HISTORY OF FLIGHTOn August 5, 2024, about 1011 eastern daylight time, a Cessna 750 airplane, N750GB, was destroyed when it was involved in an accident near Jamestown, New York. The pilot was seriously injured, and the copilot sustained minor injuries. The airplane was operated by Access Stripe, Inc as a Title 14 Code of Federal Regulations Part 91 business flight. The purpose of the flight, which was planned from Chautauqua County Dunkirk Airport (DKK), Dunkirk, New York, to Fort Lauderdale Executive Airport (FXE), Fort Lauderdale, Florida, was to pick up airplane parts and a mechanic for work on another company airplane. ADS-B flight track information, air traffic control (ATC) communication data, and cockpit voice recorder (CVR) information revealed that the pilots obtained an instrument flight rules clearance to FXE before departing from DKK. There were no checklist callouts heard on the CVR recording as the crew taxied the airplane to the runway for departure about 1003. About 1006, while the airplane was climbing to cruise altitude, the copilot stated that he could smell “something burning” and that it was “electrical,” then stated that it had “dissipated.” About one minute later, the copilot stated, “I definitely smell fire.” Immediately thereafter, audible alarms from various crew alerting system (CAS) messages could be heard, and the crew decided to divert to Jamestown Airport (JHW), Jamestown, New York. Both pilots later stated that there was an odor of smoke but no smoke was visible. The crew attempted to contact the controller but received no reply. They then transmitted the emergency transponder code as they initiated a descent toward JHW. In a postaccident interview, the pilot stated that, immediately after detecting the smell of an electrical fire, they started to receive warnings regarding the electrical system. He stated that the flight instruments were displaying “red X’s,” that the number of cautions and warnings was quickly increasing, and that he was starting to lose trim control. He attempted to use secondary trim but was unsuccessful. The copilot stated that, as they descended through about 10,000 ft, he heard the “clacker” for the pitch trim and that the airplane was “trimming down and accelerating… well over 250 [knots] with the nose trimming down.” The copilot said the “Master Caution” and panel segments illuminated, along with other CAS messages. He tried to contact the controller before he noticed that second communication radio had failed, and that the primary flight display was showing “red x’s.” The controller made several unsuccessful attempts to contact the crew, then began transmitting on the emergency frequency and noted the airplane’s emergency transponder code as it descended through 7,800 ft mean sea level (msl). The controller announced that the airplane was directly over JHW; however, the crew did not respond. Shortly thereafter, the controller established intermittent contact with the flight crew on the emergency frequency; however, the transmissions were “very garbled.” At 1008:33, the crew indicated that they had JHW in sight and would be landing there. Flight track data revealed that, after takeoff, the airplane climbed on a southerly track and leveled about 10,000 ft msl for about 1 minute before it entered a descent. The airplane passed west of JHW and initiated a left, descending, decelerating, 270° turn to align with runway 25. The final target showed the airplane aligned with the runway at 1,625 ft msl and 150 kts. In the postaccident interview, the pilot stated that he had difficulty slowing the airplane to approach speed and that they were not able to fully extend the wing flaps. The airplane then “dropped” short of the runway and impacted the ground hard. Witnesses to the accident did not report any trailing smoke from the airplane and noted that the airplane’s final approach appeared normal to other airplanes they had seen land on that runway, except that it landed short of the runway threshold and “hard.” Video surveillance from a business located north of the runway captured the airplane briefly while it was on final approach. There was no visible smoke trailing the airplane. PERSONNEL INFORMATIONThe pilot held an airline transport pilot certificate with ratings for airplane single-engine land and sea and airplane multi-engine land and sea, with multiple type ratings. His most recent Federal Aviation Administration (FAA) second-class medical certificate was issued on September 6, 2023. He reported 17,000 hours of flight experience on that date. The copilot held an airline transport pilot certificate with a rating for airplane multi-engine land with multiple type ratings, and private pilot privileges for airplane single-engine land. His most recent FAA first-class medical certificate was issued on June 7, 2024, and he reported 12,950 total hours of flight experience on that date. AIRCRAFT INFORMATIONAccording to maintenance records, the airplane was manufactured in 2014 and was powered by two Rolls-Royce AE3007C2 turbofan engines. On July 11, 2024, the airplane underwent maintenance at a manufacturer service center. The maintenance included repair of the No. 2 battery, a functional check of the horizontal stabilizer trim actuator, navigation equipment checks, corrosion inspections of the stabilizers and wings, and multiple other checks. A work report described detailed inspections for corrosion within the horizontal tail rear spar and elevator cover, and the vertical tail rear spar and rudder cover. The airplane received additional inspections of the aft vertical stabilizer and trailing edge and a lower rudder assembly corrosion inspection, as well as detailed corrosion inspections of the aft wing spar and wing fuel