat and felt like he had
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complete ability to apply all his strength to move the T-handle. The first time he pulled the T-handle it was an abrupt pull and he felt the T-handle reach the
stop. The T-handle then sank back a few inches as if a spring was pulling it back down. The gear did not extend and he pulled the T-handle multiple times
holding it at the stop for 5 to 10 seconds each time. The only sound that was heard was oil bypassing under their feet. The first officer, who was right-handed,
was using his right hand but at one point used both hands to pull the T-handle. He told the captain that the T-handle would not stay up and the captain decided
that he would try manually extending the T-handle himself.
The captain's seat was low and all the way to the aft position. This is the position he always sits when flying. He reached over with both hands, pulled the
T-handle and held it there for ten seconds, and then re-pulled the handle a second time even harder. He could feel the T-handle hit a stop at full extension. He
then took back the controls from the first officer. Neither crewmember got out of their seats to pull the T-handle.
The flight crew estimated that the elapsed time between the first landing gear selector lever down command and the airplane's touchdown was about 35 to 45
minutes.
The airplane sustained minor damage to its right wing tip, right flap assemblies, and right flap pylon assemblies. The right main landing gear door was up, and
the right main gear was in its wheel well.
PERSONNEL INFORMATION
The captain held an airline transport certificate with a multiengine land airplane rating and commercial privileges for single engine land airplanes. He held a first
class medical certificate with no limitations. The operator reported that he had accumulated about 8,618 hours of total flight time, which included about 438
hours in the Bombardier CL-600-2B19. The captain had flown about 187 hours in the last 90 days, 69 hours in the last 30 days, and 6 hours in the last 24
hours.
The first officer held a commercial certificate with single-engine land, multiengine land, and instrument airplane ratings. He held a first class medical certificate
without limitations. The operator reported that he had accumulated about 5,156 hours of total flight time, including about 2,997 hours in the Bombardier
CL-600-2B19. The first officer had flown about 197 hours in the last 90 days, 44 hours in the last 30 days, and 6 hours in the last 24 hours.
AIRCRAFT INFORMATION
The airplane was a Bombardier Canadair model CL-600-2B19, twin engine, transport category Regional Jet (CRJ), with serial number 7517. It was
manufactured on June 15, 2001. The CRJ had a maximum takeoff weight of 53,000 pounds. The engines were General Electric model CF-34-3B1 engines that
delivered 8,900 pounds of thrust each. The airplane was on a continuous airworthiness maintenance program. The last service check was conducted on June
1, 2011. At the time of the incident, the airplane had 24,969.4 flight hours and 20,132 flight cycles. The airplane was configured with 53 seats, of which 50 were
passenger seats located in the main cabin.
The airplane was equipped with a retractable tricycle landing gear system that comprised two main landing gear (MLG) assemblies mounted on the inboard part
of each wing, and a nose gear assembly mounted directly below the flight compartment. Both MLG retract inward into recesses in the wing and center fuselage,
and the nose landing gear (NLG) retracts forward. The landing gear system, operated by a selector lever, is electrically controlled by a proximity sensor
electronic unit (PSEU) and hydraulically operated by the no. 3 hydraulic system. The MLG system comprises a selector valve, run-around and bypass valve, a
left and right MLG sidestay actuator and uplock mechanism. The NLG system comprises a selector valve, extension/retraction actuator, uplock assembly,
downlock, nose selector valve, and priority valve, bypass valves, restrictors, and check valves.
Two of the three NLG doors are operated hydraulically and are sequenced to operate independently of the NLG position, while the other (single door) is
mechanically linked to the nose gear position. During extension, the forward doors open before the NLG is released from the uplock. The nose gear assembly
will then extend, simultaneously opening the rear door. Upon reaching full extension (when the NLG is down and locked), the forward doors close and remain in
that configuration until a retraction command is selected.
The cockpit is equipped with a landing gear control panel, which contains a landing gear selector lever. When the selector lever is manipulated, an electrical
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command is sent to the PSEU to extend or retract the landing gear. Both MLG are extended in the outboard direction by their respective MLG sidestay
actuators and are hydro/mechanically locked in place for landing. Each gear is retracted by the MLG side stay actuator in the inboard direction and locked in
the MLG wheel wells during flight by their respective uplock mechanism.
During normal landing gear extension, when the landing gear selector lever is placed in the gear down position, the selector lever module sends an electrical
extension command to the proximity sensor system and provides electrical signals to command the MLG and the NLG selector valves to their gear down
position. When the MLG selector valve transitions to its gear down position, the valve is designed to allow no. 3 hydraulic system fluid, from the priority valve,
to be ported, via the run-around and bypass valve, simultaneously to the uplock assembly and the extend side of the sidestay actuator for each MLG. The
hydraulic pressure causes each uplock assembly to unlatch and release the MLG assembly. When unlatched, an uplock sensor (on the uplock mechanism)
provides an input to the PSEU, which in turn signals the data concentrator units (DCUs) to generate an amber 'IN TRANSIT' gear indication on the EICAS
display for each of the gear. When hydraulic pressure is supplied to the extend side of the gear actuators (sidestay), the actuator extends causing each MLG to
extend to its full down and locked position; the extension rate is controlled by a restrictor in the actuator up line.
The landing gear alternate release system provides the flight crew with another means to extend the landing gear in the event that an electrical or hydraulic
failure within the landing gear system prevents the landing gear from being extended normally. The alternate extension system is controlled by the vertical
movement of a T-shaped alternate release handle. To extend the landing gear manually, a flight crew member must pull up on the alternate release handle.
Movement of the T-handle is transmitted by a cable circuit to the NLG uplock release mechanism and to the MLG release mechanism.
For the NLG system, the mechanisms activates the NLG door bypass valve and the NLG bypass valve and releases the NLG uplock and nose door lock. For
the MLG system, rotation of the interconnect lever results in three actions: 1) the displacement is transmitted by two cables to the release levers on the left and
the right uplock mechanism to unlock the uplock mechanism permitting the gear to extend by gravity and 2) re-positions the runaround and bypass valve into
bypass mode and 3) positions the assist valve to pressurize the assist actuator.
