N904PD – McDonnell Douglas 500N – 1999-10-25 NTSB Accident Report

1.1 HISTORY OF FLIGHT

On October 25, 1999, at 1537 hours Pacific daylight time, a MD Helicopter, Inc. (MDHI) 500N helicopter, N904PD, collided with a city street in San Jose, California, following an in-flight loss of control during an approach to the San Jose International Airport. The helicopter, operated by the City of San Jose Police Department as a public-use flight under the provisions of 14 CFR Part 91 of the Federal Aviation Regulations, was destroyed during the collision sequence. The airline transport pilot and nonrated passenger were fatally injured. Visual meteorological conditions prevailed and a company flight plan was filed. The flight originated about 1524 from Reid-Hillview airport and was en route to San Jose International Airport.

While on a routine patrol on October 23, 1999, the pilot and a police observer, who was flying the aircraft, experienced an uncommanded yaw of about 10 degrees to the right while practicing stuck pedal maneuvers. In his statement, the observer reported that after ascertaining that neither of them had made a control input, the pilot took control and made a precautionary landing at the Reid-Hillview airport. The observer said that the pilot reported the pedals felt mushy, but he had control authority. The next day a contract maintenance technician arrived and was unable to find a discrepancy in the anti-torque control system (see Aircraft Information for a more detailed discussion). The pilot and technician believed the problem was in the Yaw Stability Augmentation System (YSAS), and it was deactivated in accordance with the Rotorcraft Flight Manual (RFM) procedures in order to fly the helicopter back to the maintenance base at San Jose International. Prior to departing on the accident flight, the pilot made several left and right pedal turns while in a hover.

A review of the air-ground communications tapes from the San Jose International Air Traffic Control Tower (ATCT), revealed that the pilot contacted the facility about 4 miles east of the airport and requested landing at taxiway "Victor." The pilot was cleared to crossover the airport midfield at or above 1,000 feet agl for a left downwind. He was given the wind, altimeter setting, and transponder code. The pilot read back the instructions. Runways 30R, 30L, and 29 were in use. The pilot stated he was going to Aris Helicopters and was cleared to land on taxiway "Victor" paralleling the active runways. His next transmission at 1535:21 was "Police 1, mayday, mayday," and then again "mayday, mayday" 3 seconds later. Witnesses reported seeing the helicopter yawing erratically then descending rapidly while spiraling.

Controllers on duty in the ATCT observed the helicopter while on downwind for runway 29. One controller observed the helicopter "in erratic flight . . . (then) . . . plummet to the ground." Another controller observed the helicopter "approximately 1 to 1 1/4 mile base leg when it just fell out of the sky." Another controller assumed that the pilot was performing an autorotation and queried the others about the maneuver. The last controller observed the helicopter in a tight spiral on the left downwind and alerted the supervisor.

Twenty-one ground witnesses were identified and interviewed. Of the 21 witnesses, 15 observed the helicopter spin, although only 4 were able to quantify the direction as a "counter clockwise" rotation. Eight of the witnesses said they heard unusual engine sounds and they variously described them as "whining" or "choking/sputtering." One of the witnesses reported a "tighter and tighter" spin as the helicopter descended, while another said it was a "tight 180-degree turn." One witness said the tail "appeared to be bouncing back and fourth," while a second witness observed that the helicopter "started to rotate [one direction] and reversed direction," then spiraled straight down.

1.1.1 Radar Derived Flight Path

Recorded radar data in the form of a CDR Editor listing was obtained from the San Francisco Bay Terminal Radar Approach Control (TRACON). The data was recorded and processed by the TRACON's ARTS IIIA system using a surveillance antenna at the Moffett Federal Airport, located about 8 miles northwest of the accident site. The antenna rotates clockwise at a 4.7-second sweep rate. The data reviewed was the time of each secondary beacon return, the recorded target position, the Mode C altitude, ground track heading, and ground track speed. According to the facility, the ARTS IIIA system computes target ground speed and ground track heading based on a smoothing algorithm, which averages the last several data points.

In addition, recorded plot data was obtained from the San Jose International Airport noise abatement monitoring office. The facility obtains a direct feed of radar data from the Bay TRACON ARTS IIIA system and incorporates sound microphone decibel data to form an integrated picture of the tracks of departing, arriving, or over flight aircraft at the airport. The raw recorded radar data is processed in a computer program, Airport Noise and Operations Monitoring System (ANOMS), written by Lochard Company. The ANOMS program also uses a smoothing algorithm to compute and present target ground speed and track heading information.

Recorded raw radar data (time, position, and Mode C altitude) was also processed in a computer program Tactical Mapping by ERM, Inc. This program computes target point-to-point airspeed and heading without smoothing algorithms.

The tabular and graphically plotted radar data from all three sources is appended to this report.

Review of the data disclosed that the secondary beacon return was tracking west on a heading averaging 270 degrees at a Mode C reported altitude of 1,100 feet and a ground speed, which varied from 104 knots to 119 knots. As the target track passed over the airport midfield, the ground track heading turned left over the time frame 1534:16 to 1535:12 from 268 degrees to 126 degrees. The computed rate of turn was 2.4 degrees per second. Over the same time interval, the ground speed decreased to 115 knots. According to the air-ground communications tapes, 11 seconds after achieving the 126-degree heading, the pilot broadcasted the first "mayday" call. During this 11-second time interval, the Mode C reported altitude decreased to 900 feet as the ground speed decreased to 103 knots. Coincident with the "mayday" call, the Mode C reported altitude jumped from 900 to 1,100 feet in one 4.7-second radar antenna sweep interval, followed in the next 4.7-second sweep with a decrease to 800 feet. Over this same 9.4-second time frame, the ground speed decreased to 80 knots. The last secondary beacon return was recorded at 1535:31. Point-to-point computation of the ground track speed over the last four 4.7-second antenna sweeps yielded speeds of 79.7, 72.8, 56.3 and 23.2, then ground impact.

1.5 PERSONNEL INFORMATION

Review of the Federal Aviation Administration (FAA) Airman Certification records disclosed that the pilot held an airline transport pilot certificate with a multiengine land airplane rating. In addition, he held commercial pilot privileges for single engine airplanes land and sea, and rotorcraft-helicopter. The pilot held an instrument rating for airplanes and helicopters. He also held a flight instructor certificate with ratings for single and multiengine airplanes, rotorcraft-helicopter, and instruments. The most recent second-class medical certificate was issued to the pilot on October 30, 1998, and contained no limitations.

