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On March 9, 2005, at 1312 central standard time, a Canadair CL-600, N660RM, registered to Romeo Mike Aviation Company Inc., operating as a 14 CFR Part 91 business flight, nose gear collapsed during an aborted takeoff from runway 36 and went off the departure end of the runway. Visual meteorological conditions prevailed and an instrument flight rules flight plan was filed. The airplane received substantial damage. The airline transport rated pilot-in-command, (PIC) airline transport rated co-pilot, and five passengers reported no injuries. The flight was departing Tupelo Regional Airport, Tupelo, Mississippi, en route to Teterboro, New Jersey on March 9, 2005.
The PIC stated they were cleared for takeoff from runway 36. The flaps were set at 20-degrees and the trim was set for takeoff. The PIC advanced the thrust levers to 93 percent N1 and started the takeoff roll. The takeoff run and acceleration were normal. The airplane reached V1 (128 knots) and VR (134 knots) and the PIC attempted to rotate the airplane with the control column. The control column would not move aft from the neutral position. The forward movement of the control column was normal. The aft movement beyond the neutral position felt as if it was locked against a stop. The airplane was about 4,000 feet down the runway between 140 to 145 knots. No enunciator’s lights were illuminated. The PIC commanded the abort, extended the spoilers, applied maximum braking, and maximum reverse thrust, and maintained centerline down the runway. After the abort was initiated the PIC stated he continued to apply rearward pressure on the control column and he was not sure if he felt or heard a "crunch." He further stated, something may have given and the control column may have moved aft of the neutral position. The PIC stated the "crunch" was felt or heard after or at the abort procedure. The noise or the crunch may have come from below the flight deck. The airplane went off the end of the runway and the nose wheel collapsed in the mud.
Review of information on file with the FAA Airman's Certification Division, Oklahoma City, Oklahoma, revealed the PIC was issued an airline transport pilot rating on September 12, 2004, with ratings for airplane single engine land, airplane single engine sea, airplane multiengine land and instrument airplane. The PIC is type rated in the CE-500, CE-525S, CL-600, DA-10, DA-20, HS-125 and Lear Jet. In addition the pilot was issued a flight instructor certificate on September 5, 2004, with ratings for airplane single engine land, and airplane multiengine land. The PIC was type rated in the CL-600 on September 9, 2004. The pilot stated he has 6,157 hours of which 98.7 hours are in the CL-600. The PIC was issued a proficiency certificate on September 12, 2004 for the CL-600 by FLYITS.COM. The PIC holds a mechanic certificate with ratings for airframe and power plant issued on May 13, 1999. The PIC holds a first class medical certificate issued on September 9, 2004 with no restrictions.
Review of information on file with the FAA Airman's Certification Division, Oklahoma City, Oklahoma, revealed the airline transport rated co-pilot, was issued an airline transport pilot certificate on June 17, 2002, with ratings for airplane single engine land, airplane multiengine land, and instrument rating. The co-pilot was issued a flight instructor certificate with ratings for airplane single engine land, airplane multiengine land and instrument airplane May 30, 2003. The co-pilot holds a ground instructor certificate with ratings for advanced and instrument issued on June 5, 2000. The co-pilot is not type rated in the CL-600. The co-pilot stated he has 3,600 hours of which 10.2 hours are in the CL-600. The co-pilot holds a first class medical issued on February 1, 2005, with the limitation, "must wear corrective lenses." The co-pilot's last biennial flight review was conducted 9 months before the accident.
Review of aircraft maintenance records revealed the last 12 month and 24 month inspection was conducted on March 5, 2005. The maintenance log revealed no maintenance had been performed on the flight control system since the 12 or 24-month inspection. During the 12 and 24-month inspection the following flight control related time limit maintenance checks (TLMC) were completed: TLMC 55-10-00-202,TLMC 55-10-11-205, TLMC 55-30-00-208, TLMC 55-50-00-210 and TLMC 55-50-00-212. The airframe has 8054 hours and the airplane has flown 6.1 hours since the March 5, 2005, inspection. The airplane was topped off with 1,588 pounds of Jet "A" fuel on March 9, 2005, with prist before the airplane departed on the accident flight. There is no record that Honeywell Service Bulletin 4015373-22-A0008 dated January 20, 2005, or Challenger Service Bulletin 600-0728 dated February 16, 2005 had been completed.