In bypass mode, the runaround and bypass valve connects the extend pressure from the selector valve and both extend and retract pressure from the sidestay
actuators and the extend pressure of the uplock assembly to an independent return line. The extend pressure of the uplock assembly is sent to the return line
via the selector valve and check valves. The removal of all hydraulic pressure from the uplock mechanisms and sidestay actuators is designed to allow the
gear to free-fall regardless of the position of the MLG selector valve. The activation of the downlock assist selector valve results in the valve porting no. 2
hydraulic system pressure to the MLG downlock assist actuators to assure down locking of the main gears after free-falling.
The landing gear indication system provides the status of each landing gear position on the landing gear display area on the EICAS primary page. The primary
page contains three rectangles that will change color depending on the position of the landing gear. When a gear assembly is "up and locked", its respective
rectangle will be colored white and display "UP". When the indication system detects that a gear assembly is not "up and locked" or "down and locked", its
respective rectangle will transition to amber and when a gear assembly is "down and locked", its respective rectangle will be colored green and display "DN". If
any landing gear remains in transit for longer than 28 seconds, the amber intransit indication of that affected gear will change to red (gear unsafe).
Simultaneously a red 'GEAR DISAGREE' message will be displayed, accompanied by a 'GEAR DISAGREE' aural warning message. When any landing gear
assembly remains in its up and locked position for longer than 6 seconds after the landing gear has been commanded down, this will result in a landing gear
disagree aural warning being annunciated along with an EICAS red gear disagree warning message. This warning will also be annunciated when any landing
gear assembly remains in its downlock position for longer than six seconds when the landing gear has been commanded up.
If the landing gear selector lever remains "UP" during the manual extension, the EICAS immediately displays a "GEAR DISAGREE" message and the master
warning illuminates and the corresponding cancelable voice message sounds.
The landing gear indication and warning system comprises a PSEU and multiple proximity sensors and switches located within the control system. The PSEU
logic analyzes inputs from these various proximity sensors and switches to determine the status of the landing gear and doors. Its output is displayed on the
EICAS system (primary page) and master caution/warning panel on the glare shield.
The No. 3 hydraulic system is an independent hydraulic system that supplies the landing gear system, braking system, and certain flight control systems with
hydraulic pressure. This hydraulic system comprises two alternating current motor pumps, identified as ACMP 3A and ACMP 3B, to generate hydraulic power
(3000 psi), a pressure manifold, and a return manifold. Pressure generation comes primarily from ACMP 3A. However, because the system No. 3 accumulator
had been removed from the airplane in accordance with the requirements of FAA Airworthiness Directive 2010-22-012, ACMP 3B is also ON at all times during
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the flight.
According to the airplane's component maintenance manual, the published wear limits for the uplock assembly's latch was:
1. The maximum permitted wear limit at overhaul is 0.003 in. (0,07 mm)
on either or both upper and lower wear surfaces. Between overhauls,
the maximum permitted wear limit can be a further 0.003 in. (0,07 mm),
for a total of maximum 0.006 in. (0,15 mm) per surface.
2. If the wear on either upper or lower surface is greater than 0.006 in.
(0,15 mm), replace the latch.
The hydraulic pumps are replaced on condition of failure.
METEOROLOGICAL INFORMATION
At 2052, the MKE weather was: wind 230 degrees at 7 knots; visibility 9 statute mile; sky condition few clouds at 5,500 feet; temperature 31 degrees C; dew
point 20 degrees C; altimeter 29.77 inches of mercury.
AIRPORT INFORMATION
MKE was a field elevation of 723 feet and was five runways. Runway 13/31 was concrete-surfaced, 5,868 feet long and 150 feet wide. Runway 7R/25L was
asphalt-surfaced, 8,012 feet long and 150 feet wide. Runway 7L/25R was asphalt and concrete surfaced, 4,800 feet long and 100 feet wide. Runway 1R/19L
was concrete-surfaced, 4,183 feet long and 150 feet wide. Runway 1L/19R was asphalt and concrete-surfaced, 9,690 feet long and 200 feet wide.
FLIGHT RECORDERS
The airplane was equipped with a L-3 Communications model FA2100-1020 cockpit voice recorder (CVR) with serial number 000228060. This model is a
solid-state CVR that records 2 hours of digital cockpit audio. Specifically, it contains a two-channel recording of the last two hours of operation and separately
contains a four-channel recording of the last 30 minutes of operation. The two-hour portion of the recording is comprised of one channel of audio information
from the cockpit area microphone (CAM) and one channel that combines three audio sources: the captain's audio panel information, the first officer's audio
panel information, and the observer's audio panel information. The 30-minute portion of the recording contains four channels of audio data; one channel for
each flight crew and one channel for the CAM audio information. The CVR was received at the recorder laboratory where it did not exhibit any heat or structural
damage. The audio information was extracted from the recorder normally, without difficulty. Timing of the summary was established by correlating CVR events
to common events on the flight data recorder (FDR).
The airplane was equipped with a L-3 Communications/Fairchild model FA2100 FDR with serial number 000174026, which was designed to meet the
crash-survivability requirements of TSO-C124a. This model records airplane flight information in a digital format using solid-state flash memory as the recording
medium. The FA2100 can record a minimum of 25 hours of flight data. It is configured to record 128 12-bit words of digital information every second. Each
grouping of 128 words (each second) is called a subframe. Each subframe has a unique 12-bit synchronization (sync) word identifying it as subframe 1, 2, 3, or
4. The sync word is the first word in each subframe. The data stream is "in sync" when successive sync words appear at proper 128-word intervals. Each data
parameter has a specifically assigned word number within the subframe. The FDR was received at the recorder laboratory in good condition and its data was
extracted normally from the recorder.