According to San Jose Police Department (SJPD) records, the pilot had accumulated a total flight time of 2,586 hours, consisting of about 1,215 hours of military helicopter flight time, and about 1,266 hours in the MDHI NOTAR (No Tail Rotor) helicopter. In the preceding 90 and 30 days, the pilot had flown 88 and 31 hours, respectively, in the 500N helicopter.

1.5.1 Pilot Training

The records indicate that the pilot successfully completed a MDHI 500N NOTAR Pilot's Recurrent Flight Training Course conducted at the Mesa, Arizona, factory on May 11, 1999. The course included both ground and flight instruction in anti-torque failure emergency procedures in the helicopter. The pilot's SJPD training file is attached to this report.

The pilot was also a member of the California Air National Guard (CANG) assigned to the 129th Rescue Wing based at Moffett Federal Air Field, San Jose. He is a designated military aviator assigned to fly the HH-60G helicopter as first pilot. The HH-60 is the Air Force version of the Army UH-60, commonly known as the Blackhawk. According to the pilot's military records, he transitioned to the HH-60 helicopter during an active duty training period from April through September 1996. As of his last flight with the CANG on October 22, 1999, the pilot had accumulated a total flight time in the HH-60 of 416 hours, with 120 hours flown in the past 12 months. According to the wing safety officer, during the pilot's initial transition training in 1996, he would have been given a very thorough indoctrination in the emergency procedures specific to the HH-60, including the loss of anti-torque capability. In addition to the initial transition training, the wing safety officer reported that pilot had routine refresher training during the year with various unit instructor pilots. His most recent recurrent emergency procedures training in the HH-60G with the unit was on October 15, 1999.

Review of the anti-torque failure procedures for the MDHI 520N and the HH-60G revealed that they are diametrically opposed to one another. A detailed discussion of the two procedures can be found under AIRCRAFT INFORMATION, paragraph 1.6.3. Additional information on this topic can be found in TESTS AND RESEARCH, paragraph 1.16.6.

1.5.2 Maintenance Technician Certification and Training

The maintenance technician who examined the helicopter on October 24 is employed by ARIS Helicopters, the FAA Approved Repair Station contracted by the San Jose Police Department to maintain their helicopters. According to FAA records, the technician holds an FAA Airframe and Powerplant certificate and is the Chief Inspector of the repair station. According to MDHI records, the technician/inspector completed a 2-week maintenance course on the 500N helicopter in July 1997. The fatally injured passenger, also employed by ARIS Helicopters as the inspector's assistant, held an FAA Airframe and Powerplant certificate.

1.6 AIRCRAFT INFORMATION

1.6.1 General

McDonnell Douglas Helicopter System (hereinafter referred to as MDHS) transferred ownership of type certificate H3WE to MD Helicopters, Inc. (hereinafter referred to as MDHI) on February 18, 1999. Boeing acquired McDonnell Douglas, including the helicopter division, and subsequently spun off the commercial line into McDonnell Douglas Helicopter System. At the time of the accident, Boeing was providing engineering support functions to MDHI for the commercial helicopter product line.

The 500N helicopter is a five place, turbine powered, rotary-wing aircraft constructed primarily of aluminum alloy, while the tail boom and anti-torque thruster assemblies are primarily a graphite composite. The main rotor is a fully articulated five-bladed system.

1.6.2 NOTAR Anti-Torque System Description

According to the manufacturer, the NOTAR (no tail rotor) design provides anti-torque control by using low pressure, high volume air ducted through the tail boom. The anti-torque control system has sufficient authority to induce large and prolonged sideslip angles at cruise airspeeds. Some of the low pressure air flows out through downward oriented slots along the right side of the tail boom, which combines with the rotor downwash to create a circulation controlled low pressure area on the boom's right side. The balance of the low pressure air flows out through a pilot controlled directional jet thruster at the end of the tail boom. A variable pitch fan with the blade pitch controlled by anti-torque pedal inputs is enclosed in the aft fuselage section immediately forward of the tail boom and driven by the main rotor transmission through a fan gearbox and drive shaft. The circulation control tail boom, jet thruster assembly, horizontal stabilizer, and two vertical stabilizers make up the rest of the anti-torque system. In forward flight, the vertical stabilizers work with the thruster assembly to provide the required anti-torque forces as well as directional control to maintain fuselage heading. The left vertical stabilizer is controlled by the cockpit pedal movement, while the right vertical stabilizer is controlled independently by the YSAS system.

Control of the anti-torque system is by pilot pedal inputs and torque/push-pull tubes to a splitter assembly located over the left cabin area at FS113. From the splitter, one torque tube controls the fan blade pitch while another control goes to the left vertical stabilizer and jet thruster cone by way of a two-part thruster cable, which terminates at a bellcrank beneath the horizontal stabilizer. The bellcrank transmits motion via another cable to a pulley assembly that rotates the jet thruster cone. The cables are flexible and have a quick disconnect fitting at the tail boom to fuselage junction, which joins the forward cable to the center cable. The left pedal input pulls the cable forward and applies a tension load and the right pedal input pushes on the cable and applies a compression load. The jet thruster cone rotates on the end of the tail boom to direct the high volume, low-pressure air left or right.

At the forward end of the forward thruster cable, the cable housing is fixed to a bracket on the fuselage at station 123.3 and the rod end of the internal (Teflon wrapped) cable is attached to the control splitter assembly, also referred to as the "Station 113 bellcrank." In between the fuselage bracket and the splitter, the flexible internal cable is supported by a two-part telescoping sleeve. One part of the sleeve, the forward part, is attached to and travels with the rod end and internal cable. The aft part of the telescoping sleeve is attached to the fuselage bracket by means of a ball (swivel) coupling and remains stationary with the fuselage bracket. The two halves of the telescoping sleeve slide within one another to prevent the flexible interior cable from bowing. The angular misalignment induced by the splitter assembly travel is accommodated by a ball coupling on the telescoping sleeve.

1.6.3 RFM Emergency Procedures

Review of the Emergency and Malfunction Procedures section of the FAA approved RFM for the Model 500N revealed that sub-section 3-9 discusses ANTI-TORQUE FAILURES. The section states, "Different types of failures may require slightly different techniques for optimum success in recovery . . . therefore, it is not possible to provide a standardized solution for an anti-torque emergency." Two cautionary notes are listed for ANTI-TORQUE FAILURES. One states, "Do not attempt an autorotation from forward flight unless an actual engine failure occurs." The other note warns the pilot, "Do not attempt flight below 20 knots" during an ANTI-TORQUE FAILURE. Powered run-on landings are to be made, with throttle manipulation used, to assist in maintaining directional control.