Review of weight balance information for the flight revealed the maximum ramp weight for the CL-600 is 41,400 pounds The ramp weight for the flight was 40,610.64 pounds. The arm was 504.0 and the moment was 20467.13. The percent of MAC is 17.2. The maximum permissible takeoff weight is 41,250 pounds. The computed takeoff weight was 40,310.64 pounds. The arm was 504.0 and the moment was 20317.13. The percent of MAC is 17.3. The CG range at takeoff weight is 16 percent of MAC to 26.5 percent of MAC.
The nearest weather reporting facility at the time of the incident was Tupelo, Mississippi. The 1317 surface weather observation was: few clouds 800 feet, 1,400 overcast, visibility 5 miles with light rain, temperature 43 degrees Fahrenheit, dew point temperature 37 degrees Fahrenheit, wind variable at 4 knots, and altimeter 29.95.
WRECKAGE AND IMPACT INFORMATION
The airplane departed the departure end of runway 36, 19 feet 2 inches left of the runway centerline. The left and right main landing gear tracks were present in the muddy grassy area for 109 feet 3 inches past the departure end of the runway. The nose wheel tracks were present for 46 feet 9 inches down the ground track. The nose wheel marks began again 60 feet 6 inches down the ground track and ended 63 feet 3 inches down the ground track. The nose wheel marks began again 68 feet 3 inches down the ground track and ended 74 feet 3 inches down the ground track. The airplane came to rest in a nose down attitude 135 feet down the ground track on a heading of 350-degrees magnetic.
Examination of the airplane revealed the nose landing gear assembly and wheels were pushed aft and rotated on the nose landing gear main fitting attachment point and penetrated the bottom side of the fuselage aft of the nose gear bay. The nose landing gear wheels and assembly were embedded inside the under cockpit avionics bay. The nose landing gear drag brace separated from the drag brace fuselage attachment points. The left and right nose gear doors were damaged. The nose landing gear doors were released by the investigation team by activation of the internal manual nose landing gear door release handle. The nose landing gear retraction actuator separated from the aft nose landing gear strut attachment point. The nose landing gear doors push/pull rods were intact. Mud and debris was present inside the gear well. The landing gear retract hydraulic lines were separated from the landing gear retract actuator. The right drag brace attachment point on the fuselage was fractured. The lower two stiffeners on the left side of the nose gear wheel well were bent downward. The No. 3 hydraulic nose gear up pressure line supporting bracket was separated from the nose landing gear wheel well top structure. The nose gear strut and wheel assembly were intact. There was no evidence of hydraulic fluid loss in the nose gear-landing strut.
The lower fuselage skin just aft of the nose gear bay was separated and pushed inward into the avionics bay by the nose landing gear tires. The nose landing gear assembly and tires penetrated the lower fuselage 25 inches aft of the nose landing gear bay aft bulkhead. The nose landing gear retraction and extension actuator remained attached to the nose landing gear attachment point. The lower left and right front quarter of the fuselage was indented ¼ inch and 6 inches across at fuselage 15 inches aft of the nose landing gear bay aft bulkhead. The bulkhead locate 45 inches aft of the nose landing gear bay aft bulkhead was bent aft.
The left and right main landing gear showed no evidence of damaged. The brake actuation pistons were leaking hydraulic fluid. The right main landing gear door was bent diagonally near the main landing gear door hinge attachment point. A witness who responded to the accident stated that immediately after the aborted takeoff that all four main landing gear tires were inflated and the wheels were hot. The landing gear was sprayed with water.
The passenger door upper aft doorstop bracket was broken flush with its mounting pad. The forward upper doorstop was intact. The remainder of the airplane showed no evidence of damage.
To gain access under the flight deck, the right hand fuselage access panel was removed. In the flight deck, the left and right flight management systems, reverse thrust control panel, autopilot status and select panel, and the autopilot controller were removed. The right hand control column stop assembly exhibited no anomalies. A dynamic test of the control column mechanical up stop revealed no anomalies.