The FDR recording contained approximately 119.4 hours of data. Timing of the FDR data is measured in subframe reference number (SRN), where each SRN
equals one elapsed second. The incident flight was the last flight of the recording and its duration was approximately 1 hour and 42 minutes. The FDR incident
data was converted from SRN to the incident local time.
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WRECKAGE AND IMPACT INFORMATION
First responders found the airplane resting on runway 19R. Its nose gear and left MLG were found extended to their down and locked position and the right
MLG was found in its full up position in the wheel well. The operator's responding maintenance personnel were asked to pull and collar the FDR and CVR
circuit breakers. No circuit breakers were found extended when the recorders' circuit breakers were pulled and collared. Responding maintenance personnel
also found that the landing gear selector lever was in the down position, the emergency landing gear extension T-handle remained fully extended, and the air
driven generator was deployed.
The airplane was lifted by inflating airbags positioned beneath the right wing and the tail. A jack was placed under the right wing to support the airplane. A visual
inspection of the landing gear revealed that the right MLG remained within its respective wheel well and its door was flush with the aircraft fairing. An inspection
of the cockpit revealed that the landing gear manual release T-handle remained extended approximately 7 inches. An operator's maintenance technician pulled
up on the T-handle resulting in the T-handle moving approximately three additional inches to its fully extended position of 10 inches. The additional T-handle
displacement resulted in the right MLG extending out of its wheel well. When the maintenance technician released the T-handle from its fully extended position,
the T-handle automatically began to slowly retract causing the right MLG to stop extending. The T-handle had to be re-pulled and manually held in its fully
extended position for the right MLG to extend to its down and locked position. A slight push on the right MLG was required to bring the gear down and into its
down and locked position.
TESTS AND RESEARCH
A National Transportation Safety Board (NTSB) vehicle recorder specialist chaired a CVR group and produced a CVR factual report. The CVR group reviewed
recorded cockpit communications starting from the airplane's time while parked on the ramp at CVG. The report summarized communications during the flight's
initial approach descent at MKE and continued through the final landing. The flight crew statements are consistent with the CVR factual report's findings. The
report findings confirm the flight crew used QRH references during their six landing gear extensions attempts.
A NTSB vehicle recorder specialist downloaded, decoded, and produced a factual report to include graphic plots in reference to data from the FDR. The FDR
incident data, in part, indicated that about 20:52, while descending through a pressure altitude of approximately 2,745 feet, the left MLG data transitioned from
"Not Down and Locked" to "Down and Locked." The NLG data transitioned from "Not Down and Locked" to "Down and Locked" while the airplane's data
indicated it was approximately at a pressure altitude of 2,618 feet. The landing gear disagree warning transitioned from "Not Active" to "Active" while the
airplane was approximately at a pressure altitude of 2,217 feet. About 20:54, the right MLG transitioned from "Not Down and Locked" to "Down and Locked"
and the airplane's pressure altitude increased to approximately 2,869 feet. A second later, the landing gear disagree warning transitioned back to "Not Active."
About 20:54, the right MLG transitioned back to "Not Down and Locked." By 20:55, both the left MLG and NLG transitioned back to "Not Down and Locked."
Over approximately the next 10 minutes until 21:04, the NLG data transitioned six times from "Not Down and Locked" to "Down and Locked" and it remained at
"Down and Locked" until touchdown. From 21:00:36 to 21:01:06, the right MLG transitioned to "Down and Locked" and back to "Not Down and Locked" while
the airplane was at a pressure altitude about 4,150 feet. About 43 seconds later, the No. 3 hydraulic pressure decreased from approximately 2,720 pounds per
square inch (psi) to 14 psi. The airplane remained at about 4,150 ft.
While the airplane was at a pressure altitude 4,150 feet and over the next, approximately, 7.5 minutes until about 21:13, the No. 3 hydraulic pressure increased
to about 2,720 psi, decreased to about 10 psi, increased to about 2,700 psi, decreased to about 14 psi and then increased to about 2,680 psi where it remained
until touchdown.
About six and one half minutes later about 21:19, the left MLG transitioned from "Not Down and Locked" to "Down and Locked" and it remained at "Down and
Locked" until touchdown. The airplane's pressure altitude had decreased to about 900 feet at that time.
About 13 minutes and 14 seconds later at 21:33:04, the left MLG weight on wheels data transitioned from "Air" to "Ground" and one second later at 21:33:05
CDT, the NLG weight on wheels data transitioned from "Air" to "Ground." The FDR recorder specialist's factual report is appended to the docket material
associated with this case.
The pressure and return hydraulic filters were examined with computed tomography scans and digital radiography. A NTSB aerospace engineer produced a
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computed tomography specialist's factual report based on the radiographic examination. The examination showed that there were three particles found in the
pressure filter and one particle found in the return filter. There were several cracks noted in the epoxy material in the end caps of the return filter, and there
were some high-density areas noted in both filters within the filter material itself. The specialist's report is appended to the docket material associated with this
case.
Eight hydraulic system filters were shipped to the NTSB material laboratory along with hydraulic fluid that was collected with the filters for examination. A NTSB
chemist examined the filters and fluid and produced a materials laboratory factual report. The report indicated that all of the filters were comprised of an inner
metallic perforated tube with two additional types of filtration media laid over the top: an outer layer of stainless steel wire screen; and a filter consisting of
several layers of woven fiber mesh located between the tube and the steel mesh. A measurement of the openings in the outer wire screen found the openings
to be 166 by 198 micrometers (æm). The inner fiber mesh consisted of irregularly shaped and sized openings with an average opening size between 25 æm to
50 æm. The filter mesh layers for all of the filters were examined under a 5X to 50X stereo zoom-microscope. The examination of the filters revealed no
significant particulates within the mesh.