USAF technical order "TO 1H-60(U)A-1" is the military equivalent of the RFM for the HH-60. The emergency procedures section dealing with anti-torque drive failures in cruise were reviewed. The immediate pilot action items are; 1) Autorotate, and 2) Throttles to Idle. Three warnings are prominently positioned in the section. The first one states, "Attempts to maintain powered flight may result in unrecoverable loss of control or tail structural failure." The second warning note states, "If autorotation is delayed, excessive yaw angles will cause low indicated airspeed...[which] can make it more difficult to establish or maintain autorotation." The final warning note reads, "Left pedal application will cause . . . decreasing yaw control."

1.6.4 Maintenance History

The helicopter, SN LN032, was issued a standard airworthiness certificate in the Normal Category. Review of the helicopter maintenance records disclosed that at the time of the accident, the helicopter had accumulated 5,972.3 total flight hours. The last documented 1,200-hour and annual inspections were combined and occurred on September 30, 1999, at 5,900 total flight hours. The logbook entry notes that 100, 300, and 600-hour inspections were accomplished at the time of the annual inspection, and were performed in accordance with Boeing Form 1594 and the Chapter 5 inspection checklist. Beyond the October 23 uncommanded yaw event, no unresolved maintenance discrepancies existed. The tail boom was last demated from the fuselage on July 23, 1999, for replacement of the anti-torque system fan liner. During the annual inspection on September 30, the thruster cable tensions were adjusted. According to MDHI, tension can only be adjusted on the aft cable.

A maintenance records entry noted that both "forward and aft thruster cables" were removed for replacement on July 26, 1995, due to an unusual noise in the cable; the entry referenced part numbers 500N7201-37 and -45. According to MDHI, the part numbers referenced (500N7201-37 and -45) refer to the forward and center cables. The cables were replaced at 1,970 total hours. The representative of ARIS Helicopters produced documentation showing that the removed cable assembly was returned to MDHS on July 27, 1995. MDHI could not produce any records that the cables were received. At the time of the accident the replacement cables had accrued about 4,002 hours of operation.

Review of the maintenance manual for the helicopter revealed that at the time of the accident, there were recurring 100- and 300-hour inspections of the cable system. The cables have no in service time limits.

1.6.5 Maintenance Actions Following October 23 Uncommanded Yaw

Following the unusual feel to the anti-torque pedals during the patrol flight on October 23 and the precautionary landing at Reid-Hillview airport, the chief inspector/contract maintenance technician for the operator arrived at Reid-Hillview to inspect the helicopter on October 24, which was his regular day off work. According to the technician's statement, during a discussion of the previous night's events, the pilot stated that the anti-torque pedal movements did not feel normal to him. The pilot told the technician that he thought the left pedal was too far forward in the relaxed state. The technician said he could not recall exactly the relaxed state relative positions of the pedals and moved them left and right from stop to stop; he said they felt smooth without binding, but with what seemed to be an excessive amount of spring pressure. According to MDHI, at rest the maximum pedal misalignment shall be 0.50 inches.

The technician said he then inspected the anti-torque system in detail without finding any obvious discrepancies. According to his statement, he first rotated the thruster cone, which had a "normal" feel, and noted a corresponding movement of the left vertical stabilizer. He next removed the thruster cone and inspected the rear cable and associated bellcranks and pulleys. The technician then removed the front seats and center console and inspected the anti-torque system torque tubes and bungee in the cockpit area. He next looked at the Station 113 bellcrank/splitter by looking down the engine inlet from the front of the helicopter without removing the left-hand engine inlet fairing to gain complete access to the bellcrank. The technician observed the movement of the fan blades through the air inlet screen as the pilot moved the anti-torque pedals. As a final step, the technician examined the anti-torque cable quick disconnect fitting at the tail boom to fuselage join point and observed the cable connection slide smoothly back and fourth with the fitting cover moved to expose the connection.

As part of the system inspection, the technician performed a preflight check of the YSAS System as specified in the RFM. During the check, the trailing edge of the right vertical stabilizer is held lightly with fingertips. With the YSAS systems powered up and engaged the tail boom is moved back and forth by hand. The YSAS rate gyro is designed to sense the induced yaw and compensate by moving the right vertical stabilizer. The technician induced a simulated yaw while the helicopter was on the skids. The stabilizer did not move. The pilot, observer, and technician concluded that the YSAS system had failed the night before. The pilot's flight manual addresses the ground check in section 4-6 and 4-7, and is attached to this report.

On the next day, the day of the accident, the maintenance technician arrived at the helicopter in the late morning with the intention of performing a full yaw system rigging check. In his statement, he said that while reading through the maintenance manual, he decided not to actually re-rig the helicopter, but to check the "bottom line numbers" of the relative movements of various system components specified in a rigging procedure. He stated that he made up a matrix to record and check the various measurements. The left hand inlet fairing cover was removed to gain access to the Station 113 bellcrank/splitter assembly and the rigging of the pedals, fan blade pitch, thruster cone, and left vertical stabilizer were checked. The mechanic did not observe any discrepancies. Subsequent examination revealed that the failed portion of the telescoping sleeve ball coupling swivel fitting was on the interior side of the fitting and could not have been observed without the aid of an inspection mirror.

According to the maintenance technician, prior to departure on the flight back to San Jose International, the pilot performed three separate hover pedal turn series, both left and right. The pilot then announced over the loud speaker "everything felt good."

1.6.6 YSAS System Description

The YSAS system is composed of three components, a yaw rate gyroscope, a computer, and an electromechanical actuator connected to the right vertical stabilizer. According to MDHI, the YSAS system has a control authority equal to 10 percent of the total yaw control system. The right-hand vertical stabilizer has a range of movement of 15 degrees left and right. The YSAS system is effective at speeds above 40 knots but has little or no effect below this speed value.

1.7 METEOROLOGICAL INFORMATION

The San Jose International Airport METAR at 1537 was reporting in part: sky clear; visibility 20 miles; winds from 290 degrees at 12 knots; and altimeter setting 30.13.

According to the air-ground communications transcript, when the pilot was cleared to enter the San Jose airspace, he was given wind 300 degrees at 13, and the altimeter setting of 30.13 inHg. Subsequently, he was cleared to land on taxiway "victor" and given a wind update of 290 degrees at 12.

No unusual weather phenomena were noted during a review of the meteorological reports, or, observed by any witnesses.

