HISTORY OF FLIGHT

On June 27, 2005, approximately 1225 mountain daylight time, an unregistered CGS Aviation Hawk Two Place Arrow experimental light-sport airplane was destroyed when it impacted terrain following an uncontrolled descent during the base leg of the visual approach to runway 19 at Jackson Hole Airport (JAC), Jackson, Wyoming. The commercial pilot, who was the sole occupant and owner of the airplane, sustained fatal injuries. Visual meteorological conditions prevailed, and a flight plan was not filed for the Title 14 Code of Federal Regulations (CFR) Part 91 personal flight. The local flight departed JAC at 1218.

Several witnesses at JAC reported that the airplane departed runway 19, and immediately climbed to approximately 500 feet above ground level (agl), while maintaining runway heading. The airplane then turned left onto the crosswind leg of the traffic pattern and maintained a slight climb. Two witnesses, who were traveling southbound in motor vehicles on Highway 89 (parallel to JAC runway 01/19), reported that they observed the airplane on the left downwind leg at an altitude of 700 to 900 feet agl.

One witness, who was also in a vehicle, observed the airplane in "normal flight" at an altitude of 800 to 1,000 feet flying parallel to the east of Highway 89. The witness observed the airplane turn left (west) onto the base leg for runway 19 "in normal flight and was descending but not what appeared to be abnormal." The airplane continued a nose-low descent toward the terrain. Approximately 200 feet agl, the airplane began a "steeper" nose-low descent toward terrain. The airplane then impacted terrain in a wings level attitude, bounced, nosed over, and came to rest inverted. During the descent, the witness did not noticed any movement of the elevator or other flight control surfaces, or any in-flight structural failure of the airplane.

According to Federal Aviation Administration (FAA), the pilot was in communication with the JAC air traffic control tower during the accident flight. The pilot contacted the air traffic controller (ATC) and reported he wanted to depart JAC to the east with a left turn out. Shortly after departure, the pilot contacted ATC and stated he wanted to re-enter the traffic pattern and return to JAC for a full stop landing. The controller told the pilot to report left base for runway 19. The pilot reported left base for runway 19. No further communications were received by ATC from the pilot. During the communications, the pilot did not report any problems with the airplane.

PERSONNEL INFORMATION

The pilot, age 58, who was seated in the front seat position, held a commercial pilot certificate, issued February 25, 1970, with airplane multi-engine land, airplane single-engine land, and airplane instrument ratings. The pilot also held a type rating for Cessna CE-525S jet airplanes. The pilot was issued a third-class medical certificate on March 9, 2004, with the limitation, "Holder shall possess glasses that correct for near vision."

The pilot's logbooks were not located; however, according to the pilot's March 9, 2004, medical certificate application, the pilot reported a total time of 9,400 hours, and 203 hours in the previous 6 months.

AIRCRAFT INFORMATION

The CGS Hawk Two Place Arrow was a high wing, 2-seat tandem airplane, constructed from aluminum tubing and Dacron covers, featuring dual 3-axis control. The orange and white accident airplane was configured with conventional fixed landing gear (a tricycle gear configuration was also available from the manufacturer).

The wing (31 1/2-foot span, 147 square feet) consisted of one section per side with an aileron, 3-position flap, and aluminum wing strut. Each wing contained two 5-gallon fuel tanks, a total capacity of 20 gallons. The wings are attached to the fuselage, and a Velcro secured Dacron or Lexan "gap" cover completes the upper wing surface between the two wing sections and fuselage.

The fuselage (cockpit) was an aluminum tubing construction covered with Dacron and a Lexan windscreen, which was sewn to the Dacron cover. The fuselage had four removable zipper doors which provide a full fuselage enclosure. The accident airplane was equipped with a heater which was mounted in the nose of the airplane. A single aluminum boom tube attached the fuselage and empennage. The empennage consisted of vertical and horizontal stabilizers, with rudder and elevator control surfaces. The accident airplane was originally equipped with an optional elevator trim tab. The flight control surfaces were operated via control cables. The dual stick controls were slaved via a torque tube. The torque tube assembly was routed inside the boom tube, and the rear stick was exposed through a rectangular cut-out in the top portion of the boom tube.