Each filter was rinsed with acetone to remove any material trapped within the filter material. There was no evidence of metallic particles present in the filtrate
rinse. The filtrate from both the fluid and the filters samples was further examined and it exhibited spectra was consistent with characteristic traits of a straight
chain, aliphatic hydrocarbon. Materials containing these types of bonds are present in airplane hydraulic systems. The chemist's factual report is appended to
the docket material associated with this case.
After the removal and replacement of the pressure and return filters, the operator flushed and retained the hydraulic fluid from the No. 3 hydraulic system. The
retained fluid was shipped to the Air Force Petroleum Office (AFPET) laboratory at Wright-Patterson Air Force Base, Ohio, for fluid analysis. The operator then
acquired hydraulic fluid samples from selected landing gear components and sent them to the AFPET laboratory for fluid analysis. The AFPET reports show
that the fluids from the No. 3 system flush contained an amount of specified sized particles per 100 microliters that exceeded the recommended amount of
those sized particles, in specified ranges, per 100 microliter as indicated in service limit values in the aircraft's maintenance manual. A sample from the No. 2
system had visible particles present. This sample included white and red colored fibers. Particles included a black hydrocarbon plastic material and black
fluorocarbon grease droplets. A sample from the return manifold revealed the presence of visible particles. This sample included clear and black colored fibers.
These particles included a few red particles, clear sticky flat tape adhesive, black dirt, shiny magnesium aluminum alloy pieces, and shiny copper metal
particle. A sample from the pressure manifold filter exhibited it contained clear, brown, and black colored fibers. The majority of particles were black spongy
pieces consistent with hydrocarbon gasket material. Other particles included black flat film, shiny magnesium aluminum alloy pieces, and clear particles of
silicon oxide as sand. Particles observed in a sample from the extend side of the right hand sidestay actuator included clear, red, and brown colored fibers.
These particles included shiny magnesium aluminum alloy pieces, green flat paint chip, black spongy hydrocarbon gasket material, and clear particles of silicon
oxide as sand. Particles observed in a sample from the retract side of the right hand sidestay actuator included clear and red colored fibers. These particles
included vermiculite (a shipping material), shiny magnesium aluminum alloy pieces, black dirt, and clear particles of silicon oxide as sand. The AFPET
laboratory reports are appended to the docket material associated with this case.
A NTSB systems engineer chaired a systems group, which performed follow-on examinations of removed parts, and produced a factual on the group's findings.
The report, in part, showed that the airplane's ACMP-3A pump was originally installed on another aircraft in 2001. It was removed in March of 2008 when the
pump did not produce specified output pressure. The pump was repaired in reference to discovered insufficient clearance between the piston shoes and the
shoe bearing plate and the repaired pump was installed on the incident airplane on May 3, 2008. The incident airplane's total time on that date was 18,069.9
hours and the airplane accumulated 14,516 cycles. Functional testing was performed on the pump and its output pressure was found to be lower than
specified. The pump's compensator was adjusted to increase the pump's output pressure to within specified tolerances. Further functional testing found that the
pump was unable to meet the minimum requirement for full outlet flow delivery and exceeded the maximum requirement for case flow.
The airplane's ACMP-3B pump was installed in its position 3B at the time of delivery from the factory in 2001. Functional testing was performed on the pump
and its output pressure was found to be lower than specified. The pump's compensator was adjusted to increase the pump's and it subsequently passed all test
requirements. Disassembly of the compensator revealed its spool assembly was difficult to remove from within its sleeve. Inspection of the sleeve showed that
"coking" was present within the sleeve.
The right sidestay was examined and some items did not conform to manufacturer specifications. This actuator was rebuilt by a third party vendor. These
inconsistencies did not exhibit any functional issues with the actuator.
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The right and left MLG uplock assemblies were examined. Visual inspection of the right uplock assembly at the working area of the latch revealed wear marks
on its upper surface and wear marks and a dent on the lower surface. Visual inspection of the hook on left uplock assembly revealed contact marks on its
upper surface. According to Messier Dowty, these marks are minor and consistent with normal in-service wear patterns. Both up lock assemblies met
acceptance test specifications. The hydraulic fluid contained with the locks was analyzed. The fluid's observed color was yellow to amber and particles were
found in the fluid.
The right MLG uplock pin and assembly remained connected to the airplane at its respective attachment location. Inspection of the uplock pin revealed that it
rotated smoothly and exhibited signs of wear, consistent with in-service usage. The diameter of the pin was not measured during the on-scene activities. The
left MLG uplock pin was not documented during the on-scene activities.
The runaround and bypass valve was tested and it met testing specifications. Fluid from within the valve was observed to contain debris, which included a
non-metallic seal strand.
The functional testing examination of the landing gear manual release T-handle revealed that the locking mechanism contained within the T-handle assembly
failed to maintain the T-handle in its fully extended position when a specified retract load was applied to the assembly. When this load was applied, the
T-handle fully retracted. Additional testing found that when a lesser load was applied to the T-handle, the T-handle began to slowly retract from its fully
extended position. The T-handle stopped retracting and remained in a position outside of full retraction. Disassembly of the T-handle assembly did not find any
mechanical discrepancies with the assembly, but did reveal that its inner housing and outer slider appeared to have an oily material coating of unknown origin.
An examined sample revealed it was a lubricant. The design of the T-handle assembly does not call for the application of lubricants. The systems group's
factual report is appended to the docket material associated with this case.
ADDITIONAL INFORMATION
An NTSB preliminary report, ENG10IA055, in part stated:
On September 28, 2010, at 1710 central daylight time, a Bombardier CRJ-200, operated as Skywest Airlines flight 3074, landed with the left main landing gear
retracted on runway 7R at General Mitchell Airport (MKE), Milwaukee, Wisconsin. The flight departed Omaha, Nebraska, at an unknown time, and was destined
for MKE. There were 39 reported souls on board and no injuries reported. All airplane occupants evacuated the airplane via the main cabin door with no injuries
reported during the evacuation.