The San Jose International airport noise abatement monitoring system and the recorded TRACON radar data show there were no go-arounds or over flights of the airport by air carrier or other large airplanes within 30 minutes prior to the accident.

1.8 WRECKAGE AND IMPACT INFORMATION

Safety Board investigators examined the wreckage at the accident site location of the Alameda and Alameda Way streets in a San Jose residential/business area about 1 mile southwest of the airport. The helicopter impacted near the center lane of a five-lane city street. Trees and power lines up to 100 feet agl surrounded the wreckage, and a three-story building was immediately south of the road. The main fuel cell was ruptured and a fuel spill was observed on the street pavement.

The fuselage was observed to be oriented on a 050-degree magnetic bearing. Extensive vertical crushing and structural collapse was evident to the fuselage. The left skid was separated from the cross tubes and folded under the fuselage while the right skid was separated and folded outboard. The crush lines on the fuselage were observed to be 10 degrees nose down for the longitudinal axis and 30 degrees left side down for the lateral axis. The majority of the wreckage was confined to about a 50-foot radius of the fuselage center, with a section of the blue rotor blade located about 60 feet from the fuselage. The right forward cabin door was found about 41.4 feet northeast. Gyro stabilized binoculars were found in an office building atrium about 100 feet southwest of the main wreckage.

The blue blade fractured just outboard of the root doublers, with upward bending distortion evident. The remaining four rotor blades were attached to the rotor hub. Chordwise scuffing was found on the tip ends, and a circular scar was found on the asphalt street over the left forward quadrant in front of the wreckage at the approximate rotor disk radius. All lead-lag dampers were intact and secured to their respective blades. The pitch change links were secured to the blade pitch horns and the swashplate.

The cockpit collective, cyclic, and anti-torque control system components beneath the cockpit floor were crushed and distorted. The left and right cockpit yaw control pedals were fractured and separated from their respective posts. All fractures to the system push-pull and torque tubes were accompanied by bend deformation to the associated tubes. The lateral and longitudinal trim actuators and all associated components were crushed and bent. The upper lateral, longitudinal and collective pitch control rod end bearings at FS78.5 were intact, connected to the mixer assembly, and were undamaged.

The upper end of the FS78.5 yaw control rod remained attached to the FS97.5 bellcrank. The FS97.5 bellcrank mounting bracket was damaged with an impression of the bellcrank arm present on the front side of the bracket (rearward movement of the arm). The control tube, which connects the FS97.5 bellcrank to the FS113 splitter was fractured at the rear rod end bearing, with 45-degree shear lips evident on the grainy textured fracture face. The FS113 splitter assembly rack had punctured the upper fuselage deck with a rearward and downward motion beyond the limit stops. The splitter spring was intact and attached to its respective lugs.

Control continuity was established from the splitter to the fan blade pitch mechanism. At the FS113 splitter, the intermediate control rod (controlling pitch of the NOTAR fan blades) was attached to the splitter arm and the FS137 bellcrank, but was bent about 15 degrees. The aft control tube, which transmits motion between the FS137 bellcrank and the fan transmission, was bent about 10-degrees at the forward fan drive shaft coupling, which was fractured and separated. The fan and its gearbox transmission rotated freely.

The forward thruster cable rod end bearing remained attached to the FS113 splitter assembly. During laboratory examination of the telescopic swivel ball joint it was determined that it had been separated from it's swaged fitting. A fracture was noted that extended about 180 degrees around the swaged fitting, with a longitudinal fracture that connected each end of the circumferential fracture to the fitting end. The missing portion of swaged fitting was not located at the accident site. The circumferential fracture face appeared corroded and the longitudinal fractures had both corroded and shiny sections. Other longitudinal cracks were apparent around the remaining circumference of the swaged fitting end. Subsequent investigation disclosed that once liberated from the swaged fitting, the ball swivel telescoping rod end would move away from the swaged fitting exposing 2.95 inches of inner Teflon coated cable. The Teflon coating of the inner cable was abraded/missing over about a 2.5-inch span, with fragments of Teflon material adhering to the sharp edges of the fracture. (See the Seal Laboratory report appended to this file).

The forward and center thruster cables were disconnected from each other at the quick disconnect fitting at the tail boom to fuselage juncture (FS174). The sliding fitting sleeve, which covers the male/female cable joint fitting, was damaged, crushed, and distorted in shape. The tail boom slot for the cable immediately aft of the fitting sleeve was damaged and a rubber grommet was torn. The center cable aft rod end bearing remained attached to the left-hand horizontal stabilizer control assembly; however, the thruster and vertical stabilizer torque tube housing assembly, ref: part number 600N7200-501, was broken from the structure below the horizontal stabilizer. The FS264 composite support bracket was found fractured and separated from its mounting point. The conduit cap immediately forward of the composite support bracket was circumferentially fractured, with corrosion apparent on the fracture face.

The control rod end bearings from the thruster and vertical stabilizer control torque tube to the aft thruster cone pulley remained attached, and, the cable that runs between the two pulleys was intact. The control torque tube assembly, P/N 600N7200-501, was severed from it's structural mount. The jet thruster cone was found in the full right yaw position with evidence of over travel.

The tail boom was diagonally severed at a point just forward of the horizontal stabilizer and within the rotational radius of the main rotor disk. Witness marks and paints transfers were found on the tip end of the blue main rotor blade, which matched the fuselage color at the point of tail boom separation. The center thruster cable exhibited abrasion marks and permanent set deformation at the point of tail boom severance.

The horizontal stabilizer remained attached to the severed end of the tail boom. The right vertical stabilizer was intact. The left vertical stabilizer upper section was intact; however, the lower section was damaged, and had separated from the structure and was found adjacent to the tail boom.

The cockpit throttle and the engine fuel control index pointers were both found at the ground idle position. Instrument indications were zero except for the N1 gauge, which read 41 percent

1.13 MEDICAL AND PATHOLOGICAL INFORMATION

On October 26, 1999, the Santa Clara County Medical Examiner performed an autopsy on the pilot. During the course of the procedure, the FAA Civil Aeromedical Institute in Oklahoma City, Oklahoma, obtained samples for toxicological analysis. According to the report, the analysis was negative for carbon monoxide, cyanide, ethanol, and all screened drug substances.

An autopsy was performed on the right seat passenger by the Santa Clara County Medial examiner on October 26, 1999. The following information was extracted from autopsy report.

The cause of death was listed as "Multiple Traumatic Injuries."

