N2442A Van Nierop RISEN 915IS · 2023

On July 28, 2023, about 1410 mountain daylight time, an experimental amateur-built Risen 915 iS, N2442, was destroyed when it was involved in an accident near Glendive, Montana. The pilot was fatally injured. The airplane was operated as a Title 14 Code of Federal Regulations Part 91 personal flight.

The pilot was conducting a personal cross-country flight that originated in Georgia eight days before the accident and had proceeded to Felts Field (SFF), Spokane, Washington. The day before the accident, the pilot departed SFF and flew to the Dawson Community Airport (GDV), Glendive, Montana. The pilot initially had planned to attend the EAA AirVenture Oshkosh Air Show in Oshkosh, Wisconsin. However, the pilot mentioned to a witness that he was abandoning his attempt to make the airshow due to operational delays. The pilot’s destination when he departed GDV is unknown.

The GDV airport manager reported that he assisted the pilot with fueling the airplane with about 15 gallons of fuel, and he mentioned that the pilot was in good spirits. He expected that the pilot intended to do a fly-by the airport after takeoff to demonstrate the speed of the airplane, which had a cruise speed just under 200 knots. The airport manager reported that he observed the airplane depart runway 12 at 1408. . The engine sounded like it was at full power during takeoff; however, the airplane appeared to climb slower than he expected to about 300-400 ft agl. The main landing gear were up but the nose gear appeared to be partially extended and had not completely retracted. He watched the airplane maintain a low altitude at a low airspeed until he observed a “wing drop,” and the airplane subsequently made two turns in a “flat spin” with about a 45-60° nose-down attitude as it descended behind terrain.

A second witness, who was located near the accident site, reported observing the airplane flying low and slow in a turn at the time of the accident; it appeared that the airplane was returning to the airport. The witness stated that he heard a loud pop and then observed a wing drop before the airplane entered a rotating nose-low attitude until it disappeared behind terrain. Subsequently, he observed smoke originating from the area where he had lost sight of the airplane.

The pilot's logbook was not recovered. This pilot reported civil flight experience that included 492 total and 15 hours in last six months as of exam dated 5/17/2023.

The pilot’s long cross-country journey was delayed due to maintenance issues with the airplane that presented themselves as the flight progressed. The pilot reported to the airplane kit manufacturer via text messages that the airplane was having an overheating issue during higher operating air temperatures on the ground, during taxi, and climbing after takeoff. The pilot had been diagnosing performance issues and made modifications to the cooling system. He texted the owner of a similar make and model airplane before the accident flight and indicated that the modifications had worked. No other information was available in regard to the engine overheating. There was no mention of overheating in the squawks or airplane service records. A flight logbook was not recovered.

The pilot’s long cross-country journey was delayed due to maintenance issues with the airplane that presented themselves as the flight progressed. The pilot reported to the airplane kit manufacturer via text messages that the airplane was having an overheating issue during higher operating air temperatures on the ground, during taxi, and climbing after takeoff. The pilot had been diagnosing performance issues and made modifications to the cooling system. He texted the owner of a similar make and model airplane before the accident flight and indicated that the modifications had worked. No other information was available in regard to the engine overheating. There was no mention of overheating in the squawks or airplane service records. A flight logbook was not recovered.

Examination of the accident site revealed that the airplane impacted terrain about 1.5 miles southeast of GDV in a farmer’s field. The airplane was mostly consumed by a postimpact fire. All major structural components of the airplane were accounted for at the accident site, including the fuel tanks, fuel system, and flight controls; no debris path was observed. Examination of these components revealed no anomalies. The main and nosewheel landing gear were mostly consumed by fire and not able to be functionally tested. Windshield fragments were found 30 ft forward of the nose. Carbon fiber airplane pieces were scattered aft of the tail for about 30 ft. The debris path was consistent with a steep impact angle. The emergency parachute had been discharged, but the chute had not unfurled. No track data of the accident flight was available. Flight control continuity was established and traced from the control inputs in the flight deck to their respective control surfaces in the wings and tail.

Postaccident examination of the engine revealed it had sustained significant damage from the impact and the postimpact fire. Each of the cylinders passed the manufacturer’s hardness test to determine if the engine had experienced operational overheating. The examination found no evidence of overheating nor were any preimpact mechanical anomalies identified in the engine or accessories that would have precluded normal operation.

An autopsy of the pilot was performed by the Department of Justice Forensic Science Division Montana State Chief Medical Examiner, Billings, Montana, which listed the cause of death as “blunt force injuries.” Toxicology testing performed by the FAA Forensic Sciences Laboratory did not identify any substances that are generally considered impairing.