According to local officials, air traffic, and the FAA, at 1659, the crew reported a gear indicator problem and performed a pass for air traffic control tower
observation. The tower observed the left main landing gear in the retracted position. After some unknown troubleshooting, the crew landed the airplane with the
left main landing gear retracted.
Preliminary damage assessment by on-site FAA inspectors revealed minor damage to the left wing tip, flap assemblies, and flap pylon assemblies. The left
main landing gear door was up and locked, and the left main gear was in the wheel well.
The NTSB received notification that on March 21, 2012, about 1830 eastern daylight time, a Bombardier CL-600-2B19, N457SW, aircraft operated as Skywest
Airlines flight 4710, while on approach into the Cleveland - Hopkins International Airport (CLE), near Cleveland, Ohio, received a 'gear disagree' message. It
was verified that the right main gear indication was amber indicating that it was not down and locked which indicative of the normal 'green' indication. As per
QRH procedures, the gear handle was placed in the up position and then again in the down position. The same result occurred; the right main did not indicate
down and locked - 'green'. The Captain instructed the gear to be cycled a third time and after this, the right main gear indicated down and locked. Due to the
situation, the Captain felt best to declare an emergency and brief the passengers. The subsequent approach and landing at CLE was uneventful.
The right MLG uplock assembly and respective uplock roller were shipped to the NTSB materials laboratory for examination. The examination revealed that
wear patterns were observed along the rolling surface of the uplock latch where the uplock roller contacted the slot surfaces and on the mounting ends that
contacted the bushings. The rolling surface exhibited two regions of metal wear. The upper band had a dark and roughened appearance, consistent with
adhesive wear and was approximately 0.11 inch wide. The lower band exhibited a repeating pattern of flat deformed facets around the circumference of the
roller with steps at the edge of the wear region and was approximately 0.19 inch wide. Circumferential wear lines and zones of roughened deformed metal were
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also observed, consistent with adhesive wear and mechanical deformation from a sliding contact. The mounting ends had circumferential wear marks that were
also consistent with sliding contacts with the bushings.
Subsequent to the incident, Bombardier revised the contamination limits in reference to hydraulic fluid in the aircraft maintenance manual.
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Incident Rpt# ERA15IA198
04/16/2015 1910 EDT Regis# N610RL
Acft Mk/Mdl DASSAULT-BREGUET MYSTERE FALCON Acft SN 68
Acft TT
Eng Mk/Mdl GARRETT TFE-731-3-BR
Orlando, FL
Apt: Orlando International MCO
Acft Dmg: MINOR
9949
Fatal
0
Ser Inj
Opr Name: RLB HOLDINGS TRANSPORTATION LLC Opr dba:
Rpt Status: Prelim
0
Prob Caus: Pending
Aircraft Fire: IFLT
AW Cert: STT
Events
1. Enroute-descent - Electrical system malf/failure
Narrative
On April 16, 2015, about 1910 eastern daylight time, a Dassault-Breguet, Mystere Falcon 900B, N610RL, received minor damage from an inflight electrical fire.
The airplane had departed Westchester County Airport (HPN), White Plains, New York about 1715, destined for the Orlando International Airport (MCO),
Orlando, Florida. The two flight crew members, flight attendant, and 12 passengers were not injured. The corporate flight was being operated under the
provisions of Title 14 Code of Federal Regulations (CFR) Part 91.
According to the cabin attendant, she noticed "a weird smell" in the galley area and started checking the appliances. She could find nothing wrong with them
and asked the pilots if they had any idea where the smell was coming from. She turned around and opened the area in the galley that they kept the glasses in,
and noticed "a glow behind it." The pilot then came back to assist her. He grabbed a fire extinguisher, used it all and then asked her for another one. She
handed him the second one and he was able to put out the fire. The pilot then told her to watch the area.
According to the pilot, they noticed the strange odor when they were on the Standard Terminal Arrival Route (STAR) for MCO descending through 20,000 feet
above mean sea level. They checked that all the ovens, microwaves, and warmers were in the "off" position. All were found to be "off" but the smell continued
and then smoke appeared. The crew declared an emergency and asked for priority to the nearest airport. The pilot then "gave the flight controls" to the copilot
who took over flying and the radio duties (the autopilot was on and engaged). The copilot then turned off the power to the cabin and to the galley. The pilot then
used a fire extinguisher to extinguish fire. The fire was suppressed by getting behind the crystal storage and spraying the area with a Halon portable fire bottle.
The crew decided to continue to Orlando as it had large runways with excellent fire and rescue capabilities. The landing was normal. The crew evacuated the
passengers and fire personnel double checked the airplane for any signs of continued fire or hot spots. None were found.
According to the copilot, they noticed a smell at first and then smoke from somewhere in the airplane when they were on the STAR for MCO at approximately
1910. They determined that it was coming from the galley area, declared an emergency, and asked for vectors to the nearest airport. After the sight of smoke,
and the cabin attendant assuring that everything was powered off, the pilot "gave the controls" to him. The pilot then went back into the cabin to assist the flight
attendant by spraying the galley area with the fire extinguisher. The Halon did its job and no more smoking was noticed so they decided to continue to MCO for
an "abnormal approach and landing" with priority from air traffic control. After landing they could see the signs of burnt wiring behind the galley.
Examination of the area in back of the crystal (glass) storage area of the galley revealed the presence behind the plenum ducting of sooting and the remains of
a burnt wiring bundle.
Examination of the wiring bundle revealed that the bundle contained the 28 volt direct current wiring and ground wires for the cabin overhead lighting, and that
the associated circuit breaker in the cockpit had been tripped. Further examination of the area also revealed that the wires had been routed over the insulation
bags and not directly next to the airplane's structure, and that they had been in contact with, or in close proximity to, a soft oxygen line which had been
completely burned through.
Examination of the oxygen system revealed that the burned oxygen line was part of the passenger oxygen system of the airplane, and that the line was not
pressurized except when the passenger oxygen system was activated during situations that would require its use; such as cabin depressurization, or the
presence of smoke, or unusual odors.