Noted Under "External Injuries":

"Head: Two juxtaposed 2-inch-long lacerations split the back of the head slightly above the occipital prominence [at the back of the head], beneath which a shattered calvarium [upper portion of the skull] is palpable."

"The face presents with a bruise mark above the left eyebrow. A 2-inch-long gaping laceration occupies the left side of the chin, along the left mandible. There are two smaller, 1 1/2-inch-long lacerations overlying the right mandibular angle. There is presence of blood within the ear canals bilaterally. ..."

"The anterior aspect of the abdominal wall is covered by a well-defined, horizontal, 1-inch-wide bruise mark at the level of the umbilicus [navel] and extending laterally to the flanks. There are also some additionally faintly visible bruise marks slightly above the above described horizontal bruise mark overlying the right upper quadrant. ..."

Under "Internal Injuries" is noted:

"Head: Focal subgaleal hematoma [bruising beneath the scalp] formation at the site of the above described laceration. The calvarium and the bones of the base of the skull present with a hinge-type fracture passing across the petrous processes and the sella turcica and extending onto the right parietal bone and the occipital bone posteriorly . ..."

"Chest: ...The thoracic aorta shows a complete transection [tear] slightly distal to the origin of the left subclavian artery. ..."

1.16 TESTS AND RESEARCH INFORMATION

1.16.1 Sound Spectra Analysis

A certified copy of the air-ground communications recording from the San Jose ATCT was sent to the Safety Board's audio laboratory for a sound spectrum analysis. Three transmissions were associated with the accident helicopter. The first one occurred over a 1.5-second interval and included the first mayday transmission by the pilot. The second transmission, 0.25-seconds in duration, was a microphone keying without voice. The final 1.4-second transmission contained the second mayday voice overlay. Signals containing the fundamental frequency, and the second and fourth harmonics, of the main rotor transmission was identified in all three radio transmissions. The main rotor transmission speed was found to be near 100 percent in the first two transmissions, and decreasing through 82 percent in the third radio transmission.

1.16.2 Thruster Cable Laboratory Examination

The forward and rear thruster cables were removed from the helicopter for detailed laboratory examination at Seal Laboratories in El Segundo, California. The complete report is appended to this file.

According to the metallurgist who performed the visual and Scanning Electron Microscope (SEM) examinations, the circumferential fracture around the rear cable conduit cap revealed "a corroded and brittle intergranular fracture, typical of stress corrosion cracking." Small locations of clean ductile overload fracture were also observed near the inner diameter surface. The report states "that the conduit cap . . . cracked from the outside diameter all the way around the circumference due to stress corrosion . . . [and] . . . the final ductile overload fracture near the inner diameter surface is most probably caused by the final separation due to blade strike." The chemical composition of the conduit cap was consistent with the MDHI specification for the part.

The visual and SEM examination of the forward cable ball swivel swage fitting fracture revealed "a corroded and brittle intergranular fracture, typical of stress corrosion cracking." Small locations of clean ductile overload fracture were also observed near the inner edges of the longitudinal fractures. Energy Dispursive X-ray (EDX) microprobe spectra of a flake of translucent material recovered from both the inner diameter of the circumferential fracture and from the inner diameter surface of the ball interface were consistent with wear residue of the Teflon sleeve. A 2.5-inch section of Teflon sleeve was worn off the inner cable and two additional splits were observed about 1-inch from the worn end.

According to the laboratory report, the steel core wire in the abraded area of the rear(center) thruster cable at the point of tail boom severance was found locally stretched.

According to the Federal Aviation Administration (FAA) AC43.13-1B, section 6-20 STRESS CORROSION CRACKING: "This form of corrosion involves constant or cyclic stress, acting in conjunction with a damaging chemical environment. The stress may be caused by internal or external loading." Metallurgical analysis of the fractures revealed the presence of chlorine, typical for operation in a corrosive environment.

1.16.3 Simulation of Cable Failures

On November 3, 1999, Safety Board investigators and parties to the investigation held a meeting at the MDHI facility in Mesa, to address the P/N 500N7201-45 center thruster cable. A test protocol was devised to explore the effect of a completely severed conduit cap at the location where the stress corrosion crack was found. An exemplar cable severed at the same location as the accident helicopter's cable was installed in a MDHI supplied 500N helicopter. After installation, visual observation revealed that the two pieces of conduit cap were separated by about 0.5-inches. The anti-torque pedals were then exercised to full left and right deflection and it was noted that the conduit cap separation had little affect on the thruster cone operation. An actual rigging check was not performed.

After the meeting with the Safety Board, MDHI personnel safety wired the severed thruster cable housing together and performed a cable installation rigging procedure. They then cut the safety wire, allowing the severed thruster cable to assume a relaxed condition with the 0.5-inch of cable housing separation. The rigging was again checked. According to MDHI, the net affect was a slight over-travel of the left vertical stabilizer in both directions, but still within operational limits.

On November 17, 1999, the investigation again convened at the MDHI facility in Mesa to examine the effect of the separation of the forward thruster cable ball swivel joint. Unlike the short linear separation in the relief area of the center cable conduit cap, the tests showed that the swivel joint separation had the potential of exposing 2.95 inches of the core control cable, depending on pedal inputs. MDHI prepared an exemplar cable with the swivel ball separated from the swaged joint of the conduit cap at the location of the accident helicopter's forward cable swivel joint fracture and installed it on a 500N aircraft. Left and right full travel anti-torque pedal inputs were again applied (right pedal input pushes the cable for right thrust, and left pedal travel pulls the cable for left thrust). On application of full right pedal, the telescoping rod compressed completely, then applying full left pedal the ball traveled with the telescoping rod and exposed about 2.95 inches of core control cable. Full right travel then caused bowing of the core control cable and failed to move the jet thruster cone and left vertical stabilizer to the right. The control of the fan blade pitch was not affected.

1.16.4 Aerodynamic Loading of Cables

According to MDHI, directional control of this model helicopter in forward flight is in part a function of the aerodynamic loads placed on the left vertical stabilizer that is controlled by the anti-torque pedals. The right vertical stabilizer is controlled electronically by the YSAS. Analysis by MDHI of the flight dynamics of the NOTAR control system for the 500N indicates that, in forward, non-yawing (non-sidewise) flight, airspeed must be reduced to less than 30 knots prior to the stabilizers becoming aerodynamically "unloaded." Until such time as the airspeed is reduced, or the helicopter is flown out of trim (yawing or sidewise flight), the aerodynamic loads on the left vertical stabilizer will maintain a tension load on the thruster control cables. The MDHI analysis indicated that even with both the rear cable conduit cap and the forward cable swivel end pulled from the swaged fitting, the tension load on the thruster control cables will enable continued in-flight directional control as long as the forward airspeed remains above 30 knots and flight of the helicopter remains in trim (not in yawing/sidewise flight). MDHI further stated that the tension load, if present, would prevent any "bowing" of a fractured forward cable housing/swaged fitting until such time as the aerodynamic loads are reduced and large displacement pedal applications are made.