The CGS Hawk Two Place Arrow was powered by a 100-horsepower, water-cooled, fuel-injected Hirth model 3701 engine, equipped with a three-blade IVOPROP electric in-flight adjustable, composite propeller. The engine was mounted at the rear of the fuselage, immediately aft of the wing.

The accident airplane was built and completed in May 2005 by a CGS Aviation approved dealer located in West Virginia, in conjunction with partial assistance by the pilot. The airplane building records and logbooks were not obtained. According to the dealer and acquaintances of the pilot, the airplane had accumulated approximately 45 hours at the time of the accident.

The accident airplane was eligible for certification in the experimental light-sport aircraft category. According to CFR Part 21.191(i)(1), after January 31, 2008, an experimental certificate would not be issued for the accident airplane. The airplane was not eligible for certification in the ultralight category due to a fuel capacity of greater than 10 gallons. According to the kit manufacturer and dealer, the pilot had intended on registering the airplane with the FAA; however, he had not completed the process at the time of the accident.

METEOROLOGICAL INFORMATION

At 1215, the JAC METAR (routine aviation weather report) reported the wind from 180 degrees at 4 knots, visibility 10 statute miles, sky clear, temperature 66 degrees Fahrenheit, dew point 41 degrees Fahrenheit, and an altimeter setting of 30.03 degrees inches of Mercury.

AERODROME INFORMATION

The Jackson Hole Airport, JAC, is a public, controlled airport located approximately 7 miles north of Jackson, Wyoming, at 43 degrees, 36.44 minutes north latitude, and 110 degrees, 44.265 minutes west longitude, at a surveyed elevation of 6,451 feet. The airport features one asphalt runway, Runway 1/19, which is 6,300 feet long and 150 feet wide.

WRECKAGE AND IMPACT INFORMATION

The accident site was located in sage brush-covered flat terrain approximately 575 feet west of Highway 89, and 3/4-mile north of the approach end of runway 19. Initial examination of the wreckage and accident site was completed by Grand Teton National Park personnel. With assistance from a local mechanic, the accident site was documented by the National Park Service on-scene coordinator. A global positioning system (GPS) receiver reported the location as 43 degrees 37.36 minutes north latitude and 110 degrees 43.38 minutes west longitude, at an elevation of 6,400 feet. The wreckage was distributed along a measured magnetic heading of approximately 320 degrees. The initial impact point was consistent with the fuselage and the main landing gear. The main wreckage came to rest inverted approximately 50 feet from the initial impact point. The main wreckage consisted of the left and right wings, boom tube, empennage, and fuselage tubing. The airplane wreckage was transported to a vacant hangar at JAC for further examination by the NTSB.

On June 28, 2005, the wreckage was examined by the NTSB investigator-in-charge, two FAA inspectors, and the on-scene coordinator with the Grand Teton National Park. Examination of the wreckage revealed the left wing fabric was torn, the main spar intact, and the wing support structure was crushed aft. The forward and rear struts were fractured and remained attached to their respective wing and fuselage attach fittings. The left flap and aileron remained attached to the wing. Control continuity was established to the flap, and the aileron cable was separated and exhibited a random raveling of individual wires typical of an overload failure.

The right wing fabric was torn, the main spar intact, and the wing support structure was crushed aft. The forward strut was intact, and the rear strut was fractured at its midpoint. The struts remained attached to their respective wing and fuselage attach fittings. The right flap and aileron remained attached to the wing. The right flap control tube was separated at the flap attach point, and the aileron cable was separated and exhibited a random raveling of individual wires typical of an overload failure.

