N8376KGrumman Acft Eng Cor-Schweizer G-164B · 1982

On May 27, 2021, about 0750 eastern daylight time, a Grumman G-164B airplane, N8376K, was substantially damaged when it was involved in an accident near Celina, Ohio. The pilot sustained serious injuries. The airplane was operated as a Title 14 Code of Federal Regulations Part 137 aerial application flight.

Earlier in the morning, the pilot repositioned the airplane from Defiance Memorial Airport (DFI), Defiance, Ohio, to Lakefield Airport (CQA), Celina, Ohio. Shortly after he departed DFI the pilot noticed that he forgot to bring his GPS and returned and landed at DFI. After retrieving the GPS, the pilot departed DFI a second time and flew to Charloe Airport (53OH), Paulding, Ohio, where he retrieved a hose connector before continuing to CQA. According to the pilot, there were no issues with the airplane or its engine during his preflight inspection or the repositioning flights.

The accident occurred during the first agricultural flight of the day. Before departure, the retailer who sold the fungicide and insecticide loaded the water/product solution on the airplane while the pilot fueled the airplane with 27 gallons of Jet-A aviation fuel. The pilot estimated the fuel load at departure was about 80 gallons, and the water/product solution was likely about 330 gallons in total. The pilot stated that the loaded airplane was below the airplane’s maximum gross weight, and that he had previously flown the airplane with similar loads without any performance issues.

The pilot back taxied on runway 8 at CQA and completed an uneventful engine runup before starting the takeoff roll. During the takeoff roll, the engine torque and propeller speed gauges indicated about 52 psi and 100%, respectively. The pilot stated that he did not perform a maximum-performance takeoff and a normal liftoff was achieved without any issues.

The pilot reported that during the initial climb, about 20-75 ft above ground level, the airplane had a sudden fluctuation in engine power from high-power to low-power and back to high-power. The pilot stated that he was “thrown forward” into his safety restraints and then “thrown back” into his seat. The forward/backward motion only occurred once and was “pretty quick and hard.” The pilot stated that after the loss of engine power the propeller speed remained at 100% but the airplane’s climb performance was significantly diminished. He did not cross-check any other engine gauges.

The pilot pitched the airplane for best glide airspeed and entered a slight right turn from the runway heading to avoid trees. Believing the engine was still running, the pilot did not want to jettison the airplane’s load unnecessarily into the river near the airport. The engine lost all power shortly after he entered the right turn. The pilot was wearing an active noise-canceling helmet, but he noticed there was a noticeable lack of engine noise, and that he only heard the wind passing outside the cabin. The pilot jettisoned the airplane’s load in an attempt to maintain a climb, but the airplane descended and impacted a levee, where it nosed over into the river.

The airplane came to rest inverted on the bank of a small river. The airplane sustained substantial damage to the vertical stabilizer, rudder, upper wings, and the fuselage.

Examination of the engine before it was removed from the airframe revealed that the airframe control linkage to the propeller pitch control (PPC) was disconnected from the splined end of the shouldered shaft of the PPC, and the retaining bolt was loose and was not secured by safety wire, as shown in figure 1. The PPC shouldered shaft did have an internally threaded hole, but there was no evidence that the secondary retention feature was installed at the time of the accident. The secondary retention feature is discussed in more detail in the Additional Information section.

Postaccident engine examination and disassembly revealed rotational scoring at multiple locations, including the compressor impeller blades, impeller shroud, and propeller shaft. Light metal spray was adhered to the 3rd stage turbine stator vanes. Surfaces throughout the engine gaspath were coated with earthen debris. The torsion shaft aft spline was fractured 1.25 inches from the aft end. The disassembly and examination of the engine revealed no evidence of a preimpact failure or other anomalies that would have prevented normal operation.

The engine fuel pump was unable to hold pressure during bench testing and was subsequently disassembled. The high-pressure gear stage carbon bushing was found fractured and, according to the engine manufacturer, is consistent with impact-related damage. Otherwise, the fuel pump examination was unremarkable.

The fuel control unit (FCU) and propeller governor were tested and examined at the manufacturer. The FCU Standard Day Acceleration Schedule had significantly high out-of-limits fuel flow discharge at all test points and FCU pressure readings did not respond normally to speed input changes. The Deceleration Schedule fuel flow discharge readings were low out-of-limits during testing. All other schedules were within the expected range for similar service run units that have undergone customer adjustments. The FCU was disassembled and there was no evidence of internal failure. All springs were intact, and all bearings and levers moved freely. The propeller governor was bench tested, and all test points were consistent with similar service run units.

The propeller remained attached to the engine propeller shaft flange, and all three blades remained attached to the propeller assembly. A counterweight puncture mark on the spinner dome was consistent with the propeller in the normal range of operation at about 24° blade angle at impact. There was no damage or evidence to indicate the propeller was feathered or in the beta/reverse range at the time of impact. All three blade shanks remained attached via the retention clamps and no pilot tubes were fractured. All three blades were fractured near the tip. The damage to the blades included chordwise/rotational scoring on both camber and face sides, bending opposite rotation, leading edge gouging with material deformation towards high pitch, and progressive compound bent/twisting, as shown in Figure 2.