tanks. The report also detailed inspections of the fuselage that included the aft cabin wing attach frames and a functional test of the nickel-cadmium (Ni-Cad) batteries. The report explained that the left battery (No. 1) had thermal runaway on 5 cells, but was charged and serviced in accordance with the manufacturer’s manuals. The right battery (No. 2) failed a capacity test and it was serviced in accordance with the manufacturer’s manuals. The horizontal stabilizer trim actuator was lubricated and tested, and the primary and secondary stabilizer trim switches were functionally tested. The main electrical junction box (J-Box) cover was missing 4 retaining studs and 32 stud retaining rings. The missing rings and studs were replaced. A cannon plug was found with a pulled wire at a pass-through hole in a floor panel on the right side raised aisle, about 10 feet from the aft wall. AIRPORT INFORMATIONAccording to maintenance records, the airplane was manufactured in 2014 and was powered by two Rolls-Royce AE3007C2 turbofan engines. On July 11, 2024, the airplane underwent maintenance at a manufacturer service center. The maintenance included repair of the No. 2 battery, a functional check of the horizontal stabilizer trim actuator, navigation equipment checks, corrosion inspections of the stabilizers and wings, and multiple other checks. A work report described detailed inspections for corrosion within the horizontal tail rear spar and elevator cover, and the vertical tail rear spar and rudder cover. The airplane received additional inspections of the aft vertical stabilizer and trailing edge and a lower rudder assembly corrosion inspection, as well as detailed corrosion inspections of the aft wing spar and wing fuel tanks. The report also detailed inspections of the fuselage that included the aft cabin wing attach frames and a functional test of the nickel-cadmium (Ni-Cad) batteries. The report explained that the left battery (No. 1) had thermal runaway on 5 cells, but was charged and serviced in accordance with the manufacturer’s manuals. The right battery (No. 2) failed a capacity test and it was serviced in accordance with the manufacturer’s manuals. The horizontal stabilizer trim actuator was lubricated and tested, and the primary and secondary stabilizer trim switches were functionally tested. The main electrical junction box (J-Box) cover was missing 4 retaining studs and 32 stud retaining rings. The missing rings and studs were replaced. A cannon plug was found with a pulled wire at a pass-through hole in a floor panel on the right side raised aisle, about 10 feet from the aft wall. WRECKAGE AND IMPACT INFORMATIONThe airplane initially impacted terrain and runway approach lighting 200 ft before the approach end of runway 25. The airplane continued along the runway centerline, beyond the runway 13/31 intersection, traveled off the left side of runway 25, and came to rest about 225 ft from the runway’s left edge and about 2,150 ft beyond the initial impact point. The airplane came to rest facing about 300° magnetic. The cockpit, cabin area, inboard sections of both wings, empennage, and the root section of the vertical stabilizer were consumed by postimpact fire. The upper portion of the vertical stabilizer and the horizontal stabilizers were intact. The empennage structure collapsed as a result of the postimpact fire. The area of the vertical stabilizer was consumed by postimpact fire. Control continuity was established from the cockpit to all flight control surfaces except for the left speed brake and spoiler cables, which displayed thermal damage consistent with the postimpact fire. All primary control surfaces were still attached to the airframe in their respective locations. The fuselage exhibited significant thermal damage. The tubular structure was visible from the aft bulkhead and lavatory area forward. The passenger seats and galley were recognizable but significantly damaged. The main stairs were extended and were the exit point for the flight crew. Portions of the roof had collapsed above the passenger seats and galley area. The nose of the airplane exhibited thermal damage to the internal electrical components. The underside of the fuselage displayed damage consistent with skidding down the runway and the postimpact fire. The fuselage aft of the rear bulkhead, between the two engines, had collapsed, and the baggage compartment was destroyed. The interior of the cockpit was consumed by fire. The primary and multifunction flight display panels and their respective data cards were destroyed. The master caution lights and the co-pilot control yoke were retained for further examination. The engines were examined visually and also displayed thermal damage consistent with postimpact fire; there was no evidence of pre-impact mechanical anomalies. The copilot stated that he monitored engine function throughout the flight, that there were no CAS messages that related to the engines, and that neither engine malfunctioned. The right wing was partially separated from the fuselage and consumed by fire at the midpoint to the wing root. The left wing was intact and remained attached to the fuselage, but exhibited thermal damage across its entire span. Both main landing gear and the nose landing gear separated during the accident sequence. The left main and nose landing gear exhibited no thermal damage. The right main landing gear exhibited sooting on the strut and light thermal damage to the outward-facing surface of both tires. The gear was extended at the time of impact. The co-pilot stated that the airplane was fueled before departure. The odor of jet fuel was present at the accident site. Both wings contained an unmeasured quantity of fuel that was visible during