N307ABBarnett Allen S RV7 · 2011

On December 10, 2015, about 1347 mountain standard time, an experimental amateur built, RV-7 airplane, N307AB, experienced an in-flight break up and then impacted terrain about 3 miles west of General Dick Stout Field Airport, Hurricane, Utah. The airline transport pilot and passenger were fatally injured, and the airplane sustained substantial damage. The airplane was registered to and was being operated by the pilot as a Title 14 Code of Federal Regulations Part 91 personal flight. Visual meteorological conditions existed near the accident site about the time of the accident, and no flight plan had been filed. The local flight departed from an unknown airport at an undetermined time.

Several witnesses located near the accident site stated that they heard the airplane's engine and that it sounded like it was making power changes. The witnesses added that they saw airplane debris floating in the air. One witness stated that the engine was running during the entire descent and that he also observed the airplane spiraling and descending in a cork-screw type maneuver. Another witness reported seeing the airplane inverted at a low altitude just before impact.

The pilot, held an airline transport pilot certificate with airplane multi-engine land, single-engine land, instrument, and instructor single-engine land ratings. The pilot was issued a first-class Federal Aviation Administration airman medical certificate on October 22, 2015, with the limitation that he must have glasses available for near vision.The pilot reported on his most recent medical certificate application that he had accumulated 17,359 total flight hours, 403 of which were accumulated in the previous 180 days.

The two-seat, low-wing, fixed-gear airplane, was assembled in 2011, and it was issued an airworthiness certificate certified for aerobatic maneuvers in March 2011. It was powered by an experimental 180-horsepower ECI/Titan IO-360 reciprocating engine.The engine was equipped with a Whirlwind 200RV propeller. The last documented inspection was a conditional inspection on May 15, 2015, at an airframe time of 258.9 hours.

The airplane's kit manufacturer listed the maximum load factor as +6 g, the minimum load factor as -3 g, and the never exceed speed (VNE) as 230 mph in Section 15 of the RV-7/7A Construction Manual. Additionally, the kit manufacturer listed a maximum aerobatic gross weight of 1,600 lbs in Section 14 of the Construction Manual.

The 1355 recorded weather observation at Saint George Regional Airport, Saint George, Utah, located about 12 miles west-southwest from the accident site, reported calm wind, visibility 10 statute miles, clear skies, temperature 12° C, dew point -2° C, and an altimeter setting of 29.87 inches of mercury.

The accident site was located between a cold front to the northwest and a high-pressure area to the southwest, in an area of strong-pressure gradient. A model sounding, which included a wind profile, for the area over the accident site about the time of the accident estimated that the surface horizontal wind speed was 220° at 8 knots, with winds increasing in speed with height and veering to the west. The mean 0-to-18,000 ft mean sea level (msl) winds were from 250° at 52 knots. The model supported light-to-moderate clear air turbulence from 6,400 through 8,000 ft msl, and mountain wave development from 10,000 to 12,000 ft msl.

Pilot reports noted evidence of mountain wave activity in the region but with moderate-to-severe turbulence near the accident site, at 6,500 ft msl, consistent with the model sounding. An AIRMET for moderate turbulence below 18,000 ft was active over the accident site at the accident time. No SIGMET was active for the accident site at the accident time.

The two-seat, low-wing, fixed-gear airplane, was assembled in 2011, and it was issued an airworthiness certificate certified for aerobatic maneuvers in March 2011. It was powered by an experimental 180-horsepower ECI/Titan IO-360 reciprocating engine.The engine was equipped with a Whirlwind 200RV propeller. The last documented inspection was a conditional inspection on May 15, 2015, at an airframe time of 258.9 hours.

The airplane's kit manufacturer listed the maximum load factor as +6 g, the minimum load factor as -3 g, and the never exceed speed (VNE) as 230 mph in Section 15 of the RV-7/7A Construction Manual. Additionally, the kit manufacturer listed a maximum aerobatic gross weight of 1,600 lbs in Section 14 of the Construction Manual.

The airplane wreckage was located about 4.2 miles southwest of Hurricane, Utah, on flat sagebrush-covered terrain on top of a mesa. The debris path was about 1,460 ft long and 450 ft wide. All major components of the airplane were located in the debris path.

The main airplane wreckage was located almost at the northern extent of the debris field and included the fuselage, engine, right wing, half of the left wing, a majority of the left and right elevators, and the lower half of the rudder. The vertical stabilizer with the upper half of the rudder attached was located at the southern extent of the debris field, about 1,420 ft south-southwest of the main wreckage. The left and right horizontal stabilizers were located about 850 ft and 790 ft, respectively, south of the main wreckage. The left aileron was located about 430 ft south-southwest of the main wreckage, and the left outboard wing was located about 320 ft south-southwest of the main wreckage.