On July 29, 2011, about 1247 eastern daylight time, a special light sport airplane (SLSA) TL Ultralight sro TL 2000 Sting S3, N2442, registered to N2442 Aviation, LLC, operated by Universal Flight Training, LLC, descended uncontrolled and crashed into trees then the ground about 13 nautical miles southeast of the Sarasota/Bradenton International Airport (SRQ), Sarasota, Florida. Visual meteorological conditions prevailed at the time and no flight plan was filed for the 14 Code of Federal Regulations (CFR) Part 91 demonstration local flight from the SRQ airport. The airplane sustained substantial damage, and the certified flight instructor (CFI) sustained serious injuries while the pilot-rated student sustained fatal injuries. The flight originated from SRQ about 1230.

The purpose of the flight was demonstration of the airplane to the pilot-rated student. The pilot-in-command (PIC) seated in the right seat advised the Federal Aviation Administration (FAA) inspector-in-charge (IIC) when interviewed while hospitalized the day after the accident that they were practicing stalls and he did not recall how or why the airplane entered a spin. The FAA-IIC reported that the PIC was unable to recover from a spin and the airplane continued in a spin until contacting a tree then the ground.

The CFI was able to exit the airplane; however, due to his injuries, waited on the ground next to the airplane for rescue personnel. While trapped in the wreckage the pilot-rated student called 911 and advised the dispatcher of their last known position, and that they had crashed through a tree canopy. During the conversation he advised the dispatcher that they were flying between 2,300 and 2,400 feet, and, “we were practicing stalls and it went into a spin and we were talking about how the plane was spin proof its not the instructor couldn’t pull it out of the spin….” The dispatcher remained on the call while rescue crews were en route to the crash site and continued to talk with the left seat occupant.

Rescue teams arrived and the left seat occupant was extricated from the airplane. Both occupants were airlifted by helicopter to a hospital in St. Petersburg, Florida.

The airplane was equipped with an airframe parachute recovery system; however, it was not deployed. The FAA-IIC asked the PIC why the aircraft’s ballistic parachute system was not armed or activated, and he responded that he does not arm the system on flights that are below 3,000 feet mean sea level (msl). He later stated that he was not trained in the use of the airplane parachute system and that is why he did not remove the safety pin.

The PIC seated in the right seat, age 56, holds a commercial pilot certificate with airplane multi-engine land, airplane single engine land, and instrument airplane ratings. He also holds a flight instructor certificate with airplane single engine issued August 13, 2010. He was issued a third class medical certificate with a limitation to wear corrective lenses on October 26, 2010. On the application for his last medical certificate he listed a total time of 1,100 hours. He estimated that at the time of the accident he had 1,200 hours total time, and 10 hours make and model, all as PIC.

The left seat occupant, age 71, held a private pilot certificate with airplane single engine land rating. He was last issued a third class medical certificate with a limitation to wear lenses for near and distant vision on December 15, 1989. On the application for his last medical certificate he reported having 415 hours total time, and 12 hours in the last 6 months. He also reported weighing 275 pounds.

The left seat occupant also completed an application with the operator on July 23, 2011, indicated his total time as pilot-in-command was 1,600 hours, and he weighed 275 pounds.

While on the phone with the 911 dispatcher awaiting rescue, the left seat occupant stated that he had not flown in 16 years and the flight was a refresher flight for him.

The airplane was manufactured as a Light Sport Aircraft in 2008 by TL Ultralight, sro as TL 2000 model Sting S3, and designated serial number TLUSA174. It met the standard specification Design and Performance established by ASTM document F2245, but was not required to comply with FAA Part 23 certification processes. It was powered by a 100 horsepower Rotax 912ULS engine and equipped with a 3-bladed ground adjustable Woodcomp propeller. It was also equipped with a TruTrak Flight Systems electronic flight information system (EFIS), and a I-K Technologies AIM-Sport Engine Monitor; neither of which record and retain flight or engine data. The instrument panel was also equipped with a panel dock for a portable GPS receiver.

Review of ASTM F2245-04, revealed section 4.5.9.1 pertaining to spins which indicates that for airplanes placarded “no intentional spins”, the airplane must be able to recover from a one-turn spin or a 3 second spin, whichever takes longer, in not more than one additional turn, with the controls used in the manner normally used for recovery. That condition is with flaps retracted and flaps extended, the applicable airspeed limit and limit maneuvering load factor may not be exceeded. The section also indicates that it must be impossible to obtain uncontrollable spins with any use of the controls.