recovery of the airplane from the accident scene. Fuel control switches in the cockpit were destroyed. Both main fuel tanks and the center tank were breached. Numerous fuel panels separated from the center tank during the accident sequence. The fuel tanks and components at both wing roots both exhibited thermal damage. The wing fuel filler caps were intact and the single-point refueling panel separated from the airplane and was damaged by fire. ADDITIONAL INFORMATIONAn NTSB systems specialist conducted an examination focused on the airplane’s electrical wiring and system architecture to identify common wire locations that, if damaged, could account for the reported flight deck effects. The accident airplane was equipped with a split-bus direct current (DC) electrical architecture designed to distribute electrical power generated by two engine-driven DC generators, alternating current (AC) alternators through Transformer Rectifier Units (TRU), and DC batteries to aircraft electrical buses and avionics systems. Under normal operating conditions, the left and right electrical systems operated independently. Each electrical system was powered by its respective engine-driven generator and supported by a dedicated aircraft battery located in the aft fuselage. Electrical power distribution was managed through multiple electrical buses, including battery buses, emergency buses, avionics buses, and main electrical buses. These buses provided electrical power to avionics equipment, communication radios, flight guidance systems, and flight control interface systems. Electrical distribution components associated with these buses were installed within the aft electrical junction box located in the baggage compartment area. From this distribution point, electrical power and avionics interface signals were routed forward through wiring harnesses that supplied cockpit avionics equipment and flight control interface components. Because this junction box served as a central electrical distribution point, several aircraft systems relied on wiring routed through common harness corridors before those circuits branch toward their individual subsystems. Two wiring bundles, as well as a communications coaxial cable, were found beneath the aft baggage compartment floor in the area of this electrical distribution corridor, near the aft electric junction box and the No. 2 battery. One of the wiring bundles was identified as containing multiple horizontal stabilizer trim system-related circuit wires, as well as electrical power supply wiring for the secondary stabilizer trim controller. The other wiring bundle contained stabilizer position sensor signal wiring routed to the primary and secondary printed circuit boards. These signals provided stabilizer position feedback used by the flight control and trim systems. The communications coaxial cable provided the antenna feed for the No. 2 very high frequency communication system. According to the airframe manufacturer, the appearance of multiple CAS messages within a short period of time, including autopilot, trim, and communication system indications, was consistent with the way the installed avionics systems would respond to the loss of electrical power or data communication. Because several of these systems relied on wiring routed through the aft fuselage distribution corridor before branching into individual components, a localized disruption affecting that corridor may result in multiple systems detecting the loss of valid inputs simultaneously. The airframe manufacturer also described that the baggage compartment was normally open to the cabin through the isolation valve and shared airflow with the environmental system. FLIGHT RECORDERSThe airplane was not equipped with a flight data recorder, nor was it required to be. The airplane was equipped with an aircraft recording system (AReS) II, and its storage chip was recovered; however, the extent of the thermal damage to the chip precluded data recovery. Among other parameters about the flight, this device would have recorded the CAS messages displayed to the crew. The cockpit voice recorder (CVR) was recovered and successfully downloaded. It captured three flights, with the first flight recording beginning as the accident crew were flying a descent into DKK. The second flight was the first flight on the day of the accident. The third flight recorded was the accident flight. During the second recorded flight, the crew returned to DKK after departure following scheduling changes. There was discussion between the crew of the urgency to depart for the next flight, and the crew discussed that they would fuel the airplane and immediately depart. The recording concluded with the crew starting the auxiliary power unit (APU) once on the ground after landing. The recording indicated that the crew did not complete any checklists for the duration of the second flight. FIREA fire specialist from the NTSB Materials Laboratory responded to the accident site and attended subsequent follow-up examinations of the wreckage. Approximately two-thirds of the top surface and leading edge of the right wing was missing/melted away. The remaining wing support structure and lower surface exhibited thermal-related warpage with areas of blistered paint. There was no evidence of airflow-driven soot or smoke patterns on the remaining wing skin or winglet. The left wing exhibited heavy sooting, blistering paint, and warpage over the remaining exterior surfaces. There was no evidence of airflow-driven soot or smoke patterns on the remaining wing skin or winglet. The exterior skin and components inside the nose cone were sooted. The cockpit area was destroyed by fire, including all interior components and the pilot and copilot’s