The main wreckage was found inverted. There were no noticeable ground scars leading up to the wreckage. The fuselage was intact, but the upper half was crushed. The canopy frame was separated from the airframe and located about 55 ft northeast of the main wreckage. Most of the acrylic canopy was fractured from the frame and found in many pieces in the debris field. The engine remained attached to the fuselage. One of the composite propeller blades was fractured from the hub and the other blade was missing the tip portion. Debris consistent with propeller material was found around the main wreckage. The examination of the engine revealed no evidence of mechanical malfunctions or failures that would have precluded normal operation. The main landing gear remained attached to the lower fuselage, and there was some deformation at the attachment points.

The entire right wing remained attached to the fuselage with the flap and aileron attached. The right flap was in the "up" position. The outboard half of the right wing was deformed downward about 15º to 20º at the flap/aileron junction, located about 57 inches outboard of the wing attachment point. The upper and lower wing skins were buckled around the area where the wing was deformed downward. The right fiberglass wingtip remained attached to the wing but was splayed open at the trailing edge.

The inboard half of the left wing remained attached to the fuselage with the flap attached. The left flap was in the "up" position. The outboard half of the left wing had separated at the flap/aileron junction located about 57 inches outboard of the wing attachment point. The main spar fractured at the location where the upper and lower spar caps undergo a net section decrease from inboard to outboard. The outboard half of the left wing was mostly intact with minimal damage noted.

The Utah Department of Health, Office of the Medical Examiner, conducted an autopsy on the pilot. The medical examiner determined that the cause of death was "blunt force trauma."

The Federal Aviation Administration's Bioaeronautical Sciences Research Laboratory, Oklahoma City, Oklahoma, performed toxicological testing of specimens from the pilot. Test results were negative for carbon monoxide, cyanide, and volatiles. The testing detected doxylamine in the liver but not in the blood and ibuprofen in the blood.

Doxylamine is an over-the-counter antihistamine medication that can be used in combination with decongestants and other medications to relieve sneezing, runny nose, and nasal congestion caused by the common cold. It can be sedating. Ibuprofen is used to reduce fever and to relieve minor aches and pains from headaches, muscle aches, arthritis, the common cold and other such ailments.

Structures Examination

The vertical stabilizer separated mostly intact from the airplane during the accident sequence and had the upper portion of the rudder attached. The forward spar was fractured just above its attachment to the horizontal stabilizer front spar. The rear spar was fractured and twisted just above the stabilizer shelf, consistent with the vertical stabilizer separating leading edge left. Most of the upper half of the rudder remained attached at the upper and center hinge points. The lower half remained attached to the empennage. The rudder trailing edge was splayed open, and the trailing edge strip remained attached to the left rudder skin on both halves. There was evidence of sealant between the rudder skins and trailing edge strip. The rudder was fractured spanwise just below the center hinge. The rudder counterweight was separated from the upper end of the rudder just above the upper skin stiffeners (see figure).

Figure. Vertical stabilizer and attached rudder.

The center portion of the horizontal stabilizer forward spar, most of the horizontal stabilizer rear spar, and most of the left and right elevators remained attached to the fuselage. The horizontal stabilizer forward and rear spars were deformed down and aft. The elevators were deformed down and aft matching the spar deformation. The left and right horizontal stabilizers were found in the debris field. There was buckling damage on the lower skin of both horizontal stabilizers consistent with the stabilizers separating downward.

Examination of the inboard and outboard left wing sections at the fracture location revealed that the fractures exhibited damage and deformation consistent with the separation of the outboard portion of the wing in a downward direction.

Control continuity was established from the cockpit controls to the elevators and the right aileron. The left aileron control rod aft rod end was fractured from the control rod. The rod end remained attached to the aileron control horn at the inboard aileron hinge. Control continuity was established from the control stick in the cockpit to the left aileron bell crank and aft to the fracture point on the left aileron control rod. The rudder cables were jammed somewhere in the fuselage, and control continuity could not be established, but the cables remained attached at the rudder and the pedals.

All the fractures exhibited a dull, grainy appearance consistent with overstress separation. There was no evidence of progressive or pre-existing fractures on any of the parts.

Flutter is an aeroelastic phenomenon that can occur when an airplane’s natural mode of structural vibration couples with the aerodynamic forces acting on the airplane to produce a rapid periodic motion, oscillation, or vibration. Flutter can be somewhat stable if the natural damping of the structure prevents an increase in the forces and motions. Flutter can become dynamically unstable if the damping is not adequate or speed is increased, resulting in increasing self-excited destructive forces being applied to the structure. Flutter can range from an annoying buzz of a flight control or aerodynamic surface to a violent destructive failure of the structure in a very short period of time. Due to the high frequency of oscillation, even when flutter is on the verge of becoming catastrophic, it can still be very hard to detect. Aircraft speed, structural stiffness, and mass distribution are three inputs that govern flutter. An increase in airspeed, a reduction in structural stiffness, or a change in mass distribution can increase the susceptibility to flutter.