The boom tube was separated near the leading edge of the vertical stabilizer. The vertical and horizontal stabilizers were intact, and the elevator and rudder remained attached to the stabilizers. Elevator and rudder control continuity was established from the control surfaces to their respective cockpit controls. The elevator trim tab was not installed (According to friends of the pilot, the pilot removed the trim tab while transporting the airplane via truck after an incident in Nebraska).

The fuselage was fragmented and destroyed. The front seat floor was crushed aft, and the rear seat floor had not been installed. The control stick assembly torque tube was fractured aft of the front control stick and forward of the aileron crank. The rear control stick was not installed. The front seat was bent and crushed, and the rear seat had not been installed. The front seat aluminum back displayed an impact mark consistent with the engine water pump. The instrument panel was bent and separated from the fuselage structure. The airspeed indicator needle displayed 120 knots, and the vertical speed indicator displayed greater than 1,500 feet per minute descent. The landing gear saddle and legs were separated from the boom tube. The fuselage covering and gap cover (located overhead between the wings) had not been installed. A Lexan windscreen, which was riveted onto the fuselage structure, was separated from the fuselage.

The engine was separated from its mounting structure. The three-blade composite propeller was splintered and fragmented, and remained attached to the engine crankshaft. The propeller and crankshaft was manually rotated; continuity and thumb compression were established throughout the engine.

PATHOLOGICAL INFORMATION

An autopsy was performed on the pilot by Western Wyoming Pathology, Jackson, Wyoming, on June 28, 2005. Blood specimens for toxicological tests were taken from the pilot by the medical examiner. According to the autopsy, the cause of death for the pilot was blunt force trauma.

The FAA's Bioaeronautical Sciences Research Lab, Oklahoma City, Oklahoma, examined the blood specimen taken by the medical examiner. Toxicological tests performed on the pilot were negative for all screened drugs and alcohol. Carbon monoxide and cyanide tests could not be performed.

TEST AND RESEARCH

A Garmin GPSmap 296 receiver, located at the accident site, was examined by the NTSB's Vehicle Recorders Laboratory in Washington, DC. Due to damage sustained during the impact sequence, the receiver was sent to the manufacturer for evaluation and data extraction. Data extracted from the unit included 12 individual trackpoints recorded on June 26, 2005, between 0607 and 1148. The locations of the trackpoints were within close proximity to JAC. No data was recorded on the day of the accident.

On October 19, 2005, at the NTSB Materials Laboratory, Washington, DC, the forward portion of boom tube and enclosed controls, the front control stick assembly, the front bearing block with a portion of the fuselage, a damaged inspection mirror (found in the wreckage debris area), and an exemplar inspection mirror were examined. The examined portion of the boom tube was approximately 16 inches long and contained some of the elevator and aileron controls, and a reinforcing frame identified as an "H" member. The rear control stick protrudes out of a rectangular cutout on the upper surface of the boom tube and the aileron crank protrudes from the forward end of the boom tube. The front control stick is located forward of the front bulkhead.

When the cockpit flight control components are properly assembled, the torque tube can only rotate in the front and rear bearing blocks as two locking collars, forward and rear, to prevent any longitudinal motion. The rear bearing block is bolted to the boom tube and the front bearing block is bolted to a fuselage tube. The front and rear control sticks are connected to the torque tube at pivot points, and cables attached to the front control stick are connected to the elevator control surface. The elevator tube connects the front and rear control sticks to each other so that forward and aft motion of either control stick controls the elevator. The aileron crank is bolted to the torque tube so lateral motion of either control stick rotates the torque tube, thereby moving the crank. The crank is connected to the ailerons by cables.