The propeller examination and disassembly revealed no evidence of a preimpact failure or other anomalies that would have prevented normal operation. All observed damage was consistent with high impact forces. Blade damage and impact signatures on the cylinder and spinner dome indicated the propeller was rotating in the normal blade angle range of operation at moderate power.

Figure 1 – Airframe propeller pitch control (PPC) linkage disconnected from PPC splined shaft, and loose retaining bolt without safety wire (Honeywell Photo)

Figure 2 – Propeller (Hartzell Propeller Photo)

The engine was previously installed on a Mitsubishi MU-2B-20 airplane and was removed on November 12, 2014. On March 24, 2015, the engine was installed on the accident airplane in accordance with supplemental type certificate (STC) No. SA7987SW.

The airplane HOBBS meter indicated 1,382.4 hours at the accident site. The airplane had accumulated 14.1 hours since the last 100-hour and annual inspection completed on May 19, 2021, and May 20, 2021, respectively. The airframe and engine total time since new were 9,187.2 hours and 7,733 hours, respectively. The engine had accumulated 521.2 hours since the last overhaul completed on May 17, 2019.

On July 29, 1975, the Federal Aviation Administration (FAA) issued Airworthiness Directive (AD) 75-16-20 for several Mitsubishi MU-2 models requiring repetitive inspections of the propeller pitch control lever for security and proper rigging.

On December 21, 2011, the engine manufacturer released service bulletins (SB) TPE331-72-2190, TPE331-72-2191, and TPE331-72-72291 that recommended TPE331 operators inspect the PPC assembly to verify the splined end of the shouldered shaft had an internally threaded hole to accommodate a secondary retention feature for the aircraft linkage control interface. If the hole was not present or damaged, the SB included instructions to drill, tap, and countersink a threaded hole. The aircraft linkage control and secondary retention (bolt, washer, and lockwire) are considered airframe components.

On December 29, 2020, the FAA issued AD 2020-26-14 for the Mitsubishi MU-2 fleet that superseded AD 75-16-20 and required the installation of a secondary retention feature on the PPC lever linkage, repetitive inspections of the PPC lever linkage, and reporting inspection results to the FAA. However, because AD 2020-26-14 was only applicable to Mitsubishi MU-2 airplanes it did not include other airplane make/models that were modified by STCs with a TPE331 engine and a comparable PPC linkage like the accident airplane (Grumman G-164B AgCat).

In response to the accident, on May 17, 2023, the Federal Aviation Administration released Airworthiness Directive (AD) 2022-01801-A to mandate the installation of the PPC secondary retention feature on Allied Ag Cat Productions, Inc. model G-164A and G-164B airplanes to include those airframes that were equipped with a TPE331 engine installed under several identified STCs.

According to the Honeywell TPE331 overhaul manual:

The propeller pitch control (PPC) is mounted at the rear of the [engine’s] reduction gear section on the propeller shaft centerline. The PPC is composed of a ported sleeve, which is positioned by cam. The control end of the beta tube (which also has oil-supply ports) rides inside the ported sleeve.

The positioning cam-control shaft is connected to the main metering valve power-lever shaft by mechanical linkage. During propeller-governing mode, the propeller pitch control serves no basic function other than oil passage and housing for the beta tube. In beta-mode (under-speed governing) the propeller pitch control provides for operator control of propeller blade pitch angle. Operator control is accomplished by manually positioning the propeller pitch control cam. The beta tube oil supply holes are then aligned with the ported sleeve so that the pressure supplied to the propeller balances the propeller piston spring.

During takeoff, the propeller governor controls propeller pitch. The only function the PPC has in propeller governing mode is to act as an oil passageway. The PPC and FCU main metering valve are connected through the concentric shaft and are rigged to move together. The cockpit power lever is connected to the splined end of the shouldered shaft of the PPC in the Grumman G-164B (AgCat) installation. The cockpit condition lever is connected to the propeller governor. According to Honeywell Service Bulletin TPE331-76-2002, a disengagement of the airframe PPC linkage will result in the pilot’s inability to change fuel flow and/or propeller pitch (blade angle) during flight and landing.

The pilot stated that he was spraying a swath of field bordered by power lines. After making a pass, the pilot's attention was focused on a pole and he struck nearby power lines, which wrapped around the propeller. The pilot noticed a sudden loss in power and thought he could make it back to the airport. The power loss increased and the pilot initiated a forced landing into a field, resulting in a hard landing. The engine detached from the airplane and both wings and fuselage sustained substantial structural damage. The pilot reported no mechanical failures or malfunctions that would have precluded normal flight.