The pilot and co-pilot control column contained a microphone plug near the control column base fairing. The microphone jack receptacle was oriented in a vertical position on the pilot's side. The microphone jack receptacle was oriented about 90-degrees to the control column vertical axis on the co-pilot side. The control column base fairing on the co-pilots side was missing paint where the microphone jack receptacle overlaps the control column base fairing during control column rotation. No paint was missing on the pilot's side of the control column base fairing. The missing paint on the co-pilots side extends 1/2 inch and is 1/2 to 1/4 inch wide. The co-pilot microphone jack receptacle plug was extended 3/8 inch from the fully plugged position. The spacing between the microphone jack receptacle and the top of the control column base fairing was 1/4 inch.
Review of the airplane log books revealed Arkansas Modification center, Inc., Little Rock, Arkansas, installed Supplemental Type Certificate Number SA4900SW (STC), Installation of a Dual Baker M1045 Cockpit Audio Systems with a Baker M1050 Cabin Paging/Chime Amplier on N660RM. The airplane was inspected in accordance with the factory recommended checklist and determined to be in an airworthy condition on January 13, 1984. Raytheon Aircraft Company acquired Arkansas Modification center, Inc., and owns STC4900SW. The STC was issued to Raytheon Aircraft Company by the Federal Aviation Administration on January 26, 1983.
All circuit breakers on panels A, B, C, and D were pulled in the cockpit. Electrical power was provided to the airplane with an electrical ground power unit. Smoke was noted under the cockpit floor and the electrical ground power unit was immediately disconnected. The source of the smoke was not determined. The nose landing gear penetrated the avionics compartment and electrical wiring was displaced.
The left side elevator control rod assembly, which connects to the elevator power control unit torque tube, was disconnected to isolate the power control unit from the elevator control circuit because hydraulic power was not available to conduct a dynamic test of the elevator control system.
The pilot's control column was pulled aft by the accident pilot to simulate the displacement used for rotation on the accident flight. The control column was pulled further aft to the nose up stops with no jamming indications. On a subsequent control column cycle, a light downward finger pressure was applied on the co-pilot's microphone jack as the pilot's control column was rotated. When the microphone jack box contacted the fairing at the control column base, the control column could not be moved further aft. This position corresponds to a control column position about 1/2 inch aft of the neutral position. The accident pilot stated the position of the pilot's control column in this jammed condition appeared to be similar with the control column position encountered during the accident sequence.
To gain access under the flight deck, the left hand fuselage access panel was removed. The left hand control column stop assembly exhibited no anomalies. A dynamic test for clearance between the pitch disconnect mechanism and the flight spoiler mechanism support bracket revealed no contact conditions. The elevator control was rotated aft and contacted the mechanical up stop and no anomalies were noted.
The right side autopilot elevator servo actuator access panel was removed and no anomalies were noted.
The right side elevator power control unit access panel was removed. The elevator power control units were inspected and no anomalies were noted. No dynamic test of the elector power control units was performed due to lack of hydraulic power.
The right side elevator flutter damper access panel was removed. The inboard and outboard elevator flutter dampers were inspected. The fluid level indicators on both dampers were indicating red. No other anomalies were noted. No dynamic test of the elevator flutter dampers was performed due to lack of hydraulic power.
The pitch disconnect handle was pulled out to disconnect the pilot's elevator control from the co-pilot's elevator control system to prevent the right hand elevator control rod from contacting the horizontal stabilizer structure.
The left autopilot servo turnbuckle access panel was removed. The left elevator servo cable turnbuckle was safetied and no anomalies were noted.
The elevator autopilot servo cable drum cover was removed for visual verification of the cable run. A dynamic test was completed. The elevator control servo was inspected for rotation and no anomalies were noted.
The left hand elevator pitch feel simulator access panel was removed. The left elevator control pitch feel simulator and elevator control rods were connected and safetied. The horizontal stabilizer input rods were connected to the pitch feel simulator unit and were safetied. The left and right pitch feel simulators were intact. No anomalies were noted. A dynamic check was conducted on the left hand elevator pitch feel simulator and no anomalies were found.
The aft elevator control quadrant was inspected and was intact. The elevator control cables had tension and no anomalies were noted. A dynamic check was completed for the aft elevator control quadrant and no anomalies were noted.
The elevator gain change mechanism access panel was removed and no anomalies were noted. A dynamic test was completed and no anomalies were noted.
The left side elevator power control unit access panel was removed. The inboard and outboard elevator power control units were inspected and no anomalies were noted. No dynamic test could be performed on the elevator power control units due to lack of hydraulic power.
