ERA19LA072

An in-flight loss of control following an encounter with supercooled large droplet icing conditions, which ultimately resulted in an uncontrolled descent and subsequent inflight breakup. Also causal was the pilot’s failure to maintain an appropriate airspeed for flight in icing conditions.

The pilot and his pilot-rated passenger departed under instrument flight rules for a personal cross-country flight and climbed first to 19,000 ft mean sea level (msl), then to 23,000 ft msl several minutes later. Weather and air traffic control radar information indicated that the airplane had been operating in an area of heavy precipitation for about 20 minutes before it entered a descending right turn. A section of the right wing separated as the airplane descended, and the airplane impacted water. Performance calculations revealed that, shortly before the airplane began its descent, its airspeed decreased to between 77 and 90 knots indicated (KIAS). The airplane’s published stall speed at maximum gross weight was 69 KIAS. Because accumulation of ice on the unprotected areas of the airframe increased the airplane’s stall speed, with aerodynamic buffeting occurring up to 19 knots above the normal stall speed, the manufacturer stated that flight in icing conditions should be conducted at a speed not lower than 130 KIAS. Postaccident examination of the airplane’s flight controls and icing protection system components revealed no evidence of preimpact failure or malfunction. While the filament of the stall warning fail and windshield heat fail warning light bulbs were broken and stretched, potentially consistent with illumination at impact, it could not be determined at what portion of the flight the bulb(s) might have illuminated. Based on the environmental conditions at the time, the stall warning system was likely not accurate or reliable. Although fatigue cracks were noted in the right wing spar web, the cracks did not contribute to the inflight break-up. The pilot received a preflight weather briefing that included a convective SIGMET outlook and an AIRMET for icing between about 14,000 ft and 27,000 ft msl, valid for the area of the accident site about the time of the accident. Although the pilot did not access specific icing forecasts, which likely understated the potential for icing conditions, there was sufficient information available to the pilot to indicate possible icing at his chosen cruise altitude. The airplane was operating above the freezing level near the top of a mature cumulus cloud formation, which is known to have higher liquid water content, and within a heavy rain shower band. Those conditions likely resulted in an encounter with supercooled large droplet icing conditions, which exceeded the capability of the airplane’s icing protection system. Given this information, it is likely that the flight encountered icing conditions, which resulted in the airframe accumulating ice in excess of that able to be shed by the airplane’s icing protection system. It is also likely that, because the pilot was operating the airplane below the minimum icing airspeed, it encountered an aerodynamic stall at an airspeed that was higher than normal, which resulted in a loss of control, an uncontrolled descent, and subsequent inflight break-up.