According to Federal Aviation Administration records, the airplane was manufactured in 1988. The airplane's most recent continuous airworthiness inspection
was completed on April 13, 2015. At the time of the accident; the airplane had accrued approximately 9,949 total hours of operation.
Portions of the wiring bundle and the burnt oxygen line were retained by the NTSB for further examination.
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04/16/2015 2245
Regis# N2691W
Rifle, CO
Acft Mk/Mdl FAIRCHILD SA227 AC
Acft SN AC-655B
Rpt Status: Prelim
Eng Mk/Mdl AIRESEARCH TPE331 SERIES
Acft TT
Fatal
Opr Name: KEY LIME AIR CORPORATION
Opr dba: DENVER AIR CONNECTION
26855
0
Apt: Garfield County Rgnl RIL
Ser Inj
0
Prob Caus: Pending
Aircraft Fire: NONE
AW Cert: STC
Events
1. Takeoff - Loss of engine power (total)
Narrative
On April 16, 2015, about 2245 mountain daylight time, a Fairchild SA227-AC, multi-engine airplane, N2691W, operating as Key Lime flight 168, was
substantially damaged after an uncontained engine failure during climb at Rifle, Colorado. The pilot was not injured. The airplane was registered to CBG LLC;
Wilsonville, Oregon; and was operated by Key Lime Air Corporation; Englewood, Colorado. Dark night instrument meteorological conditions (IMC) prevailed at
the time of the accident and an instrument flight rules (IFR) flight plan had been filed for the 14 Code of Federal Regulations Part 135 scheduled cargo flight.
The airplane departed from Garfield County Regional Airport (RIL), Rifle, Colorado, at 2237 and was destined for Denver International Airport (DEN), Denver,
Colorado
The pilot reported that during climb, when still well below the tops of nearby mountains, he heard a "bang" followed by a complete loss of power and engine fire
indications from the right engine. After completing appropriate checklist items the pilot declared an emergency and diverted to Grand Junction Regional Airport
(GJT), Grand Junction, Colorado, for an instrument approach and an otherwise uneventful landing at 2311. A postaccident examination of the airplane revealed
the second stage turbine rotor from the right engine had separated. One portion of the rotor exited through the left side of the engine and nacelle structure,
penetrated the right side of the fuselage, and came to rest inside the fuselage wall. Other portions of the separated rotor exited through the right side of the
engine. There was thermal damage to the engine and the inside of the nacelle structure, but no evidence of a sustained fire in that area.
At 2253 the Automated Surface Observation System at RIL reported wind from 130 degrees at 4 knots, visibility 10 miles in light rain, ceiling overcast at 2,000
feet above ground level (agl), temperature 3 degrees C, dew point 2 degrees C, altimeter 30.04 inches of mercury. Data from the United States Naval
Observatory indicated that moonset occurred at 1740 and sunset occurred at 1950.
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Accident Rpt# ERA15FA178
04/06/2015 1300 CDT Regis# N555JC
Cherokee, AL
Apt: N/a
Acft Mk/Mdl HUGHES 369D
Acft SN 180256D
Rpt Status: Prelim
Eng Mk/Mdl ROLLS ROYCE 250-C20B
Acft TT
Fatal
Opr Name: HAVERFIELD INTERNATIONAL INC
Opr dba:
9813
1
Ser Inj
0
Prob Caus: Pending
Aircraft Fire: NONE
AW Cert: STN
Events
1. Maneuvering-low-alt flying - Loss of visual reference
Narrative
On April 6, 2015, about 1300 central daylight time, a Hughes 369D, N555JC, was substantially damaged when it impacted the Tennessee River adjacent the
Natchez Trace Bridge, near Cherokee, Alabama. The commercial pilot was fatally injured. Low ceilings and fog prevailed. A company flight plan was filed for
the flight, which originated at Roscoe Turner Airport (CRX), Corinth, Mississippi, destined for Scottsboro Municipal Airport-Word Field (4A6), Scottsboro,
Alabama. The positioning flight was conducted under the provisions of 14 Code of Federal Regulations Part 91.
According to a witness, a former private pilot, he heard the helicopter land in a National Park Service field contiguous to his property, about 3,900 feet from the
1-mile-long, north-south Natchez Trace Parkway Bridge. He couldn't see the bridge at the time due to fog and light mist.
The helicopter remained on the ground for about 45 seconds, still powered with rotors turning; then power increased and it took off smoothly, clearing trees by
about 30 feet. The helicopter subsequently headed toward the bridge, and after about 10 to 15 seconds, the witness lost sight of it in the fog. As the helicopter
flew, the witness heard no anomalies, and the engine sounded "healthy." He subsequently heard the helicopter hit the water with no change in sound until
impact.
According to another witness, he was fishing under the south end of the bridge when the accident occurred. The weather was foggy with low visibility and rain.
The witness heard the helicopter for about 10 to 15 minutes before seeing it coming toward him, paralleling the west side of the bridge. When he first saw the
helicopter through the fog, it was level with the top of the bridge. It began a gradual descent, then about 10 seconds before water impact, dropped (nose-dived)
to about 25 feet above the water. It subsequently descended at a 10- to 15-degree angle, and impacted the water near the center of the river, about 50 to 100
feet east of a green buoy (about 100 yards west of the bridge.)
There was no change in sound before the helicopter hit the water, with the same "whining" noise until impact. At impact, the witness saw the helicopter's tail
"kick over" the top of the main rotor blades and snap off. The helicopter did not hit the bridge.
The helicopter was recovered from the river on April 9, 2015. It was missing the aft part of the tail boom, including the tail rotor and gear box, from about 33
inches (fuselage station 230) aft of the tail boom mount, and only remnants of one main rotor blade were subsequently recovered; the other blades remained
missing. The left skid was also missing.
Damage began at the helicopter's front, lower left side, and extended upwards. There was no hydraulic crushing (water impact damage) to the bottom of the
fuselage.