The airplane was equipped with a Galaxy GRS ballistic parachute rescue system which is activated by a red “T” handle installed behind the co-pilot’s seat. According to data provided by the manufacturer, the parachute system design is purposefully constructed for the fastest possible opening.

Review of the maintenance records revealed the airplane was last inspected in accordance with an annual inspection on December 10, 2010. The airplane total time at that time was recorded to be 178.8 hours, while the hour meter reading at the time of the accident was 247.1, or an elapsed time of 68.3 hours since the annual inspection had been signed off as being completed.

A surface observation weather report taken at SRQ at 1253, or approximately 6 minutes after the accident indicates the wind was from 080 degrees at 7 knots, the visibility was 10 statute miles, and scattered clouds existed at 4,000 feet. The temperature and dew point were 33 and 23 degrees Celsius, respectively, and the altimeter setting was 30.08 inches of Mercury. The accident site was located approximately 13 nautical miles and 119 degrees from SRQ.

The airplane was equipped with a GPS; however, it was not located in the wreckage when it was examined by NTSB and a representative of the U.S. Field Technical Director following recovery.

Examination of the accident site by an FAA airworthiness inspector revealed the airplane came to rest at the base of a large oak tree and was obscured from view by low branches. The accident site was located at 27 degrees 17.43 minutes North latitude and 082 degrees 20.55 minutes West longitude. The FAA-IIC reported that the first responders removed the left wing during the extrication process of the left seat occupant, while the right wing remained attached. All components necessary to sustain flight remained attached or were found in close proximity to the main wreckage. No fire was observed on any component. The right wing was cut to facilitate recovery of the airplane.

Examination of the airplane and engine following recovery was performed with Safety Board oversight by a representative of the U.S. Field Technical Director of the airplane manufacturer and a representative of the engine manufacturer. The examination of the airframe revealed the fuselage was fractured circumferentially approximately 12 to 18 inches aft of the firewall. The rudder remained attached to the vertical stabilizer which also remained attached; however, evidence of overtravel of the rudder to the right was noted. Further inspection of the lower portion of the rudder revealed the rudder shaft was fractured. Rudder control cable continuity was noted between the rudder torque tube and the rear bellcrank near the control surface. The fractured rudder shaft was retained for further examination. The horizontal stabilizer remained attached; however, both sides were fractured about 30 inches from the fuselage centerline. The elevator remained connected by the anti-servo tab push/pull rod, and the anti-servo tab remained attached to the elevator at all hinge locations. The left side of the elevator was fractured in 2 pieces, while the right side of the elevator was full span. Inspection of the aileron and elevator flight control systems revealed no evidence of preimpact failure or malfunction.

Examination of the cockpit revealed the pilot’s control stick was bent forward, and was approximately 6.25 inches forward of the position of the right stick; however, both control sticks remained interconnected. The flap selector was in the full down position. The fuel shutoff valve was open, and the throttle was full forward with control cable continuity confirmed. The auxiliary fuel pump switch was separated from the instrument panel; however, electrical power was applied directly to the switch and it was found to operate satisfactory. Examination of the pilot’s (left seat) restraint system revealed the outboard lapbelt remained attached to structure which was structurally separated, while the inboard lapbelt and shoulder harness remained attached to the structure. Testing of the pilot’s shoulder harness inertia reel by hand revealed it tested satisfactory. Examination of the pilot’s seat revealed the seat back was pulled out, and the bottom side of the lower seat pan exhibited impact mark approximately 5.6 inches aft of the seat base screws associated with contact by the flight control tube. The instrument panel contained a panel dock for a GPS receiver; however, the receiver was not located.

Inspection of the co-pilot’s seat revealed no obvious seat frame pull-out. The inboard and outboard portion of the lapbelts in addition to the shoulder harness remained attached to structure. The co-pilot’s shoulder harness tested satisfactory when tested by hand. Examination of the bottom side of the lower seat pan exhibited impact mark approximately 5.6 inches aft of the seat base screws associated with contact by the flight control tube.

Examination of the left wing revealed it was fragmented and in multiple pieces, though the flap and aileron remained attached. The aileron push/pull rod was fractured in bending overload about 2/3 span, but remained connected to the bellcrank near the control surface. The flap torque tube indicated the flap was fully extended, which agreed with the position of the flap selector in the cockpit. No obstruction of the fuel supply from the wing root to the fuel strainer was noted.