Boeing-Mesa Accident Investigation Report 991025/500N/LN032, revised: September 19, 2000, pertaining to this accident (attached), states in section 6, Directional Control System Loads, that several flight conditions are possible wherein the tensile loads on the thruster cable, referenced above, decrease to zero and become compressive loads. One of the conditions that produce a compressive load in the cable is "Right Sideward flight" (right sideslip/left yaw angle). In the same section of the report, Table 1: Thruster Cable Loads, MD500N, shows that compressive loads as high as 26.6 pounds were measured on the thruster cable in right sideward flight during flight testing. In a separate correspondence from MDHI to the Safety Board on June 29, 2000, on the subject of model MD500N anti-torque control power, an engineer at MDHI said he "would predict that any of our aircraft would spin 360 [degrees] if full left pedal were input and held at 60 knots for a variety of initial conditions." The engineer also noted "Some aerobatic flights in the MDHI 520N have done 360 [degree] spins down the runway at 40 knots or more." Other maneuvers that produced compressive loads on the cable were rearward flight, some hover maneuvers (yawing turns to the left), steep partial-power descents at less that 30 knots, and during the landing flare.

1.16.5 Effect of Power Manipulation on Yaw Control

In the MDHI analysis of a cable failure scenario, proper pilot response will depend on the direction of yaw. For a right yaw, reduction of power will help arrest the yaw rate, while an increase in power will exacerbate the problem. For a left yaw, a power increase will reduce the yaw rate, while a decrease in power will aggravate the yaw/rate. According to MDHI, left yaw rates in the event of a power reduction can reach intolerable proportions. MDHI stated that a pilot experiencing an anti-torque failure resulting in a fixed thruster in the left direction should: 1) use a shallow to normal approach into the wind or with a right crosswind; 2) plan to touchdown with little or no forward speed; and 3) maintain directional control with small adjustments in throttle and or collective control. MDHI noted that it is unlikely that an autorotation can be accomplished. These conditions were not flight-tested and the published emergency procedure in the RFM was derived following system analysis.

1.16.6 RFM Emergency Procedures

Review of the RFM disclosed that section 3.9 addresses the emergency procedures concerning coping with various anti-torque failures. Fixed thruster conditions (jammed anti-torque controls that make control of the direct jet thruster direction impossible) are addressed. The RFM states that power must be adjusted to maintain a speed between 50 and 60 knots. A right fixed thruster condition requires the use of a shallow approach and running landing, with touchdown into the wind at a speed between 20 and 30 knots; directional control is maintained with small adjustments of the throttle and collective controls. A left fixed thruster condition also requires a shallow to normal approach either into the wind or with a right crosswind, with touchdown at little or no forward airspeed; directional control is maintained with small adjustments of the throttle and collective controls. The RFM emergency procedures section does not address methods to minimize airflow to the jet thruster and circulation control slots.

1.16.7 Yaw Pedal Input Effect on Air Supply to Tail Boom

As noted in the aircraft information section of this report, air is supplied to the tail boom circulation control slots and the jet thruster nozzle by a variable pitch fan, which is powered by a drive shaft from the main rotor transmission. The pitch of the fan blades is controlled by the cockpit anti-torque control pedals. Pedal motion left and right is transmitted through torque tubes and bellcranks to the FS113 splitter assembly. One arm of the splitter is attached to a push/pull tube, which goes to the fan blade pitch change mechanism, while the second splitter assembly arm is attached to the forward thruster telescoping sleeve ball swivel coupling. With the cockpit pedals in a neutral position, the fan blades operate at a 26.5-degree pitch. At full right pedal, the fan blades move to 59 degrees of pitch. As the pedals are moved from full right through neutral to full left, the blades return to 26.5 degrees then increase to 71 degrees. MDHI stated that a failure and separation of the telescoping sleeve ball swivel coupling, and a subsequent bowing of the forward thruster cable, would have no effect on the pitch change capability of the fan blades; the fan blade pitch would continue to respond to the cockpit pedal movements.

1.16.8 History of Corrosion Cracking in Thruster Cable

MDHS, the predecessor company, verbally stated during the November 3, 1999, meeting that they became aware of a cracking problem in the relief area of the center thruster cable in 1998. MDHS stated that they supplied those cables with cracked conduit caps to Cablecraft for evaluation. Cablecraft evaluation determined the cause of cracking to be fatigue. MDHS disputed the findings of the cable manufacturer and performed a metallurgical analysis of the fractured components and determined that the cause of the cracking was due to stress corrosion. MDHS requested Cablecraft at that time to change the heat treatment of the conduit cap from H900 to the H1025 condition. Cablecraft agreed to make the change to the heat treatment of the conduit cap in November 1998, although the actual change did not occur until April 6, 1999, at drawing revision C of the conduit cap design.

According to MDHI, Cablecraft, the manufacturer of the forward and center thruster cables, was first made aware of a center thruster cable-cracking problem on July 24, 1998, by way of MDHI SOR (Service Operating Reports) reports. Cablecraft believed the problem to be a result of vibrations that were fatiguing the conduit cap because the cap is hard mounted to a bulkhead. MDHC conducted metallurgical examinations of the same exhibits and reported fracture features typical of stress corrosion cracking, with no fatigue observed.

The cable manufacturer reported that a heat treat process was changed in the drawings from H900 to H1015 on April 15, 1999. The Cablecraft drawing number was changed from Revision B to Revision C, which was dated April 18, 1999. MDHI followed with a change to the part number 500N7201 nomenclature, which added new dash numbers in order to discern between cables that had the new heat treatment and those in the old condition. Any cables with a date stamp after August 1999 were manufactured with the altered heat-treat process. The markings on the accident center thruster cable were 02731/500N7201-45, Rev, C, MFR 06970, 10/93. The forward accident thruster cable was marked 02731/500N7201-37, Rev. C, MFR 06970, 01/93. The specification for those dash numbers calls for a H900 heat treat condition.

