NTSB Factual Report

HISTORY OF FLIGHT

On February 19, 2004, approximately 0245 central standard time, a Schweizer 269D single-engine turbine powered helicopter, N255TP, sustained substantial damage during a hard landing following a loss of engine power during cruise flight near San Antonio, Texas. The commercial pilot and his pilot-rated observer sustained minor injuries. The helicopter was registered to and operated by the San Antonio Police Department for police missions. Dark night visual meteorological conditions prevailed, and a company flight plan was filed for the Code of Federal Regulations Part 91 public use flight. The local flight originated from the Stinson Municipal Airport (SSF), near San Antonio, Texas, approximately 5 minutes prior to the accident.

The pilot reported in the Pilot/Operator Aircraft Accident Report (NTSB Form 6120.1/2) that while en route to a police emergency; he noticed a "slight change in engine noise," and the observer asked the pilot if he had "overtorqued" the engine. Flight instruments, including the torque gauge, were normal. Subsequently, the pilot heard a "bang," followed by an aural engine warning horn, and a loss of engine power. The pilot initiated an autorotation to a vacant grocery store parking lot. During the autorotation, he "pulled up on the collective" to avoid striking power lines. The helicopter impacted power lines, landed hard on the parking lot surface and came to rest upright.

The pilot-rated observer reported that he heard the engine surge and felt the helicopter yaw to the left. He immediately looked at the torque gauge and noted the torque needle "in the red" and told the pilot to watch the torque. As the engine surged again he observed the torque needle "spiking back and forth." Subsequently, the observer "heard a loud bang from the engine" as the engine lost power.

A witness located near the helicopter's flight path observed the helicopter flying in a northwest direction when the "engine started to sputter as if the engine was not firing properly." The helicopter appeared to be swaying back and forth, as the witness heard a "loud backfire." The helicopter "veered sharply to the right with the engine continuing to sputter." As the nose of the helicopter "dropped forward," the witness was no longer able to hear the engine.

A second witness also located near the helicopter's flight path observed the helicopter flying north as he heard a "loud backfire", and then saw white smoke coming from the helicopter.

PERSONNEL INFORMATION

The pilot held a commercial pilot certificate with a rotorcraft helicopter rating, which was issued on November 21, 2002. The pilot's most recent second-class medical certificate was issued on September 30, 2003, had no limitations.

The pilot reported on NTSB Form 6120.1/2 that he had accumulated a total of 493.7 hours of flight time, of which 428.2 hours were in the Schweizer 269D helicopters. The pilot further reported he had accumulated 56.6 hours of flight time within the previous 90 days, and 23.5 hours within the previous 30 days, prior to the accident.

According to the training records, the pilot satisfactorily completed his last company proficiency flight check on November 11, 2003.

AIRCRAFT INFORMATION

The 1997-model Schweizer 269D helicopter was configured for aerial law enforcement response with a maximum capacity of three occupants. The helicopter was equipped with a Rolls Royce 250-C20W turbo-shaft engine (serial number 845043), rated at 420 shaft horsepower.

Review of the aircraft maintenance records revealed the last annual/400-hour inspection was completed on January 27, 2004 at an airframe and engine total time of 3,600.2 hours. The engine was last overhauled on June 20, 2003, at an engine total time of 3,326.1 hours. No open maintenance discrepancies were observed in the engine or airframe logbooks. At the time of the accident, the airframe and engine had accumulated a total of 21.3 hours since the last inspection.

According to the overhaul records, the compressor impeller was replaced, a new rear labyrinth seal, and compressor spline adapter were installed. Entries in the compressor overhaul records stated that the impeller clearance (bump clearance) was measured to be .010 inches at the time of overhaul. The impeller clearance is the measured gap between the compressor impeller and the compressor shroud.

METEOROLOGICAL INFORMATION

At 0253, the automated weather observing system at the San Antonio International Airport (SAT) reported wind from 190 degrees at 8 knots, visibility of 10 statue miles, sky clear, temperature 52 degrees Fahrenheit, dew point 50 degrees Fahrenheit, and an altimeter setting of 30.14 inches of Mercury.

WRECKAGE AND IMPACT INFORMATION

The helicopter impacted the surface of a lighted parking lot in a level attitude and slid approximately 30 feet before it came to rest upright.

Examination of the helicopter revealed that both landing skids were compressed upwards and spread apart. The Forward Looking Infra-Red (FLIR) system was crushed upward into the fuselage structure. The bottom of the fuselage was deformed and buckled along its length.

The cockpit area was crushed from the bottom, upward into the seat structure. Both seat bottoms were deformed and compressed. The fuel tank support shelf was distorted and rivet tear out was noted on both outboard panels of the fuselage. Video of the accident sequence indicated that the fuel cap became loose and an unspecified quantity of fuel exited during the impact with the hard surface.

One main rotor blade was fractured approximately 2 feet outboard of the blade grip and displayed leading edge damage. The blade tip exhibited forward bending. The second main rotor blade remained attached to the rotor hub assembly and exhibited chord bending aft. Marks were noted on the inboard bottom of the blade consistent with power line contact. The third main rotor blade was separated from the rotor head. Two leading edge marks were observed approximately two feet outboard from the blade attach point.