Boom Tube

The boom tube extends from slightly forward of the forward bulkhead to the empennage and is the main structural member of the aircraft. Examination of the boom tube revealed that two doublers and a reinforcing plate had been riveted to the tube. The first doubler and was wrapped around the boom tube and attached with 9 rivets. The second doubler was wrapped around the first doubler; started at the forward bulkhead and extended approximately 12 inches to the rear. A reinforcing plate, which also started at the forward bulkhead, extended 9 inches to the rear and wrapped around the second doubler. The second doubler was attached to the boom tube with 38 rivets, of which 19 were common to the reinforcing plate, and 4 were common to the internal "H" member. The "H" member consisted of two longitudinal tubes and two lateral tubes welded together to form an "H" shape. The "H" member was riveted into position, horizontally across the boom tube, to reinforce the boom tube at the landing gear saddle location. The "H" member, painted gray in color, was centered approximately 10 inches from the front bulkhead. According to a photograph provided by the accident airplane builder, the original "H" member was painted white, and the second doubler and reinforcing plate were not installed. The original "H" member was replaced during a repair, which was completed by the owner, after a hard landing incident (See Additional Information section of this report).

The forward lateral tube of the "H" member had been deformed and bent to the rear and upwards. The plating at the lower portion of the rear control stick had been damaged, consistent with contact with the "H" member. An indentation was observed, and plating was missing, at the forward lower portion of the rear control stick, consistent with contact of these two areas together. Both sets of damage marks are consistent with the torque tube being moved to the rear.

Examination of the rear edge of the rectangular cutout for the rear control stick revealed disturbed paint at three locations with no material deformation.

The interior surface of the boom tube was examined and revealed a number of scratches that, in some case, had removed the paint to reveal bare material. One significant scratched area, was located above the "H" member. This area, was longitudinally oriented and accompanied by circumferential surface scratches that extended downward. A second significant area was located forward of the rear control stick and below the level of the "H" member. This area consisted of two longitudinally oriented surface scratches with circumferential surface scratches oriented downwards from the lower of the two longitudinal scratches.

Before the removal of the control assembly from the boom tube, repeated attempts were made to find an orientation of the exemplar mirror where the controls could be operated, and the damage to the accident inspection mirror might be approximately duplicated. The attempts were unsuccessful. The rear bearing block attachment bolts were removed and the control assembly extracted. Adhesive tape was utilized to pick up any particles that may have been trapped between the bearing block and the boom tube. The tape was examined and revealed minute tan colored granules, consistent with sandy soil, and one small white metal particle that appeared to be an aluminum chip from drilling operation.

Continued internal examination revealed a rippled surface on the lower surface of the boom tube at the forward end. Examination of the corresponding outer surface revealed no indications of any ripples.

Aileron and Elevator Controls

The aileron crank, which was attached to the torque tube, was no longer perpendicular to its hub, and bent in the forward direction. The torque tube was bent downward in the vicinity of the hub, and fractured at the forward side of the aileron crank hub. The fracture faces revealed that they were inclined at 45-degrees, consistent with overload. The aileron crank attaching bolt displayed a sheared end, consistent with the circumferential displacement of the hub relative to the torque tube. The rear locking collar was located adjacent to the aileron crank hub, further forward of the installation instructions placement, and could not be moved by hand.

In order to examine the torque tube, the rear bearing block was slid rearwards. The screw in the rear locking collar had to be loosened before it could be slid rearwards. The torque required to loosen the locking collar screw was measured at 3-inch pounds.

Examination of the torque tube revealed a distinct bend, longitudinal smears, surface discoloration, and circumferential features. A band of light longitudinal smearing, approximately 1/2-inch wide, with a circumferential origin, was the original location for the rear inside edge of the rear locking collar. Immediately aft, the surface of the torque tube displayed a lightly discolored band, approximately 1 inch wide, which was consistent with the original location of the rear bearing block, which was 1 inch thick. For approximately 5/16-inch forward of the locking collar's original location, the surface of the torque tube was almost featureless with the exception of an area approximately 1/4-inch by 1/4-inch consisting of faint circumferentially oriented scratches, and two distinct features that were oriented circumferentially.