The left side elevator flutter damper access panel was removed. The inboard and outboard elevator flutter dampers were inspected. The fluid level indicators on both dampers were indicating red. No other anomalies were noted. No dynamic test of the flutter dampers could not be performed due to the lack of hydraulic power.
The left autopilot servo access panel was removed. The pitch autopilot servo drive assembly was is intact. The elevator servo cable has tension. The elevator servo drum assembly was intact. The two elevator servo electrical connectors were connected. The autopilot servo cable fairing was attached to the servo cable drum assembly and floated. No anomalies were noted. The pitch autopilot servo drive assembly was removed for further examination.
The pitch autopilot servo drive assembly was sent to Chicago, Illinois to undergo computed tomography scanning. The pitch autopilot servo drive assembly was subjected to x-ray computed tomography (CT) scanning to document the internal condition of the component. The scanning was conducted on March 17, 2005. The scans were performed by Bio-Imaging Research, Inc. (BIR) under the direction of the NTSB. BIR used the ACTIS 800/450 CT system with an x-ray source strength of 420 kV and the system recorded the x-ray attenuation information on a 1024 channel cadmium tungstate linear array detector system. The data set was evaluated using the VGStudioMax software package to create a three-dimensional reconstructed image of the component. The pitch autopilot servo drive assembly images were examined for any signs of missing or damaged components or any other anomalies. No anomalies were noted.
The rear equipment bay access door was opened. The elevator control cables were intact with tension. The elevator pulleys in the rear equipment bay were verified for rotation. No anomalies were noted in the rear fuselage. The ram air tube was removed from the hydraulic fluid heat exchanger for access to inspect the tail section of the rear fuselage. The No. 2 hydraulic system reservoir indicated 15 percent without the hydraulics pressurized. The No. 1 electrical hydraulic system reservoir sight gage was not readable. The No.1 and No. 2 hydraulic reservoir ecology bottles contained 5 percent of hydraulic fluid. The No.1 and No. 2 hydraulic by-pass filter pop outs were not activated. The No.1 hydraulic system pressure accumulator indicated 1300 psi. The No.2 hydraulic accumulator indicated 1475 psi.
The right main landing gear wheel well fairing was removed. The cannon plug for the brake overheat sensor was removed. The No. 3 hydraulic system reservoir content indicator indicated 53 percent. The hydraulic system reservoir ecology bottle contained 5 percent of hydraulic fluid.
The left main landing gear wheel well fairing was removed. The cannon plug for the brake overheat sensor was removed. The hydraulic system reservoir ecology bottle contained 5 percent of hydraulic fluid.
The left and right aileron system quadrants and cable circuits were inspected and no anomalies were noted. The left and right flap power drive unit and flex shaft revealed no anomalies. The left and right anti skid control valves were inspected and no anomalies or leakage was present. The left and right volumetric brake fuse was inspected and found in the closed position and no leakage was found. The hydraulic system accumulator was inspected and the pressure contents indicator indicated 1475 psi. A general visual inspection between the left and right center wing and the lower external fuselage revealed no anomalies. The flight spoiler quadrant and the push-pull flex shaft cable was inspected and no anomalies were noted. The left and right No. 3 hydraulic pressure, case drains, and return filter bypasses were not activated.
MEDICAL AND PATHOLOGICAL INFORMATION
The airline transport rated pilot-in-command, co-pilot and 5 passengers reported no injuries. No toxicology specimens were requested from the flight crew.
TEST AND RESEARCH
The Fairchild Cockpit Voice Recorder (CVR), model GA100 was forwarded to the NTSB Recorders Division. Examination of the CVR by the NTSB determined that the recording did not include any pertinent information. A CVR group was not formed. The airplane was not equipped with a flight data recorder.
The FAA adopted a new Airworthiness Directive, AD 2005-11-04, effective May 27, 2005 for certain Bombardier airplanes modified by STC4900SW. The AD requires revising the airplane flight manual (AFM) to require repetitive visual checks of the microphone jack assemblies on both control columns to detect damage that may interfere with movement of the control column. in addition, the AD also requires modification of the microphone jack assembly, related investigative actions, and corrective actions if necessary which allows the AFM revision to be removed from the AFM. The AD was issued by the FAA "to prevent a damaged microphone jack assembly from interfering with movement of the control column, which could result in loss of control of the airplane."
The airplane, cockpit voice recorder, and servo were released to the registered owner on March 25, 2005.