Drive line continuity was established for the main rotor drive shaft and rotor hub through the main transmission. The upper pulley clutch assembly operated properly in both directions, drive and free wheel.

The tailboom remained attached to the fuselage and displayed three impact marks consistent with the main rotor blade strikes. The tail rotor remained attached to the tail rotor gearbox and displayed damage on both blade tips. The tail rotor drive shaft was decoupled from the main transmission aft pinion spline. Continuity was established throughout the tail rotor drive shaft, tail rotor gearbox, hub, and blades.

Examination of the engine revealed that the front of the compressor support was crushed and the compressor would not rotate freely. The upper compressor case halve was circumferentially cracked in the area of the 2nd stage vane band. N1 (Gas Producer) would not rotate. N2 (Power Turbine) rotated freely when the fourth stage turbine wheel was rotated by hand through the exhaust collector. The engine chip plugs were clean with no debris noted.

TEST AND RESEARCH

The engine was examined and disassembled on March 2-5, 2004, at the facilities of Rolls Royce, near Indianapolis, Indiana, under the supervision of the NTSB investigator-in-charge (IIC) and representatives from Rolls Royce and Schweizer Aircraft Corporation. Examination of the engine (serial number CAE845043) revealed the accessory gearbox, N1 (Gas Producer) and N2 (Power Turbine) rotated freely when rotated by hand after the compressor assembly was removed. The mounting nuts on the compressor assembly were observed at 30-40 pounds of torque during the removal from the accessory gearbox. The compressor front support was removed and the number one bearing was found oiled and undamaged.

A rub mark, approximately one-quarter inch in length, was observed on the first stage compressor case blade path. The third, fifth, and sixth stage vane assemblies were bent in the direction of rotation. Damage was also noted on the second and fourth stage vane assemblies. Corresponding damage was observed on the trailing edges of the blades on the adjacent forward compressor wheels.

Prior to the removal of the compressor rotor from the rear support, the impeller travel (bump clearance) was measured at 0.021 inches. No evidence of contact or rub was noted on the compressor impeller or compressor shroud.

The control accessories were removed from the gearbox and were found free of anomalies. After removal of the fuel pump, the splines and shaft were found intact. Rotational continuity was established throughout the shaft. When rotated by hand, residual fuel came out of the fuel pump outlet.

The spur adapter gear shaft and the turbine-to-compressor coupling were intact. The turbine section was disassembled and the number eight bearing sump cover nut was found collapsed and contacted the gas producer tie bolt end which subsequently penetrated the sump cover nut.

All turbine wheels and nozzles were found to be free of anomalies. All bearings were oil wetted, and rotated freely. The exhaust collector was crushed and distorted.

The fuel control unit and power turbine governor were bench tested at the facilities of Honeywell Aerospace Systems, in South Bend, Indiana, on March 5, 2004, under the supervision of the NTSB IIC. No discrepancies were noted during the examination.

Examination of the compressor case assembly by a Rolls Royce metallurgist under supervision of the NTSB IIC revealed a circumferential crack approximately 4.5 inches in length along the second stage blade track approximately 1.5 inches aft of the forward flange. The fracture surface was found to be consistent with overload.

Examination of the compressor rotor assembly revealed that four of the first stage compressor blades were "wrinkled" on the suction side of the blades at approximately 80 percent span. Consistent damage was observed on two adjacent blades on one side of the rotor and two adjacent blades on the other side of the rotor. The fifth and sixth stage compressor blades were bent in the opposite direction of rotation. Small areas of the blades were fractured and missing.

Debris was collected from the first stage nozzle shield area after the turbine section was disassembled. The debris was consistent with compressor blade and vane material.

Scoring and localized fretting were observed on the compressor impeller stub shaft. Localized fretting was observed approximately .025 inches aft of the impeller rotor, and measured 0.1 inches circumferentially. The fretting was consistent with axial movement.

The inner and outer stationary seals on the compressor rear support displayed multiple wear marks consistent with axial movement of the compressor rotor. When compared, the score marks on the inner seal and the width of the rotating labyrinth seal knifes seals indicated an axial movement of approximately 0.016 inches. The comparison of the score marks on the outer seal and the compressor impeller knife seals indicated an axial movement of approximately 0.010 inches.

The rotating labyrinth seal displayed bright score marks on the inner surface of the seal, consistent with damage during disassembly. Localized fretting was observed on the forward face immediately outboard of the bore and on the inner face of the oil slinger. This fretting was consistent with axial movement of the compressor assembly.

Examination of the compressor diffuser revealed cracks in the bond joints between the diffuser aft wall and the turning vanes. Cracks were observed through the diffuser aft wall adjacent to the mounting studs, and localized deformation in the diffuser wall.

ADDITIONAL INFORMATION

The wreckage was released to the owner's representative on July 6, 2004.

NTSB Probable Cause

The loss of engine power due to the axial movement of the compressor rotor blades contacting the compressor vanes resulting in a subsequent compressor stall. A contributing factor was the improper assembly of the compressor section during the engine overhaul by unknown maintenance personnel.