At the time of the accident, the FAA Service Difficulty Safety Reporting System in Oklahoma City was queried regarding anti-torque system failures or problems for the MDHI NOTAR series helicopters. No anti-torque system problems were found in the database.

Safety Board investigators requested a history of known thruster cable malfunctions or defects from MDHI. On November 2, 1999, MDHI provided a copy of their Service and Operations Reporting System (SOR) addressing the thruster cable performance. Of the 17 SOR condition reports supplied, the earliest was dated October 19, 1991, and the most recent was June 16, 1999, with no reports for 1993 or 1994. The list covered both 500N and 600N models. A breakdown of the 17 reports revealed 6 were cracks in the relief area of the conduit cap. The six reports were dated June 1998 and November 1998. The other reports were for rough operation, "noise" and installation difficulty. One of the reports was from Aris Helicopters, Ltd. on behalf of SJPD for "noise in cables." There were no reports of cracks in the swivel ball retaining cap/swaged fitting. The SOR reports are attached.

The FAA Aircraft Certification Office (ACO) with oversight responsibility for MDHI was contacted regarding the history of the anti-torque control system. According to the engineer responsible for the 500N and 600N series helicopters, MDHI reported the corrosion cracking condition in the center thruster cable in October 1998. A meeting between the engineer and MDHI personnel occurred on January 26, 1999. The FAA and MDHI agreed that the cable cracking was not a safety of flight issue and the company was given until January 2000 to solve the cracking problem in the relief area. The FAA and MDHC engineers agreed that the issue presented was considered a minor service difficulty and was not determined to be an unsafe condition, which would warrant the issuance of an airworthiness action. According to the ACO, the possibility of failure at the telescopic swivel end of the forward cable end was not discussed or reported in any correspondence to the FAA prior to the accident.

1.16.9 Other Component Tests

The trim switches were removed from both cyclic control sticks and examined for condition. The switches appeared to be near new condition. Seal Laboratories report is attached.

The Power Analyzer and Recorder (PAR) system was removed and shipped to the manufacturer, Avionics Specialties, Inc., for readout. According to the report, no exceedence or caution timing events were in progress at the time of the accident. The PAR computer appeared to be operating correctly until power was removed by the crash. Their report is attached.

The Turbine Outlet Temperature (TOT) gauge was shipped to the manufacturer, Diamond J, Inc., for recovery of non-volatile information. According to one of the regular pilots, the TOT gauge was not routinely reset. No current useful information was retrieved from the memory. Their report is attached.

The helicopter's YSAS system components were removed and sent to the manufacturer for functional testing. All the components met acceptance standards.

The annunciator light examination revealed stretched filament elements on the "GENERATOR OUT" and "160F BATTERY OVER TEMPERATURE" warning lights. The "BATTERY 140F" filament element was not stretched. Photographs taken by the SJPD shortly after the accident revealed an illuminated "BATTERY TEMP 140F" warning light with the master electrical switch in the battery position.

The Ni-cad aircraft battery was removed and taken to an FAA approved repair station for examination and functional testing. According to the attached report, "based on the service history and physical evidence available, the battery appears to have been functioning normally at the time of the accident."

1.16.10 Engine Examination

The Rolls-Royce Allison (RRA) 250-C20R/2 450 Shaft Horse power engine, serial number CAE 295404, showed extensive damage as a result of impact forces. The N1 section rotated freely but was not continuous to the starter/generator. There was extensive crushing damage to the starter/generator cooling fan. Removal of the starter/generator revealed all of the blades on the cooling fan were sheared off of the wheel, and the remainder of the wheel was bound as a result of the shroud crushing onto it. The starter/generator drive shaft was torsionally sheared. The N2 section could not be rotated as a result of deformation of the exhaust collector. Engine mounts were broken and the engine had shifted in the airframe. The power turbine governor pointer was indicating 60 and the fuel control pointer was indicating 30. The indications were considered to be unreliable because the power turbine governor linkage was broken and the fuel control linkage was bent. The engine was removed from the airframe and transported to National Airmotive Corporation, Oakland, California. The engine was not capable of being run on a test stand because of extensive impact damage and was disassembled for internal examination.

A pneumatic system check was performed with no leaks noted. One tang on the fuel pump's mounting flange was broken as a result of impact damage. The fuel control throttle shaft was bent and the drive body was cracked as a result of impact forces. Both the upper and lower magnetic chip indicator plugs were clean. Engine oil contained in the accessory gearbox was clean and had a normal aroma. All internal engine shafting was intact. Following removal from the exhaust collector, the N2 rotor rotated freely. All engine bearings were well oiled and rotated free and smooth.

The cooling fan on the starter/generator was broken and bound as a result of the generator cooling fan cover crushing from impact forces and contacting the cooling fan. The inner surface of the cooling fan housing showed deep rotational gouging. The starter/generator drive shaft was torsionally sheared and the fracture was consistent with a sudden stoppage of the generator with the engine operating.

The fuel nozzle was heavily carboned and the bench spray test indicated slight streaking, but all test points were within the normal range, with the exception of one high point at 125 psi. According to RRA specification data, the maximum fuel flow for this point should have been 70 pph (pounds per hour) and the reading was 70.4 pph. There were no indications of hot spots on the first stage nozzle as a result of streaking.

Some compressor vanes had small pieces of aluminum foil folded over and shaped to the vane. There were indications of aluminum spray on the first stage nozzle shield. All wheels and nozzles appeared normal with the exception of the third stage turbine wheel and fourth stage nozzle, which exhibited heavy rotational scoring.

1.16.11 Power Turbine Governor Tests

The power turbine governor was bench tested at National Airmotive Corporation. The unit did not meet serviceability test criteria. It was subsequently bench tested two additional times; once at Allied Signal, Burbank, and once at Allied Signal, South Bend. The engine governor, S/N 18946, was installed on this engine September 3, 1997, and had accumulated 1,972 hours since overhaul. The governor has a 2,000-hour overhaul requirement, and had 28 hours service life remaining. The data from the three bench tests did not differ significantly. The unit was subsequently disassembled at Allied Signal, South Bend. There were no failed components or abnormal conditions noted during the disassembly inspection. According to Allied Signal, the functional test data indicated an early rpm condition at the Pg spring set point and a shallow governor slope. The early Pg spring set point is field adjustable during the power turbine governor rigging procedure. The shallow governor slope noted during functional testing was consistent with normal wear and tear of a high time governor. The condition indicators would be N2 rpm droop during collective application and N2 rpm increase during collective reduction.