Examination of the right wing revealed the aileron and flap remained attached, although the aileron was delaminated full span at the trailing edge. The flap torque tube indicated the flap was fully extended, which agreed with the position of the flap selector in the cockpit. A tree limb had penetrated the lower wing skin near the pitot static port, and evidence of a tree contact was noted at the inboard portion of the aileron. The fuel vent and fuel supply system were free of obstructions.

Examination of the rudder shaft was performed by the NTSB Materials Laboratory located in Washington, D.C. The results of the examination revealed the tube on the shaft consisted of two pieces, one long and one short, which could be rotated. The tip of the longer portion of the tube, adjacent to the circular flange, had been locally deformed and was bent. Examination of the fractures on the shaft revealed grainy surfaces on slant planes, consistent with bending overload, no evidence of preexisting cracks was noted.

Examination of the engine revealed the propeller remained attached to the engine and the engine remained attached to the airframe. One propeller blade remained connected to the propeller hub while the other 2 blades were fractured. Impact damage was noted to the carburetor sockets, air filters, oil tank, and coolant lines. The engine was removed from the airframe for further inspection which revealed continuity and compression on all cylinders. Inspection of the ignition system, carburetors, oil system components, cooling system, exhaust system, and auxiliary fuel pump which was operationally tested revealed no evidence of preimpact failure or malfunction. Inspection of the fuel strainer revealed the remains of fuel and some debris; however, the fuel screen was not blocked. Rotation of the engine using the starter revealed the engine-driven fuel pump would not pick up fluid from the source. The pump was removed from the engine and actuated by hand which produced the same results. The pump was retained for further examination.

Inspection of the propeller revealed the blade that remained connected to the propeller hub exhibited delamination at the blade tip, while the remaining 2 blades were fractured at the propeller hub. Inspection of the separated blades revealed minimal damage to the leading edges of both blades. The blade angles of all 3 blades were at 22 degrees (lowest blade angle is 16 degrees while the maximum blade angle is 26 degrees).

The certified flight instructor and pilot-rated student seated in the left seat were rescued and airlifted to a hospital in St. Petersburg, Florida, for treatment of their injuries. The left seat occupant expired at 0034 hours on August 1, 2011.

A postmortem examination was not performed of the left seat occupant; however, an external examination was performed. According to the external examination report, he weighed 340 pounds, and was identified to be “overly-nourished.” The cause of death was listed as complications of blunt trauma, while contributory conditions were Arteriosclerotic Cardiovascular Disease and Diabetes Mellitus. The report also indicates that a bandage and cast are present on the lower portion of the left leg and foot.

Blood specimens of the left seat occupant taken upon admittance to the hospital were submitted to the FAA Bioaeronautical Sciences Research Laboratory, Oklahoma City, Oklahoma, and also to the Pinellas County Forensic Laboratory, Largo, Florida. The toxicology report by FAA stated testing for carbon monoxide, cyanide, volatiles, and listed drugs could not be performed due to the insufficient quantity of blood submitted. The results of testing by Pinellas County Forensic Laboratory were negative for volatiles, drugs of abuse, and other tested drugs.

The airplane restraint system is designed to comply with ASTM standard 2245, Chapter 5.10, titled “Emergency Landing Conditions” which indicates that the structure must be designed to protect each occupant during emergency landing conditions when occupants (through seat belts or harnesses or both), experience the static inertia loads corresponding to 3 G’s up, 9 G’s forward, and 1.5 G’s lateral ultimate load factors. Section 5.1.2.2 of the same standard also indicates that special ultimate load factor of 2.0 shall be applied to seat belt/harness fittings including the seat if the seat belt or harness is attached to it.

Calculations by the airplane manufacturer were performed in an effort to determine the design load amount for each lapbelt attach point for the forward G loading limit of 9 G’s, and was based on the maximum seat limit of 250 pounds, or 113.4 kilograms (kg’s). The formula specified 60 percent of the forward load limit distributed to the lapbelt while the remaining 40 percent of the forward load limit distributed to the shoulder harness. The formula used for the calculations was:

113.4 kg (9 G’s) (2) (.6) = 612.4 kg

2 (Number of lapbelt attach points)

Using that same formula and the left seat occupant’s actual weight in kg’s (154), at the maximum 9 G forward design limit, the G loading of each lapbelt attach point of his seat was calculated to be approximately 832 kg, which was approximately 1.3 times the ultimate design load factor limit of each lapbelt attach point. No calculations were performed to determine the actual G loading at the moment of impact.

