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On August 6, 2010, about 0900 Pacific daylight time, a PZL M-18, N70461, was substantially damaged during a forced landing following a complete loss of engine power shortly after takeoff from runway 16 at Sanborn Airport (38CN), Meridian, California. The pilot/owner received minor injuries. The flight was operated under the provisions of Title 14 Code of Federal Regulations (CFR) Part 137, doing business as Martin's Dusters. Visual meteorological conditions prevailed, and no flight plan was filed for the agricultural agent application flight.
According to the pilot, the flight was not the first flight of the day. The airplane was loaded with 3,600 pounds of fertilizer and about 120 gallons of fuel; the respective maximum capacities were 5,000 pounds and 150 gallons. The pilot stated that he took off to the south, that the airplane got airborne easily, and that he maintained a low altitude for his application work. Just after he began his turn to the east, he heard a loud "crack," and felt the airplane decelerate. He recognized that the engine had stopped developing power.
He leveled the wings and attempted to touch down in the water-filled rice paddy adjacent to, and south of, the airstrip. After touchdown, the airplane nosed over onto its back. The pilot was unable to extricate himself, and a nearby witness assisted him in exiting the airplane.
The wreckage was recovered to a secure storage facility a few days after the accident. On September 16, 2010, the wreckage was examined by personnel from the NTSB, the Federal Aviation Administration (FAA), and the engine manufacturer. Subsequent to that examination, the engine was removed from the airplane and shipped to the Honeywell Product Integrity Investigation Laboratory, Phoenix, Arizona. In early December 2010, the engine was examined in detail at that facility by representatives of Honeywell and the NTSB.
The pilot reported that he had a total flight experience of about 20,610 hours, including about 205 hours in the accident airplane make and model. He held an FAA commercial pilot certificate, with airplane single-engine land and rotorcraft-helicopter ratings. His most recent FAA second-class medical certificate was issued in January 2010, and his most recent flight review was also completed in January 2010.
FAA registration documentation indicated that the airplane was imported new into the United States in 1986. In March 1992, the ASZ62IR M-18 radial (reciprocating) engine and PZL-KALISZ propeller were removed, and a Lycoming T53-L-7A turboshaft engine and a Hamilton Standard 53C51 propeller were installed. That same T53-L-7A engine was the one installed on the airplane at the time of the accident.
The airplane was first registered to the accident operator in May 2009. However, the FAA documentation current at the time of the accident incorrectly specified that the airplane was equipped with a reciprocating engine, and the airworthiness category of the airplane was missing from the registration documentation.
The recorded weather observation at an airport 15 miles east of the accident location included winds of 4 knots from 180 degrees, clear skies, and a temperature of 18 degrees C about the time of the accident.
WRECKAGE AND IMPACT INFORMATION
The airstrip was surrounded by agricultural fields interlaced with a series of levees. After the power loss, the pilot deviated slightly to the east to avoid a levee during the forced landing. Ground scars indicated that the airplane touched down about 850 feet south of the south end of the runway, and came to rest within about 150 feet of the initial touchdown point. The fuselage fractured just aft of the firewall, so that the bulk of the airplane came to rest inverted. The engine section was folded back, with the nose facing aft. The fuselage and cockpit section remained relatively intact, but the canopy section was partially embedded in the soft mud and water, which prevented the pilot form readily exiting the airplane without assistance. The fixed, conventional landing gear remained intact. The vertical stabilizer was crushed and canted forward, and the wings and spray rig were damaged. The engine remained attached to the firewall, the propeller remained attached to the engine, and all three propeller blades were bent aft. There was no fire.
T53 Series Engine Background Information
The T53 series engines were originally designed and manufactured by Lycoming in the 1950s. In 1994, AlliedSignal acquired the Lycoming Turbine Engine Division of Textron. In 1999, the resulting company became part of Honeywell Aerospace. Honeywell Aerospace provided technical assistance for the investigation of this accident.