The data gathered from the three functional bench tests of the accident power turbine governor was inserted into a computer program analytical propulsion system model where it was combined with the 500N rotor system dynamics. A second analytical model was run on a governor at the maximum acceptable serviceable limit. Testing involved an aggressive and rapid collective increase from 25 degrees to 90 degrees within 1 second. The maximum accepted power turbine governor resulted in a 3 percent rotor droop and the San Jose PD governor resulted in a 3.2 percent rotor droop, showing a variance of 0.2 percent greater than the norm. The second maneuver performed was a collective dump from 95 degrees to 5 degrees over a 4 second time period. The maximum accepted governor resulted in a 1.2 percent N2 overspeed and the San Jose PD governor resulted in a 1.7 percent overspeed, showing a variance of 0.5 percent. The complete governor investigation report is appended to this file.

1.18 ADDITIONAL INFORMATION

According to three operator maintenance technicians, the current maintenance manual available at the time of the accident, revision 18, provided inadequate information to properly remove, install, and rig the thruster cables. According to the maintenance technicians, improper installation technique can induce a bending stress on the housing. As an example, if the particle separator/mist eliminator were not removed during forward cable installation, the cable and the telescoping sleeve ball swivel coupling would require bending forces to align with routing holes.

MDHI reported that they reviewed the maintenance manual instructions with technicians at the factory and in the field who have performed thruster cable replacements. These technicians did not have difficulty with the maintenance manual instructions concerning removal, installation, or rigging of the cables. Some technicians stated they did remove the particle separator/mist eliminator to provide better access for cable replacement; however, they did not believe that removal of those components was required to prevent damage to the thruster cables. In addition, a MDHI Field Service Representative witnessed the removal, installation, and rigging of the forward and center thruster cable and confirmed the installation of the thruster cable with the particle separator installed would not damage the cable.

During the investigation and at the request of the FAA, MDHI issued mandatory Service Bulletins. The first, SB500N-021 SB600N-028 was dated November 19, 1999, and was entitled "Forward and Center Thruster Control Cables Conduit Cap at Telescopic Swivel End Inspection. The second was SB500N-020R1 SB600N-027R1, which was dated November 24, 1999, and was entitled "Forward and Center Thruster Cables, Conduit Cap Relief Area, Inspection for stress corrosion cracking." The FAA issued Emergency Priority Letter Airworthiness Directive 99-25-08 to mandate compliance.

At the direction of the owner, the wreckage was released to the insurance company representative on June 6, 2000.

FINDINGS

1. Radar data showed the flight to the destination airport was uneventful until 11 seconds after the helicopter completed a left turn to a downwind heading and began a normal descent for landing.

2. Eleven seconds after the turn's completion, a spike is observed in the Mode C altitude readout. Large sideslip angles will induce static system anomalies. The yaw control system has sufficient authority to induce large and prolonged sideslip angles in cruise flight.

3. The pilot's first "mayday" call over the ATC frequency is coincident with the spike in the Mode C altitude readout.

4. Ground witnesses observed the helicopter begin a yaw to the right, then reverse direction and enter into a rapid rotation as it descended to the ground.

5. According to metallurgical examination, the thruster cable telescoping sleeve ball swivel coupling fitting stress corrosion failure and separation preceded this flight by some length of time.

6. The Teflon inner cable was severely abraded and had been operating out of its sleeve for some length of time. The roughness on the abraded area was in proximity to the sharp edges of the cracked telescoping sleeve ball swivel coupling fitting.

7. Two days before the accident, this pilot experienced a yaw control anomaly and he made a precautionary landing at another airport.

8. Over a 2-day period, maintenance technician(s) examined the thruster control system at the precautionary landing location and could not determine the reason for the discrepancy. During the detailed examination of the cable runs and control freedom checks, the maintenance technician did not remove the access panel over the FS113 splitter assembly (the location of the failed telescoping sleeve ball swivel coupling) to fully examine the thruster control cable run.

9. The pilot and the maintenance technician incorrectly attributed the yaw control anomaly to a YSAS failure, and a joint decision was made between the pilot and the maintenance technician to ferry the helicopter with the unresolved discrepancy to the maintenance base for further diagnostic work.

10. At zero or very small sideslip conditions, the nature of the vertical stabilizer design produces a tension load in the thruster control cable, which would pull the forward cable telescoping sleeve ball swivel coupling fitting together; at larger right sideslip angles (left yaw), the tension load in the thruster cable decreases and eventually becomes a compressive load.

11. Movement of the anti-torque control pedals changes the fan blade pitch to produce more airflow to the circulation control slots and the anti-torque thruster nozzle. As the pedals are displaced from center toward either extreme of travel, the airflow increases proportionally.

12. The RFM procedures for a stuck thruster condition are incomplete and do not contain procedures to minimize airflow to the thruster nozzle.

13. The pilot was concurrently flying the UH-60 helicopter for the California Army National Guard. The emergency procedures for an anti-torque failure in the MD500N are diametrically opposed to those for the UH-60. The pilot successfully completed a MDHI 500N NOTAR Pilot's Recurrent Flight Training Course conducted at the Mesa factory on May 11, 1999.

14. MDHI and Cablecraft failed to expeditiously identify the stress corrosion cracking problem in the telescoping sleeve ball swivel coupling fitting, and failed to change the component specifications in a timely way to prevent the failure and separation of the fitting.

15. The FAA concurred with MDHI that the cable cracking was not a safety of flight issue and the company was given until January 2000 to solve the cracking problem. The FAA concurred with MDHI engineers that the stress corrosion cracking in the thruster cable fittings was considered a minor service difficulty and was not determined to be an unsafe condition, which would warrant the issuance of an airworthiness action.

16. Witness marks and color transfers establish that the blue main rotor blade flexed downward during the ground impact sequence, severing the tail boom. The center thruster cable was impacted by the blue main rotor blade and forcibly separated at the aft bellcranks and the quick disconnect fitting. The post impact position of the thruster can was in an over travel condition.

17. There are no physical limit stops on the travel of the thruster can and it can rotate independently of the thruster control cable and bellcranks.

18. Sound spectrum analysis shows the main rotor transmission was operating at 100 percent until after the Mode C altitude readout spike, and then it rapidly decayed during the descent to ground impact.

19. The twist grip throttle control and the fuel control indices were found at idle. The emergency procedures for an anti-torque failure in the HH-60 helicopter require the pilot to immediately reduce the power to idle and enter an autorotation.

20. The right seat passenger had occipital lobe injuries corresponding to damage on the right entrance door.