Control continuity was confirmed from the cockpit to the rotor head, both vertically through the collective, and laterally and longitudinally through the cyclic. Yaw
control through the rudder pedals was confirmed from the cockpit to the remnants of the "long tail rotor control rod" in the severed tail boom.
Rotor system drive continuity was confirmed from the engine to the transmission, the transmission to the rotor hub, and from the transmission aft to where the
tail rotor drive shaft was severed along with the tail boom.
Three of the five rotor blades were separated just outboard of the doubler at the main rotor root fitting, and two blades were separated through the strap
assemblies and blade pitch housings, consistent with full power on the rotor system at water impact. Extensive damage was also found on the hub upper shoe
in the vicinity of all five pitch change housings, consistent with a medium-to-high collective setting at the time of impact.
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Accident Rpt# ERA14LA098
01/14/2014 1447 EST Regis# N720QS
Savannah, GA
Apt: Savannah/hilton Head Intl KSAV
Acft Mk/Mdl ISRAEL AIRCRAFT INDUSTRIES
Acft SN 085
Eng Mk/Mdl P&W CANADA PW306A
Acft TT
Fatal
Opr Name: NETJETS AVIATION, INC
Opr dba:
7493
0
Ser Inj
0
Aircraft Fire: NONE
AW Cert: STT
Summary
The airplane had undergone maintenance for an unresolved landing gear position indication issue. The flight crew reported that, during the third
postmaintenance operational check flight (OCF), the airplane experienced a "thud" and momentary shudder as it climbed through 16,000 ft at 280 knots. The
flight crew observed no other anomalies and did not experience any abnormal handling characteristics with the airplane during the remainder of the flight.
Upon landing, the flight crew observed that the auxiliary power unit (APU) service door was open and bent but that it remained attached at its aft hinge point.
The door's lower latches were partially engaged, and a detailed examination revealed no anomalies with the latch assembly. The lower portion of the rudder
was substantially damaged due to contact with the service door. Review of maintenance work orders revealed that maintenance had been performed in the
APU compartment before another OCF that had been conducted several days before the accident flight and that it was subsequently inspected in accordance
with the maintenance provider's procedures. Maintenance personnel should have noted that the APU service door was not properly secured during the
inspection. The flight crew's preflight inspection did not include inspecting the APU service door because it was only accessible with the aid of a ladder.
.
Cause Narrative
THE NATIONAL TRANSPORTATION SAFETY BOARD DETERMINED THAT THE CAUSE OF THIS OCCURRENCE WAS: Maintenance personnel's failure
to properly secure the auxiliary power unit service door before returning the airplane to service, which resulted in substantial damage to the rudder.
Events
1. Enroute-climb to cruise - Part(s) separation from AC
Findings - Cause/Factor
1. Personnel issues-Task performance-Inspection-Post maintenance inspection-Maintenance personnel - C
2. Aircraft-Aircraft handling/service-Maintenance/inspections-Return to service-Incorrect service/maintenance - C
3. Aircraft-Aircraft handling/service-Maintenance/inspections-(general)-Inadequate inspection - C
Narrative
On January 14, 2014, approximately 1447 eastern standard time, an Israel Aircraft Industries Gulfstream 200, N720QS, operated by NetJets Aviation, Inc., was
substantially damaged when the auxiliary power unit (APU) service door opened in-flight near Savannah, Georgia. Both airline transport pilots were not injured.
Visual meteorological conditions prevailed, and an instrument flight rules flight plan was filed for the local flight, which departed Savannah/Hilton Head
International Airport (SAV) at 1436. The post-maintenance operational check flight (OCF) was conducted under the provisions of Title 14 Code of Federal
Regulations Part 91.
The airplane had been undergoing general maintenance, as well as maintenance for a recurring landing gear position indication issue. The airplane was
returned to service on January 9, 2014, and during the subsequent OCF, the crew discovered that the landing gear position indication issue had not been
resolved. Additional maintenance was performed, and during the return-to-service inspection on January 10, a large quantity of fuel was noted in the APU
compartment due to a leak in a fuel line fitting to the APU fuel control unit. An o-ring was removed and replaced at the location of the leak, and a subsequent
leak test revealed no anomalies. The airplane was returned to service, and a second OCF revealed that the landing gear indication issue remained unresolved.
The airplane underwent further maintenance from January 10 to January 14, and was returned to service on January 14.
Prior to departure on the accident flight, the crew performed a preflight inspection and observed no anomalies. They described the takeoff as normal, and
stated that as the airplane climbed through 16,000 feet at 280 knots, they experienced a "thud" and momentary shudder. No other anomalies were observed,
and the crew did not experience any abnormal handling characteristics with the airplane. Upon landing at SAV, the crew observed that the APU service door
was open and bent, but remained attached at its aft hinge point. The door's lower latches were in the down and locked position, and the side latches were in the
open position. The crew noted that the side latches were "bent as if torn away." The lower portion of the rudder exhibited substantial damage due to contact
with the door. According to the operator, the crew's preflight inspection included only a visual confirmation of the security of the service door, because due to its
location on the airplane, the door could only be accessed with the aid of a ladder.
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Review of work orders indicated that prior to each OCF on January 9 and 10, the airplane received a full post-maintenance inspection, which included verifying
aft fuselage accesses, antennas, and vents for leakage, condition, and security. Prior to the accident flight, the airplane underwent an inspection only to the
areas that had received maintenance since the previous OCF. Between the OCF on January 10 and the accident flight on January 14, the APU door was not
accessed and no maintenance was performed in the area of the APU.
According to the airframe manufacturer, between 2005 and 2014, there were 5 other reports of in-flight APU service door events involving G200 airplanes,
resulting in minor damage. In October 2008, the manufacturer issued a maintenance and operations letter (G200-MOL-08-0009) to remind flight crews to
inspect all service and access doors for security prior to flight. On February 21, 2014, the manufacturer issued Maintenance and Operations Letter
G200-MOL-14-0002, advising maintenance and operations personnel to verify that the APU service door was properly seated and latched when being secured.