According to the Honeywell Aerospace representative, T53 engine models that contained an "L" in the variant designators were military engines. The T53-L-7A engine was a life-extension conversion of the previous T53-L-7 model. The "A" conversion extended the Time Between Inspection (TBI) to 600 hours, and the Time Between Overhaul (TBO) to 1,800 hours.
The Honeywell Aerospace engine examination report contained the following statement in its introduction: "The T53-L-7A engine was designed and manufactured to meet the requirements set forth by the United States Military; as such, the U.S. Military was the design holder and had sole responsibility for developing and maintaining the continuing airworthiness requirements for the T53-L-7A series engine. Honeywell had no knowledge of, nor provided input regarding the installation of the T53-L-7A series engine in the accident aircraft."
Engine Maintenance Guidance
Since the U.S. military was the design holder and had sole responsibility for developing and maintaining the continued airworthiness requirements for the T53-L-7A series engine, maintenance guidance was only available via U.S. military documentation. Department of the Army Technical Bulletin TB 55-2800-200-31/1 (T53 Inspection Guide) presented detailed information for the subject T53-L-7A engine and many of its components. The inspection guide referred to the "hot end," which included the combustion turbine assembly and the gas producer system. Colloquial industry terminology typically referred to this as the "hot section;" for the purposes of this report they are considered synonymous. Since the -7A variant was a "Long-Life" engine, the specified inspection interval for the hot end was 600 service hours, compared to 400 hours for other variants. The inspection guide provided specific textual and diagrammatic information regarding the wear and damage limitations for the fuel vaporizers. The cited limitations included characteristics and allowable dimensions for burn-off, warping, bulging, out-of-round, cracks, and holes.
FAA Approval of Aircraft Modifications
Modifications to FAA-certificated aircraft must be approved by one of two means; either a "field approval," or a Supplemental Type Certificate (STC). Any STC- or field-approved modification must comply with all applicable regulations. The FAA Flight Standards division was responsible for field approvals, while the FAA Aircraft Certification Office (ACO) branch was responsible for STC approvals. There were no explicit FAA requirements for an applicant to involve either the engine or airframe manufacturer in a field- or STC-approvals; the only requirement was that the application for approval of the modification must contain all the relevant data required for certification. However, approval of a military version of an engine that did not have a civil type certificate (TC) would require a much greater certification effort than the installation of an engine with a civil TC.
FAA Order 8900.1, entitled" Flight Standards Information Management System (FSIMS)" provides field approval guidance to FAA inspectors. Volume 4 (AIRCRAFT EQUIPMENT AND OPERATIONAL AUTHORIZATIONS), Chapter 9 (SELECTED FIELD APPROVALS), Section 2 (Field Approvals of Turbine/Turboprop Engine Installations on Piston-Engine Powered Aircraft), dated September 2007, presented information regarding "the methods of approval for alterations converting aircraft from reciprocating to turbine/turboprop powerplants." The document stated that "As a result of recent concerns expressed by the Federal Aviation Administration (FAA) Aircraft Certification Service, it has become apparent that engine changes to aircraft have been field approved that are beyond the scope of this process. One example is the field approval of the installation of Lycoming T53-L-7A engines on Polskie Zaklady PZL M-18A model aircraft. These engines are not type-certificated products."
Paragraph 4-1213 of the Order stated that "Engine changes that alter an aircraft from a reciprocating engine to a turbine engine require the use of either an amended Type Certificate or a Supplemental Type Certificate (STC). The use of field approvals for these changes is not appropriate because these are major changes to type design. Additionally, some of these alterations were approved using engines and propellers that were not type-certificated products. One source of these engines, propellers, and associated systems is surplus military airplanes. It is recommended that these modifications be removed and the affected airplanes returned to proper FAA approved configuration."
The primary STC process guidance for FAA inspectors was contained in FAA Order 8110.4C. Supplemental or amplifying guidance was provided in policy memos from FAA directorates. Current FAA guidance required that the subject of instructions for continued airworthiness (ICA) be addressed for modification approvals. According to an FAA inspector very familiar with the STC process, when the accident engine was installed in the airplane 20 years ago, the level of emphasis that the FAA placed on ICA was significantly less than current standards.