Van’s contracted with an outside company to perform a flutter analysis for the RV-8 airplane. The analysis was completed in October 1998. The company performed a ground vibration test (GVT) on an RV-8 airplane at the Van’s factory in the zero fuel and full fuel configurations to establish the natural modes of vibration for the airplane. This information was used to perform a flutter analysis at a simulated altitude of 10,000 ft up to Vd of 256 mph or 220 knots. The results of the analysis showed the RV-8 airplane to be free from flutter above the design envelope with the control surface balance weights provided by Van’s.

Since the RV-7 airplane is a derivative of the RV-8 with a wider cabin, the results of the RV-8 flutter analysis were used for the RV-7. The RV-7 flutter analysis was completed by the same company in April 2001. A GVT was performed on the RV-7 airplane with zero fuel to evaluate the changes in vibration modes. The GVT showed that the wing vibration modes differed enough to require a flutter analysis, while the tail modes were essentially the same as for the RV-8. The analysis concluded that the RV-7 wing was free of flutter to speeds well above the dive speed of the airplane. The tail flutter analysis results from the RV-8 were applicable to the RV-7.

According to Van’s, the rudder used on the RV-7 airplanes was sized to meet the spin recovery requirements in the Part 23 regulations. The GVTs conducted for the RV-7 airplanes show the first flutter mode that manifests with increasing airspeed is a fuselage side bending vibration mode that couples with a rudder flutter mode. The rudder damping for this mode is dependent on the density of the air flowing over the rudder and thus is dependent on the true airspeed of the airplane. The tests accounted for changes in the rudder counterbalance mass. The tests showed the existing rudder counterbalance mass was sufficient to account for variations in paint and the addition of a taillight to the rudder. The GVTs and flutter analyses performed showed that the airplanes are free of flutter beyond the design never exceed speed. Van’s also showed the airplanes to be free of flutter up to the design demonstrated dive speed for each of the models. Van’s noted that the rudders analyzed and tested were built according to the design and did not incorporate any additions such as servo-controlled trim tabs.

Also, according to Van’s, the GVTs showed that at even higher airspeeds flutter modes involving wing symmetric bending, aileron rotation, and horizontal stabilizer bending and torsion manifest. The tests accounted for changes in the fuel load on the airplane. The analysis indicated that these flutter modes may interact with each other to exacerbate the response. Based on flight testing, GVT results, and flutter analyses, Van’s concluded that the airplanes meet the requirements in Part 23 regulations and are free from flutter as designed and built when operated within the prescribed flight envelope.

Electronic Devices

No flight data for the accident flight could be recovered from the electronic devices found in the wreckage. However, a GoPro Hero 4 camera, which had sustained significant impact damage, contained two files recorded on previous flights in which the accident airplane performed an aileron roll to the right. In 2022, the National Transportation Safety Board began using an updated file carving process to retrieve data from electronic devices. The updated file carving method produced a number of video fragments consistent with the files recovered in December 2015; however, the process also produced a number of JPEG files that were not recovered previously. Of the recovered JPEG files, 29 files were consistent with having been recorded on the accident flight. These files’ metadata suggested they were captured in a time lapsed still image mode, which is a capture mode that initiates recording of a JPEG sequence when the user presses the device’s shutter button. More than half the images in the timelapse mode captured the accident sequence.

Based on comparison of terrain features, buildings, and lakes between several photos and Google Earth, the airplane was initially heading southbound over a sparsely populated mesa southwest of Hurricane. In some of the photos, the pilot and passenger were looking toward the pilot’s lap. In some, they were looking towards the GoPro, and at times during the maneuver, the pilot was looking downward. The pilot appeared to roll the airplane right to an inverted position and appeared to pull back into a split-S maneuver that resulted in a northbound heading. As the airplane almost returned to wings level, it appeared that the vertical stabilizer was no longer with the aircraft. Between two of the later photos, the airplane went from an almost wings level attitude to inverted with an apparent substantial negative g load.

Weight and Balance

The distribution of the airplane contents throughout the debris field prevented an accurate weight and balance assessment and the airplane's most recent weight and balance records were not located. Therefore, an estimated weight and balance calculation was conducted. According to the airplane's kit manufacturer, the airplane had a maximum factory basic weight of 1,114 lbs and a useful load of 686 lbs. The medical examiner reported that the total weight of the occupants was 306 lbs. Assuming a total fuel load of 42 gallons, the airplane would have been about 72 lbs above its maximum aerobatic gross weight of 1,600 lbs at the time of the accident.