FAA records indicated that, on January 2, 2018, the pilot received notice of disapproval of application for the flight portion of his instrument rating practical test. The FAA designated examiner indicated that he failed to maintain altitude within 100 ft during level flight and failed to maintain airspeed within 10 knots during a precision approach. On June 20, 2018, the pilot completed the practical test, which comprised 1.8 hours flight in a Piper PA-32-301 airplane. At that time, he reported 289 total hours of flight experience, of which 70 were instrument. The pilot obtained ground and flight training in the accident airplane between December 10, 2018, and December 14, 2018, to meet an insurance requirement of 25 hours of dual flight instruction. The training was performed by a flight instructor who was an FAA designated pilot examiner (DPE). During that training, the pilot received endorsements for a flight review and instrument proficiency. A review of the training records revealed that the pilot received training in operating within icing conditions and the use of the airplane’s ice protection system. According to notes from the flight instructor/DPE, the accident pilot needed to review the missed approach procedure and work on his instrument flight rules phraseology, but was “good” with autopilot use. The instructor stated that, as part of his ground training, he discussed cumuliform and cumulonimbus cloud types, the amount of precipitation that would occur in each, and the icing hazards associated with each. He also discussed and trained what airspeed (130 to 140 knots) to maintain during climb and cruise when operating in icing conditions to avoid ice accumulating on the bottom of wing. According to the report submitted by the pilot’s attorney, at the time of the accident, the pilot had a total flight experience of 390 hours, of which 290 were as pilot-in-command, and his total time in the accident airplane make and model was 30 hours. The airplane was equipped with an enhanced digital display indicator, annunciator panel, a portable Garmin 496 GPS map receiver, vertical profile weather radar system, and stormscope. The ice protection system was designed and tested to allow for continuous maximum and intermittent maximum icing specified in 14 CFR Part 25, Appendix C. The system comprised pneumatic wing and empennage boots, a wing ice detection light, electrothermal propeller deice pads installed on each propeller blade, an electrically heated windshield, heated lift detector, heated pitot head, two operating alternators, two vacuum pumps and the alternate static source. The surface deice system was manually engaged by a switch on the environmental/deice switch panel, which provided power to the solid-state timer. According to the manufacturer of the solid-state timer, it was not possible to determine positional information such as where it was in a cycle, nor did it contain any non-volatile memory. Although the enhanced digital display indicator was retained, no attempt was made to download any nonvolatile memory. The environmental/deice switch panel was examined by NTSB Materials Laboratory personnel, but the internal mechanisms of the switch panel were too damaged to determine switch position at impact. Forensic toxicology on specimens of the pilot was performed by the FAA Forensic Sciences Laboratory. The report indicated that unquantified amounts of N-Butanol and Propanol were detected in the liver and muscle specimen, while 62 mg/dL and 59 mg/dL were detected in the liver and muscle specimens, respectively. Ethanol is water soluble, and after absorption it quickly and uniformly distributes throughout the body’s tissues and fluids. The distribution pattern parallels water content and blood supply of the tissue. A small amount of ethanol can be produced after death by microbial activity, sometimes in conjunction with other alcohols, such as isopropanol, butanol, and n-propanol. A second search located the wreckage on February 5, 2019, about 525 ft north-northwest of the last radar target with altitude. The wreckage was recovered and transported to a salvage facility for examination. Examination of the airframe revealed that the fuselage crown in the cabin area near the aft seats was displaced down. The aft fuselage was separated at fuselage station (FS) 280. The following components were separated from the airplane: firewall with attached engine mount and nose landing gear; engine assembly with attached propeller; horizontal stabilizer; elevators; vertical stabilizer; rudder; section of right wing; outboard 1.5-foot section of left wing; left aileron; left flap; majority of right aileron; nose baggage door; and fuselage upper crown over the cockpit and cabin area. Components that were not recovered consisted of: emergency fuel pump, gascolator, baggage door, right wing inboard stall strip, inboard area of the right aileron, horizontal stabilizer; right elevator and section of left elevator; vertical stabilizer; and rudder. The nose landing gear was extended. The firewall was impact damaged, but there was no evidence of fire on it or any observed components. Both wings exhibited extensive impact damage, spar fractures, and deformation. The aft spar of the left wing was fractured consistent with overstress and displaced down at WS 128, while the spars of the right wing were fractured between WS 83 and 93. The fracture surfaces of the right-wing spars were excised and retained for further examination by NTSB. Examination of the aileron flight control system and flap system for both wings revealed no evidence of preimpact failure or malfunction. The speedbrakes of the left and right wings were fully and partially deployed, respectively. The landing gear were extended, and the radar pod of the right wing was separated. The blade of the lift transducer was restricted due to impacted sand. Both cabin doors were present. Both primary static ports were clear of obstruction on the exterior, while the alternate static port and static port of the pressurization outflow valve were blocked on the exterior consistent with sand found in the fuselage. Examination of the aft fuselage revealed that the forward attach point of the vertical stabilizer was fractured/pulled up. The vertical stabilizer aft spar remained attached and was bent in multiple directions with a portion of rudder attach bolt attached. The left and right main spars of the horizontal stabilizer were fractured in the down direction. Excised sections of the structural pieces of the vertical stabilizer, and horizontal stabilizers were retained for examination by the NTSB. Examination of the elevator and rudder flight control system revealed no evidence of preimpact failure or malfunction. The elevator trim barrel assembly was extended 1.0 inch, consistent with a neutral elevator trim setting. Examination of the pneumatic deice pressure control valves and pressure switches revealed that all the deice valves appeared to be in an unpowered state. The solenoid of the upper deice pressure control valve (A416) was separated and not located. All wires related to the deice control valves were cut consistent with recovery forward of the P402 connector, including the pressure switch wiring. Due to the condition of the control valves and pressure switches caused by the prolonged saltwater submersion, they were not tested. Examination of tail deice system components revealed that sections of deice boot remained attached to the left and right inflation and deflation hoses for the horizontal stabilizer. The vertical stabilizer deice “T” fitting was fractured, but hoses remained attached at both sides of the “T” fitting. Operational testing of each wing pneumatic deice boot was performed by using compressed air blown into the respective line that inflated the upper and lower chambers of each wing deice boot. Aside from impact-damaged areas o