Longitudinal bands of smeared surface initiated approximately 1/4-inch forward of the circumferentially orientated scratches. The bands of smearing were continuous for several inches up to a narrow band displaying two faint circumferential marks. A circumferentially orientated, narrow band of unsmeared surface was located approximately 2-1/2 inches forward of the locking collar's original position. The longitudinal bands of smearing continued forward of the narrow band, but were not continuous across this position. The rear edge of a band of discoloration that was approximately 1/2-inch wide was located immediately to the rear of the narrow band of unsmeared surface. The discolored band was almost continuous around the torque tube and was evident on both unsmeared and smeared surfaces. The total displacement of the rear locking collar was measured at 3.25 inches.

Diameter measurements were taken in the vicinity of the narrow band of unsmeared surface located approximately 2-1/2 inches forward of the locking collar's original position and found to be 0.997 inch on the vertical axis and 1.002 inches on the horizontal axis, consistent with deformation imparted to a tube when it is bent. Measurements of the torque tube outside diameter, at undamaged locations, were found to be consistent at 0.999 inch.

Examination of the aileron crank revealed two areas devoid of paint and a distinct contact mark located on the rear face. The areas appeared to be the locations of maximum deformation associated with the bending of the aileron crank and the discoloration matched the location of the weld onto the opposing face. A closer view of the contact mark revealed that it consisted of two distinctly featured areas, one displaying a lightly rubbed area with laterally oriented arched lines on the surface and one displaying a thin band of bare metal with surface scratches orientated on the longitudinal axis of the crank with a second thin band of bare metal extending almost perpendicular and to the right of the longitudinal band.

A laterally oriented smeared surface was noted on the front bulkhead upper tube. The smeared surface would be aligned with the lower portion of the bare metal with the aileron crank in its normal position, perpendicular to the torque tube.

The front control stick was still attached to a portion of the torque tube. The front locking collar was still affixed to the torque tube, forward of the control stick. The portion of the torque tube forward of the locking collar is normally located in the front bearing block. Smeared material on the rear face of the locking collar corresponded with a mark on the forward face of the control stick clevis, consistent with impact as the torque tube was moved to the rear. In order to examine the forward end of the torque tube, the locking collar was removed. The torque required to loosen the front locking collar screw was measured at 25-inch pounds.

For comparison, the front and rear locking collar inner surfaces were examined. The inner surfaces of both collars exhibited circumferential smears that matched the pattern of machining marks. On the rear locking collar, bands of longitudinally oriented smears and scratches obscured some of the circumferential smearing. The screws were removed from both locking collars and the threads were found to be clean and intact, without visual evidence of any type of locking compound.

Torque Tube Testing

Tests were devised in an attempt to produce surface features that could be compared with the features on the accident aircraft's torque tube and to quantify the force required to move a locking collar on a torque tube. The aircraft kit manufacturer provided three new torque tubes and two new locking collars for the test. The exemplar tubes were identified as "A", "B", and "C".

The first series of tests consisted of installing an exemplar locking collar on an exemplar torque tube using one of the accident airplane locking collar screws and torque loading the screw (the accident airplane screw was used because the exemplar screws contained a locking compound that would have increased the tare torque between the screw and collar threads). The torque tube and locking collar assembly was located in a test frame and a lever arm was attached to the test frame and the end of the torque tube. A digital force gauge was attached to the other end of the lever arm to measure the load, through the leverage provided, that was required to overcome the frictional lock produced at that particular torque loading on the locking collar. The locking collar was torque loaded at 5, 10, 15, 20, and 25 inch-pounds at individual locations on each torque tube.

An examination of the torque tube surfaces produced by testing revealed that the discoloration at 5 inch pounds was almost negligible, but the discoloration darkened as the torque loading increased. The surface discolorations produced were almost uniform for the individual torque loading but did not produce the smearing observed on the accident torque tube.

The test results for the torque tubes A, B, and C, are presented in the following chart with an average figure and with the force applied to the stick (for forward and aft movement) to overcome that average figure.