seats. The seat frames had collapsed. The fuselage skin and structure in this area were melted down to the lower window frames, with the right side sustaining more damage than the left side. The windshield and all the windows were broken out of the frames and the remnants exhibited sooting and various levels of thermal damage. The passenger area sustained similar damage. The crown of the airplane was missing in the area around the door. The rest of the crown in this area was partially consumed and collapsed into the passenger area. The interior components exhibited significant thermal damage. The galley area was destroyed. The passenger seat upholstery was mostly consumed but the frames were still intact. The empennage and tail section sustained heavy thermal damage. The right and left elevators were intact. The right elevator was melted and exhibited heavy sooting, paint blistering, and skin warpage. The left elevator exhibited airflow-driven sooting. The rest of the tail was intact and exhibited sooting, paint blisters, some localized melting, and skin warpage. The area aft of the pressure bulkhead was destroyed, with only a section of the lower skin remaining. One section of separated hydraulic line appeared to have a foreign material adhered to the surface. The line section was retained for further examination. The entirety of the airplane wiring could not be accounted for on scene and in subsequent wreckage examinations. Most of the remaining wire bundles were intact and showed no obvious signs of electric arcing. A complete section and several pieces of wire conductor were recovered from the wreckage that exhibited arcing, welding, and beading. SURVIVAL ASPECTSBoth pilot seats were equipped with a 5-point harness. The emergency locator transmitter (ELT) did not activate and was consumed by fire. According to the flight crew, they initially attempted to egress the airplane via the main entry doorway. The co-pilot recalled attempting to open the main cabin door, but he was unable to open it and moved back into the cabin toward the emergency exit. He noted that when he arrived at the emergency exit, fire was visible out the window. At this point, he witnessed the pilot attempting to open the main entry door, which was ultimately successful. TESTS AND RESEARCHThe master warning/caution lights were examined using a planar X-ray imaging system to determine if there were light bulbs within the annunciator lights. The radiographs showed that none of the annunciator lights had intact bulbs nor filaments; therefore, filament status could not be determined. The co-pilot yoke switch was X-ray imaged to determine the internal trim switch position. The yoke switch sustained thermal damage during the postaccident fire, and the outer case was deformed. The radiograph showed that the deformation of the outer case was pressing on the internal switch components, which precluded determination of the switch position. A section of titanium hydraulic line had no fittings on either end. The fracture surfaces on the separated ends were consistent with overstress. There was a small globular silver metallic deposit on the surface of the line section. The metallic deposit was examined using an x-ray fluorescence alloy analyzer. The deposit was found to be aluminum. There were no other deposits or material transfers on the exterior surface. A large section of wire bundle was recovered from the empennage during the on-scene wreckage examination. The melted remnants of a cannon plug were attached at one end. The other end of this section had separated from the mating end. The bundle sustained heavy thermal-related damage. At the center section of the bundle, a 3-inch-long area of the metal braid mat was covered with a copper-colored metallic splatter. Several smaller areas (approximately 1 inch or less in length) of the metallic copper splatter were found on the interior side of the braid. Brittle, carbonized residue was present around each individual conductor. No other sign of wire insulation was present. No identifying markings were found on the outer covering of the bundle or on the individual conductors. In the area adjacent to the splatter pattern on the braiding, several conductors (approximately 20 to 25 inches from the cannon plug) exhibited welding of the conductors, narrowing and necking on fractured conductor ends, and signs of erosion on the surface. A smaller area with similar damage was found approximately 3 inches from the cannon plug. The remaining conductors on the opposite side of the cannon plugs exhibited beading and melting. A shorter wire bundle section, approximately 20 inches in length, exhibited the melted remnants of the cannon plug on one end. The remaining conductors were melted together in several places and there was beading on the conductor ends, as well as some signs of narrowing and necking on a few of the individual conductor ends. The wire bundle was shipped to the airframe manufacturer for examination and identification. Based on the number of wires, size of wires, and the location of the bundle found within the wreckage, the bundle was identified as an engine harness. The bundle was x-rayed and it was determined that no single-conductor cables were present within the bundle, consistent with the configuration of an engine full authority digital engine control (FADEC) harness. The flight crew did not report any engine controllability issues during the accident flight. Two additional cannon plugs with a few attached wire conductors were recovered. The plug with the longer section of wiring exhibited signs of narrowing and necking of the conductor ends. On the cannon plug with the shorter section of connected wiring, the conductor ends exhibited one large bead