Accident Engine Installation
The replacement of the reciprocating engine with the turboshaft engine was documented and approved via an FAA Major Repair and Alteration Form 337, also referred to as a "field approval." The applicable FAA Form 337 stated that the "alteration was accomplished by…the adaptation of the T53-L7A Lycoming turbo prop engine, propeller, prop governor, nose cowl and all engine accessories as removed from the Military OV-1B "Mohawk" type aircraft (firewall forward)." The installation was accomplished by a certificated mechanic with an Inspection Authorization rating, and was approved by an FAA maintenance inspector from the Seattle Flight Standards District Office (FSDO).
The FAA Form 337 also stated that the "alteration was additionally installed in accordance with previous identical installation as approved under STC NO. SA5154NM." FAA records indicated that in May 1991, STC SA5154NM was issued to an individual by the FAA Seattle ACO. That STC approved the installation of a Lycoming T53-L7 turboprop engine and Hamilton Standard propeller on a different PZL M-18 airplane. The STC explicitly stated that the approval was limited to the installation only in PZL M-18 serial number 1Z008-06, registration N27237.
Examination of the FAA documentation for the installation of the T53-L-7A engine in the accident airplane revealed that it did not contain any information regarding ICA. Substantiating documentation for the STC was not located during the investigation.
Accident Engine Maintenance History
The accident engine was first installed in the airplane in 1992, but the engine maintenance records obtained by the investigation only dated back to January 1996. The specified "hot end" inspection interval was 600 hours; the maintenance records catalogued four hot end inspections, with calculated intervals (between inspections) of 644, 475, and 732 hours, respectively. The accident occurred about 370 hours after the most recent hot end inspection. The specified overhaul interval was 1,800 hours; the records catalogued a single hot end overhaul, which was accomplished 1,621 hours after the previous overhaul. At the time of the accident, the engine had accumulated about 2,090 hours since its most recent hot end overhaul. Additional maintenance irregularities are detailed below.
In June 1997, at a total time (TT) in service of 2,450 hours and a TT since overhaul (TTSOH) of 721 hours, the engine was removed and subjected to a "hot end" inspection by Garlick Helicopters Incorporated (GHI) in Hamilton, Montana. Maintenance activities were cited to be in accordance with "TM55 2840-229-23." Cross reference of that TM55 document number revealed that it was the maintenance manual for the Honeywell T53-L-13B/BA and/or T53-L-703 engine, but not the subject T53-L-7A engine. Some previous inspections and repairs on the engine had also been conducted by GHI.
In March 1999, numerous hot end components, including most but not all of the fuel vaporizers, were replaced. No specific information regarding those fuel vaporizers, including condition or service history, could be located. In January 2003, many engine components were replaced, but there was no mention of the fuel spray hardware. The entry indicated that the overhaul was accomplished in accordance with "DMWR55-2840-104," which was the correct designation for the overhaul manual for the T53-L-7A engine. In March 2007, the engine was removed and a hot end repair was accomplished by Cappsco, an FAA certificated repair station not associated with GHI. The investigation was unable to locate any specific information regarding that maintenance activity. No other significant maintenance was documented for the engine between that March 2007 activity and the accident. Maintenance records indicated that the most recent annual inspection of the engine was accomplished in April 2010. At the time of the accident, the engine had a TT of about 4,452 hours, and a TTSOH of about 1,102 hours. Refer to the public docket for this accident for additional details.
Engine Maintenance Provider GHI
According to FAA and U.S. Department of Justice (DOJ) information, GHI (the company which replaced the fuel vaporizers, and which also conducted at least one overhaul of the accident engine) had been cited at least twice for violations of Federal Aviation Regulations pertaining to helicopter maintenance. On August 14, 1998, the FAA issued an emergency revocation of the GHI repair station certificate. In 1993 and 1994, GHI was cited for knowingly misrepresenting the service time of helicopter rotor blades that the GHI had maintained.