Torque collar Tube A Tube B Tube C Avg. Load Stick Force Stick Force
screw (in-lb) Load (lb) Load (lb) Load (lb) (lb) Forward (lb) Aft (lb)

5 45 101 81 76 13 10
10 126 180 72 126 22 16
15 191 297 139 209 36 27
20 231 335 187 251 43 32
25 385 435 259 360 62 46

A review of the Human Factors Design manual revealed that, with a slight bend at the elbow, an average man could exert a forward push of 33.6 pounds, and a rearward pull of 44.8 pounds, on the control stick.

The second series of tests consisted of installing the locking collar on an unused end of an exemplar torque tube using one of the locking collar screws from the accident airplane, torque loading the screw and moving the collar rapidly. The collar was moved rapidly by installing a stock tube on the torque tube to contact the locking collar. The stock tube was then struck by a hammer. The screw in the locking collar was torque loaded at 5, 15, and 25 inch-pounds at individual locations on two of the exemplar torque tubes. The features produced by the rapid movement of the locking collar exhibited bands of longitudinal smearing that had been produced on a discolored surface. Although the smearing was not as pronounced as that observed on the accident torque tube, it did display a similar texture under all three torque loading conditions.

The third test consisted of deforming the unused end of one torque tube to duplicate the dimensions measured at the bend on the accident torque tube. A locking collar was installed at an adjacent non-deformed location and the screw was torque loaded to 5 inch-pounds. A force of 290 pounds was required to pull the tube through the locking collar.

Aircraft Kit Document Review

A search of all the sections in the assembly instructions did not reveal any torque loading criteria for any bolted attachment or screw, such as the screw in the locking collars.

The elevator installation section of the assembly instructions described the cable tensioning procedure as "Raise or lower the tail boom until stabilizer is at 0 [degrees] angle or horizontal with the floor. With a bubble protractor, check the elevator up and down travel. The minimums are 25 [degrees] up and 15 [degrees] down. Adjust cables so that when the stick is in a neutral position, the stabilizer is neutral. There should always be a slight amount of tension in the cables when the elevator is full up or down. If one cable goes slack, snug it up. If you run out of turnbuckle adjustments, then back off the turnbuckles and pick up the slack with the cable clamps. When all is adjusted, be sure that 4-5 threads are showing on the turnbuckles before safety wiring them."

The cable tensioning procedure section of the owner's manual (page 9) described the procedure for the elevator as "Lock the elevator to the horizontal stabilizer. See sketch 1. At proper tension it should take 10 lbs. pressure to deflect the control stick 2 inches at a point 13 inches up from the pivot point. See sketch 2". Sketch 1 illustrates the locking method and sketch 2 illustrates an aft movement of the control stick, which would move the elevator up.

Any pretension on the elevator cables creates a load that pulls the torque tube in the aft direction. Assuming that the down elevator cable has no tension and that the up elevator cable is resisting the aft load on the stick, the rigging procedure would produce an aft load in the torque tube of 52 pounds [(10 x 13) divided by 2.5]. Therefore, until the stick force exceeds 10 pounds, the aft load on the torque tube would remain relatively constant.

Inquiries with the manufacturer revealed that the elevator control system is what is termed as a pull/pull system where one cable is loaded above the system tension when operated. Performance characteristics for the elevator were also received from the manufacturer with calculations indicating that the force on the elevator surface, at 100 mph, would be 203.3 pounds (the angle of attack, relative to the oncoming air, defines the force on the elevator but was not indicated). Further calculations using the control component dimensions indicated that, under the same conditions, the load in one of the elevator cables would be 385.2 pounds and that the control cable used (3/32 inch diameter, 7x7 construction, stainless steel) had a breaking strength of 920 pounds. The manufacturer had also performed a test on an aircraft where elevator rigging had been carried out in accordance with the owner's manual and then the turnbuckles were loosened. The manufacturer stated that the control stick struck the seat when the turnbuckles were extended 3/16 inch.

Accident Inspection Mirror

The inspection mirror was found at the accident site in the debris field. The glass portion of the mirror was missing and no residual glass could be found in its frame. The mirror frame had been bent almost double with the axis of the bend located diagonally across the frame. Examination of the mirror frame revealed that two indentations were located on opposing edges of the mirror frame and an area of mechanical damage was observed at the bottom of the mirror frame.