as well as melt/weld on the ends of several other remaining conductors. Seven sections of battery cable conductors were examined. All fractured conductor ends were consistent with mechanical (cold) fracture. There was no sign of electrical arcing on any of the conductor surfaces. The horizontal stabilizer trim actuator (HSTA) was removed from the airplane after its recovery from the accident site, and sent to the airframe manufacturer for examination. The rod end tube was secured to preserve the “as found” length of the HSTA. External wiring was present and appeared continuous, with no missing or melted insulation noted. However, the strain relief tabs on all the Electronic Control Module (ECM) wire connections were broken. Besides the broken strain relief tabs and heavy soot layer, the HSTA appeared to be in good condition with no obvious signs of mechanical damage. The spherical rod end bearings were both noted to be freely rotating. The actuator length from the center of one mounting bolt hole to the other was measured to be 36 1/16 inches. This corresponded to a horizontal tail position of approximately -1° (inconsistent with the HSTA exhibiting a runaway condition to the fully-extended position). The HSTA was fixed to a control box on “Side A” of the HSTA and a power supply. With power applied, the control box was directed to “extend” and the HSTA operated smoothly in the extend direction until the extend mechanical stop was reached. The control box was directed to “retract,” and the HSTA operated smoothly in the retract direction until the retract mechanical stop was reached. The control box was swapped to “Side B” and the test was replicated and yielded similar results. The HSTA was placed in a load frame and was positioned at the “as found” length of approximately 36 1/16 inches pin to pin. A load of 9,100 lbs was chosen because it closely aligned with the HSTA maximum operating load of 9,090 lbs. A 9,100 lb compressive load was applied using a ramp rate of 1,000 lbs/sec. No backdriving was observed or recorded during the load application. In order to verify the proper function of the mechanical no-back devices within the HSTA, both motor clutches were powered open. Powering open the motor clutches effectively disengages the motor stacks from the geartrain, eliminating the ability for the motor brakes to hold any backdriving torque resulting from the applied load. In this simulated failure scenario, the mechanical no-back devices acting on the ballscrew assembly (one device for compressive loads, another for tensile loads) are the only thing preventing the actuator from backdriving when axial load is applied. The same load regime was applied as described above, with nearly identical results: no backdriving was observed or recorded. The no-backs were therefore determined to be fully functional. During a post-recovery examination of the wiring system, the No. 2 battery, located on the right side of the airplane near the baggage compartment, displayed arcing signatures and a large hole in the inboard portion of the outer steel casing. The battery cells were contained within the steel casing, but had visible thermal damage. The over temperature sensor was separated and destroyed by fire. The No. 2 battery and the suspected battery cables were retained for further examination. The No. 2 battery, a nickel-cadmium type, was examined by the NTSB materials laboratory. All pieces were covered in a thin layer of soot and a layer of surface rust, consistent with exposure to a postaccident fire. As received, the remaining polymer cell cases were solidified into a single mass, and a layer of black ash “crust” covered many of the terminals at the top of the battery. After external examination, the battery was cut open in the vicinity of the arcing and hole for further inspection. Damage was noted on both the inside surface of the casing and to cell 7, which was directly adjacent to the damaged section. The interior surface of the casing near the through-holes was rough, with nodules of metal buildup scattered around the damaged area. The damage started approximately 3 5/16 inches above the bottom wall and was 3 1/8 inches tall by 2 13/16 inches wide. This roughly corresponded to the size and location of the thermal damage on the outer surface. Microscopic examination of the damaged area showed clusters of roughly spherical nodules of metal, consistent with melting and re-solidification spatter caused by electrical arcing. The polymer casings on visible cells were largely missing, and electrode plates were exposed. A mass of solidified polymer was present at the bottom of the battery. Cell 7 showed damage consistent with exposure to high heat. The polymer casing was melted, and portions of the exposed electrodes plates were eroded away. The area of damage started approximately 3 inches above the bottom wall and extended 3 1/4 inches upwards. At its widest point, the damage extended across the full width of the electrode plate. This dimension roughly corresponded to the damaged areas seen on the inside and outside of the case. The erosion penetrated through approximately seven electrodes, with the deepest areas roughly corresponding to the location of the two aforementioned through-holes on the case. The perforated steel mesh substrates of several electrodes were visible around the edges of the eroded area. Further, the edges of the eroded layers showed damage consistent with melting and re-solidification, and a buildup of metal nodules consistent with spatter from electrical arcing was present in these regions. Further cells adjacent to face 3 showed exposed electrodes, but none were eroded. The examination of the battery did not reveal evidence of any preimpact abnormalities.