Accident Engine Examination Results
The engine was examined in detail at a Honeywell Aerospace facility a few months after the accident. The data plate identified the engine as a model T53-L-7A turbo-propeller engine, serial number LE-05431. There was no evidence of fire damage. The engine was mostly intact, and the propeller assembly remained installed. The propeller blades had been cut to facilitate shipment of the engine.
No metallic debris was present on the accessory drive gearbox chip detector. The inner skin of the exhaust diffuser was deformed, and partially pulled out of the exhaust. The four struts of the exhaust diffuser assembly were deformed and displayed impact damage.
The inlet stator and first stage compressor blades appeared intact, with earthen debris in the engine inlet. The inlet and diffuser housing assemblies, axial and centrifugal compressor cases, accessory drive and reduction gearbox assemblies, overspeed governor, tachometer drive gearbox, fuel control assembly, starter, and propeller control system were examined. No anomalies other than impact-related damage were noted.
All power turbine (PT) blades were separated approximately 1 1/2 inches from the root platform, and had rough edges, with varying levels of reverse (opposite direction of rotation) bending. The outboard portions of all gas producer (GP) turbine blades were separated, with the remaining material displaying tearing and battering damage. The GP turbine blade tip shroud was deformed and separated from the GP turbine nozzle assembly, and there was tearing, battering and burning damage to the nozzle. The combustion chamber deflector outer support and many vanes were fractured and separated from the GP nozzle assembly, and metallic debris had accumulated on the deflector.
The inner wall of the combustion chamber liner had liner material displaced radially outward, with the damage to the combustion chamber liner in the approximate axial position of the plane of the GP section. All of the fuel vaporizers were burned, damaged, or both. Refer to the section below, and the public docket for this accident, for additional details.
Combustor Hardware Configuration and Failure Modes
The combustor was a reverse flow-type combustor with multiple vaporizing fuel injectors (referred to as "T-tubes" or "T-canes") and injectors located adjacent to complementary igniters. The fuel vaporizers utilized a coarse injection process, a small amount of mixing air, and prescribed fuel-residence time to vaporize the fuel before its injection into the combustion chamber. Since the fuel residence time in damaged fuel vaporizers is less than that in intact units, fuel can exit the vaporizers in a partially evaporated state.
All the fuel vaporizers exceeded the allowable damage limits specified by the maintenance guidance, and that damage that was consistent with gradual deterioration over time. There was no evidence that any of the separated blade- or nozzle-fragments came in contact with the fuel vaporizers. Large segments of three of the fuel vaporizers had been liberated. According to Honeywell, their failure sequence began with thermal mechanical fatigue, followed by crack initiation and growth, and terminated with segment liberation. All remaining fuel vaporizers exhibited distortion, oxidation, cracking and discoloration associated with sustained moderate temperatures or shorter, high-temperature operation.
Damaged fuel vaporizers, particularly those with missing sections, adversely affect the combustor aero-thermal performance by altering the combustor exit temperature profiles. Alteration of those temperature profiles can result in damage to the engine hot end. In that scenario, engine hot end damage is more likely at low power conditions, where heat transfer is reduced.
The combustor was perforated and distorted in several locations where the turbine blades or other components had impacted the inner liner. Those impact sites had damage signatures that were consistent with them not being exposed to the high temperatures associated with an operating engine. The combustor had some deposits of a soft material, which was consistent with mud from the ground impact. There was no other visible combustor damage.
The outer transition liner exhibited cracking near its inner diameter. It could not be determined whether that cracking was associated with the engine failure and ground impact, or had occurred over time. The engine manufacturer determined that those cracks did not have a significant effect on combustor operation.
Blade Material Failure Modes
The failed blades were examined in the Honeywell Materials laboratory. Those examinations indicated that all failures were consistent with short-term exposure (minutes or less) to metal temperatures in excess of 2,000 degrees F. The separations of the taller/longer blades in the rotor assembly were consistent with a stress rupture fracture mode produced by short-term, high-temperature exposure. The separations of the blades that failed closer to the platforms, and the separations through the platforms of some blades in the rotor assembly, were consistent with overload due to impact by the taller/longer blades. No material defects were observed.