Microscopic examination of the indentations revealed that they were predominantly located on a flat portion of the mirror edge and there was minimal deformation of the adjacent radii. The indentations contained smears of gray colored paint. The mechanical damage appeared as a portion of the back that had been bent inwards, with a tear at one end of the bend. Microscopic examination of the mirror frame did not reveal any particles of glass at the edges. Further microscopic examination revealed no glass particles in the cardboard backing piece or in the edges when the backing piece was removed.

In order to determine if the mirror had been somehow trapped in the boom tube, in a position not envisaged earlier, the mirror frame was removed from the handle and with samples removed from the boom tube, the elevator tube and aileron crank, was subjected to Fourier Transform Infra Red (FTIR) analysis. The FTIR spectra indicated that the paint in the indentations on the edge of the mirror frame did not match the paint on the elevator rod, the aileron crank, or the inside of the boom tube.

ADDITIONAL INFORMATION

According to the CGS dealer, prior to accepting the completed airplane in West Virginia, the pilot performed several flights in the accident airplane. During one flight, the pilot left a cordless drill on the top of the engine prior to takeoff. During the flight, the cordless drill fell off the engine and contacted the composite propeller. The propeller sustained damage, and the pilot successfully landed the airplane without further incident. The damaged propeller was removed and replaced. The dealer reported the pilot, on previous flights, had "no trouble handing the aircraft, seemed comfortable with the aircraft."

In May 2005, the pilot departed West Virginia for Jackson, Wyoming, in the accident airplane. According to witnesses, the airplane landed at Cram Field Airport (BUB), Burwell, Nebraska, to refuel. During takeoff from BUB, a piece of luggage fell from the rear of the airplane. The pilot returned to the airport to retrieve the luggage. A short time later, the pilot contacted the witnesses and reported that he had damaged the landing gear during the landing at BUB.

According to a friend of the pilot, the pilot stated that he did not properly latch the flaps during the landing at BUB. While over the runway, with the flaps set in the full position, the flap handle disengaged, and the flaps went to the full retracted position. Subsequently, the airplane landed hard on the runway damaging the landing gear and fuselage structure. The pilot then partially disassembled the airplane, transported the airplane via truck to Jackson, and performed repairs to the airplane.

After the incident in Nebraska, the kit manufacturer and the dealer offered to personally assist the pilot in Jackson with the repairs to the airplane; however, the pilot declined the offer for on-scene assistance. The pilot received a various number of new replacement components from the kit manufacturer. Several of the pilot's friends assisted the pilot during various phases of the repairs. At some point during the repairs, the pilot removed the rear floorboard, rear control stick, rear seat and restraint system, the nose-mounted heater, the fuselage cover, and the gap cover. Because the windscreen was sewn into the fuselage covering, the pilot crafted a Lexan windscreen and riveted it onto the fuselage structure. The pilot did not replace the fuselage cover, gap covers or doors after the repairs were completed. The pilot's friends also reported that the pilot had been using an inspection mirror to search for metal shavings in the rectangular cutout of the boom tube during the days preceding the accident.

The accident flight was the third flight since the hard landing in Nebraska and subsequent repairs. During the first flight, the pilot experienced a "tail-heavy" feeling, and he thought that the removal of the heater had shifted the center-of-gravity (CG). The pilot replaced the smaller aft-mounted (near rear seat) battery with a heavier battery, which the pilot mounted in the nose of the airplane. The pilot reported no further problems to friends and family regarding a CG issue. During the second flight, the pilot experienced an overheating engine. The pilot removed the engine cooling air scoops from the fuselage cover (which were originally sewn onto the fuselage cover), and directly attached the air scoops to the fuselage structure with rivets. The pilot reported no further discrepancies with the airplane to his friends.

The wreckage was released to a family representative.