Review of the previous in-flight APU service door events by the manufacturer revealed that, during manufacturing, the forward APU service door latches may
have been trimmed to prevent interference with the frame, possibly resulting in insufficient overlap between the latches and the doubler. On May 30, 2014, the
airframe manufacturer issued Service Bulletins 200-52-403 and 280-52-136, calling for inspection of the forward APU service door latches on G200 and G280
airplanes, respectively. Examination of the latches on the accident airplane immediately following the accident revealed the proper overlap between the latches
and doubler.
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Accident Rpt# CEN15CA066
11/26/2014 2321 CST Regis# N902LC
Acft Mk/Mdl MD HELICOPTER INC MD 900
Acft SN 900-00118
Apt: N/a
Acft Dmg:
Fatal
Eng Mk/Mdl P&W CANADA PW207E
Opr Name:
Highlandville, MO
0
Ser Inj
Opr dba:
0
Aircraft Fire:
AW Cert: STN
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11/26/2014 2321 CST Regis# N902LC
Highlandville, MO
Apt: N/a
Acft Mk/Mdl MD HELICOPTER INC MD 900
Acft SN 900-00118
Eng Mk/Mdl P&W CANADA PW207E
Acft TT
Fatal
Opr Name: AIR METHODS CORPORATION
Opr dba:
3110
0
Ser Inj
0
Aircraft Fire: NONE
AW Cert: STN
Events
1. Approach - Miscellaneous/other
Narrative
On November 26, 2014, about 2321 central standard time, a McDonnell Douglas model 900 helicopter, N902LC, was substantially damage when the main rotor
blades struck a power line during landing near Highlandville, Missouri. The pilot was executing an off-airport landing to a field in order to pick-up a patient at the
time of the accident. The pilot sustained minor injuries; the two medical crew members were not injured. The helicopter was registered to the Lester E Cox
Medical Centers and operated by Air Methods Corporation under the provisions of 14 Code of Federal Regulations Part 135 as an air medical flight. Night visual
meteorological conditions prevailed for the flight, which was operated on a company visual flight rules flight plan. The flight originated from the Cox Medical
Center facility in Springfield, Missouri, about 2310.
The pilot reported that he had completed two orbits of the accident scene and selected a large nearby field as the landing site. Using the helicopter search and
landing lights, the pilot identified a pole at the southeast corner of the field with wires running toward the north. However, about 20 feet above ground level, he
identified a second set of power lines paralleling the helicopter's flight path. He observed those power lines in his peripheral vision from under the night vision
goggles. He attempted to move the helicopter to the left to avoid them, but the main rotor blades inadvertently struck one of the lines. The pilot maintained
control and landed safely.
A postaccident examination revealed substantial damage to three of the main rotor blades. The remaining two rotor blades sustained minor damage. Debris
from the power line struck and broke the pilot's door window. The pilot did not report any failures or malfunctions with the helicopter before impacting the power
lines.
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Accident Rpt# ERA15FA181
04/12/2015 1625 EDT Regis# N119RL
Fort Lauderdale, FL
Acft Mk/Mdl PIPER PA-31T1
Acft SN 31T7904002
Acft Dmg: DESTROYED
Eng Mk/Mdl P&W PT6A-11
Acft TT
Fatal
Opr Name: JOHN P VAN OMMEREN
Opr dba:
3267
4
Ser Inj
Apt: Fort Lauderdale Executive FXE
Rpt Status: Prelim
0
Prob Caus: Pending
Aircraft Fire: GRD
AW Cert: STN
Events
1. Approach-VFR pattern final - Fire/smoke (non-impact)
Narrative
On April 12, 2015, about 1625 eastern daylight time (EDT), a Piper PA-31T1, N119RL, collided with terrain on final approach to runway 13 at Fort Lauderdale
Executive Airport (FXE), Fort Lauderdale, Florida. The private pilot and three passengers were fatally injured and the airplane was destroyed by impact forces
and a post-crash fire. The airplane was registered to a private company and was operated by the pilot under the provisions of 14 Code of Federal Regulations
Part 91 as a personal flight. Day, visual meteorological conditions prevailed for the flight, and an instrument flight rules flight plan was filed. The local flight
originated from Orlando Executive Airport (ORL), Orlando, Florida, about 1520.
According to information obtained from air traffic control, the pilot checked in with FXE tower personnel and initially did not report any problems with the flight.
N119RL was cleared to land on Runway 9. A short time later, the pilot reported to the local controller that he had smoke in the cockpit. The controller cleared
N119RL to land on any runway. The pilot responded that he would take runway 13, and the controller cleared him to land on runway 13. The pilot then called
out "mayday" several times before the airplane crashed approximately ¬ mile from the approach end of Runway 13.
The pilot, age 51, held a private pilot certificate with airplane single engine, multi-engine, and instrument airplane ratings. He reported 1,221 hours total flight
time on his most recent Federal Aviation Administration (FAA) medical certificate, dated February 18, 2015. Records recovered from the wreckage also indicate
that he completed a PA-31 initial training course one week prior to the accident.
The 7-seat airplane was manufactured in 1979 and was equipped with two Pratt and Whitney Canada PT6A-11 turboprop engines. The aircraft maintenance
records indicated that an annual inspection of the airframe, engines, and propellers was completed on April 3, 2015.
The wreckage was located inside a fenced nature preserve within the boundary of the airport. The main wreckage was located directly under the extended
centerline for runway 13. The wreckage debris field was about 167 ft in length and about 50 ft wide, oriented on a heading of about 112 degrees. All major
structural components of the aircraft were found within the confines of the debris field. Both engines were separated from the airframe and both propeller
assemblies were separated from the engines during the accident sequence.
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Air Data Research
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Air Carrier, Turbine Powered & Large Aircraft Reports

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