As previously mentioned, both the pilot and co-pilot bottom surface of the lower seat pans exhibited impact marks approximately 5.6 inches aft of the seat pan screws, which correlated with contact by each flight control tube. The airplane manufacturer was asked to correlate the position of the impact marks on the lower seat pan with the position of the elevator flight control position at impact and it was determined that the elevator control was approximately full nose-up to have caused the marks at the documented locations, although deformation of the seat pan was required to have the contact marks occur.

The PIC stated that he performed weight and balance calculations based on the provided weight of the passenger, and determined that the airplane’s weight was at the top of the envelope but within weight and balance.

Postaccident weight and balance calculations were performed using the empty weight of the airplane (822 pounds), the weight of the left seat occupant per the external examination report (340 pounds), the weight of the right seat occupant per his interview (185 pounds), and the total reported usable amount of fuel on-board at takeoff (9.5 gallons or 57 pounds). The calculations indicate that at the moment of engine start, the gross weight was 1,404 pounds and the center of gravity (CG) was calculated to be 83.25 inches aft of datum.

Postaccident weight and balance calculations were also performed using the empty weight of the airplane (822 pounds), the weight of the left seat occupant per his completed application form (275 pounds), the weight of the right seat occupant per his interview (185 pounds), and the total reported usable amount of fuel on-board at takeoff (9.5 gallons or 57 pounds). The calculations indicate that at the moment of engine start, the gross weight was 1,339 pounds and the center of gravity was calculated to be 82.91 inches aft of datum.

The Pilot Operating Handbook (POH) indicates that the maximum pilot or co-pilot seat load is 250 pounds, the maximum ramp weight is 1,340 pounds, the maximum takeoff weight is 1,320 pounds, and the forward and aft center of gravity limits are 80.2 inches and 86.7 inches aft of datum, respectively. The POH also indicates that it, “…is certified as a Light Sport Aircraft and is not approved for aerobatic flight including spins.” Section 3 of the POH contains information on the aircraft parachute system, and also a checklist of items to accomplish to deploy the aircraft parachute system. The same section also discusses the steps to accomplish to recover from an inadvertent spin, which indicates to bring the throttle to neutral, neutralize the ailerons, apply full rudder opposite the direction of rotation, and apply forward elevator to break the stall. Following ceased rotation, neutralize the rudder, apply aft elevator to recover from the nose-low attitude, and in the event the aircraft does not recover, deploy the parachute. The pre-taxi and before takeoff checklists found in Section 4 or the normal procedures section both indicate to either remove or verify removal of the airframe parachute GRS safety pin.

The airplane manufacturer performed spin testing at maximum gross weight in the most forward and aft CG positions. Spins with and without power were performed and with the flaps retracted and fully extended. The flight test report considered proprietary indicates that in the various configurations, airframe buffet proceeds a stall, the position of the flaps did not aggravate the spin process but care must be taken to avoid a flap overspeed during the descent, the airplane had a tendency to return to normal flight after about ½ turn, the ailerons must be kept neutral, and upon spin entry immediately push the control stick forward and apply opposite rudder input.

The engine-driven fuel pump was sent to Rotech Flight Safety, Inc., in Vernon, British Columbia for further testing with Safety Board oversight. The pump was installed on a new Rotax 912ULS engine which was installed on a test stand. The engine-driven fuel pump would not supply an adequate amount of fuel to sustain engine operation; the fuel pressure reading was 0.32 psi (normal psi range 2.2 to 5.8 psi). The auxiliary fuel pump was turned on and fuel was noted to pass internally through the pump indicating there was no blockage. Although there was no damage to the fuel pump, fuel leakage was noted at the base of the pump. The new engine was started using the auxiliary fuel pump and the engine was found to operate normally. The fuel pump was disassembled which revealed the inlet screen was clean and no debris was inside the filter, though the diaphragm was noted to be cracked on the air side. The pump was then sent to Austria Civil Aviation Safety Investigation Authority for examination at Rotax’s facility.

Examination of the engine-driven fuel pump at Rotax’s facility in Austria with Austria Civil Aviation Safety Investigation Authority revealed superficial cracks on the dry side of the diaphragm in the area of the diaphragm plate. An adapter was used to perform an airflow test of the diaphragm with connected push rod which revealed a slight pressure reduction. The internal components of the pump were reconstructed in a housing, and placed in a fixture which revealed fluid leakage at the weep holes; however, functionality in terms of fuel flow and fuel pressure was verified. The reconstructed fuel pump was installed on a new Rotax 912UL engine which was started and operated for 30 minutes without the electric fuel pump operating. The engine was noted to start normally and run smoothly with the system fuel pressure within limits over the full range of different rpm settings.