Felda, FL, USA
UNREG
INNOVATOR TECHNOLOGIES MOSQUITO XEL
After about a 10-minute flight, the pilot returned to the airport and attempted to land the helicopter. A witness reported that, when the helicopter was about 30 ft above the ground, it seemed to be "unstable" and began to oscillate from side to side. The pilot then aborted the landing, flew for a few minutes, and then returned for landing. The witness reported that during this landing attempt, when the helicopter was again about 30 ft above the ground, it began spinning to the left and then impacted the ground. According to a friend of the pilot, about 2 months before the accident, the pilot had experienced a similar loss of control in the accident helicopter in which the helicopter rapidly spun to the left three times just before landing. The pilot then gained altitude, regained control, and flew away from the landing site. The pilot checked the controls and then came in and landed without incident. The pilot told his friend that, after the helicopter started spinning left, he shut off the automatic throttle governor and was able to recover. When the pilot was interviewed about 1.5 years after the accident, he reported that, during the previous flight when he lost helicopter control, he believed that he came in a little too fast and, since the governor did not work well and the helicopter did not have a lot of horsepower to correct or recover, that could have caused the spins. Regarding the accident, he stated that "I would like to think it was the helicopter, but it could have been me." The majority of the helicopter, including the engine governor and engine management system, was consumed by a postcrash fire and could not be examined. Examination of the surviving components did not reveal any evidence of a preexisting failure or malfunction of the flight control system or engine. Although the loss of control was consistent with the pilot failing to maintain control during the landing approach and experiencing a loss of tail rotor effectiveness, the postcrash fire damage precluded determination of whether a mechanical failure played a role in the loss of control.
HISTORY OF FLIGHT On September 29, 2013, at 1109 eastern daylight time, an unregistered Mosquito XEL Helicopter was substantially damaged during landing at Lazy Springs Recreation Park, Felda, Florida. The non-certificated pilot was seriously injured. Visual meteorological conditions prevailed, and no flight plan was filed for the local personal flight, which was operated under the provisions of Title 14 Code of Federal Regulations (CFR) Part 91. According to a witness, the pilot liked to fly the helicopter whenever he had a chance. On the day of the accident, the pilot had trailered the helicopter into the park to do some flying in the local area. The takeoff was uneventful, but instead of the pilot going out, and flying around the area for 30 to 45 minutes which was his usual habit, approximately 10 minutes later he returned and attempted to land. As the helicopter was approximately 30 feet from touchdown it seemed to be "unstable" and began to oscillate from side to side. The pilot then aborted the landing and flew off for a few minutes, and then returned. This time as the helicopter was once again about 30 feet from touchdown, it began spinning to the left and impacted the ground. A postcrash fire then ensued. The pilot was pulled out of the wreckage by the witness and another person, and was later airlifted to a hospital. PERSONNEL INFORMATION The pilot did not hold any type of pilot certificate or rating for rotor wing aircraft. He had attended a basic helicopter orientation course which consisted of 24 hours of ground instruction and 10 hours of flight instruction in a Schweizer 300C which was designed to familiarize the course attendees with safety procedures, guidelines, aerodynamic forces, forces in flight, flight control systems, safety of flight, hazards of helicopter flight, basic navigation, aviation physiology, federal aviation regulations, aeronautical decision making, and pilot judgment. Review of pilot records also revealed that he had received instruction prior to the course in a Robinson R22, and that he had received 14 CFR Part 61, Special Federal Aviation Regulation Number 73 (SFAR 73) required ground training which required that before a pilot could manipulate the flight controls of a Robinson R22 or R44 Helicopter, they must be trained in energy management, low rotor rpm which could lead to a low rotor rpm stall, and low or negative G, which could lead to mast bumping. Further review of pilot records also indicated that he had received approximately 20 hours of dual instruction and at the time of the accident, he had accrued approximately 40 total hours of flight time. AIRCRAFT INFORMATION The helicopter was of conventional composite and metal construction. The airframe was made of fiberglass in a vinylester matrix. It was powered by a 60 horsepower, two cycle, two cylinder engine, equipped with a 180-watt alternator which provided power to run the helicopters electrical system. The drive train's primary reduction was bolted directly to the engine. A centrifugal clutch on the engine crankshaft permitted startup of the engine without a load from the rotor system. Power was transmitted from the clutch to the driven pulley of the reduction through a cogged belt. The driven pulley housed a sprag clutch which would permit the rotor to overspeed the engine during autorotation. Review of the helicopter manufacturer's records revealed that the helicopter was manufactured in 2012 and had been equipped with floats. It weighed 314 pounds which would allow it to be operated under 14 CFR Part 103 ultralight regulations however, the pilot had changed the configuration of the helicopter by removing the floats, and adding an engine governor which rendered it ineligible for operation under Part 103 and placed it into the experimental category. This would have required the pilot to possess a private pilot certificate, the helicopter to be registered with the Federal Aviation Administration (FAA), and an airworthiness inspection to be performed by an FAA designated airworthiness representative prior to the first flight, as described in FAA Advisory Circular (AC) 20-27F, "Certification and Operation of Amateur Built Aircraft." At the time of the accident the helicopter and engine had accrued approximately 20 hours of total operating time. METEOROLOGICAL INFORMATION The recorded weather at Southwest Florida International Airport (RSW), located approximately 17 nautical miles west of the accident site, at 1053, included: winds from 070 degrees at 10 knots, 10 miles visibility, sky clear, temperature 28 degrees C, dew point 21 degrees C, and an altimeter setting of 29.98 inches of mercury. WRECKAGE AND IMPACT INFORMATION Examination of the accident site and wreckage revealed that the helicopter came to rest on a 15- degree embankment on the edge of a 27 acre lake, on a magnetic heading of 095 degrees. The majority of the helicopter including the cabin, seat, floor panel, and tail boom sections were consumed by the postcrash fire. The rotor head showed marks consistent with mast bumping. The control mechanism was connected and moved freely. The swash plate was consumed by post-crash fire. Rotor blade "A" was delaminated and thermal damaged from the blade root to 5 feet outboard. The rotor blade was still connected to the rotor hub. There was no chord or span wise scratching on the blade. The pitch change rod was connected and the pitch change horn was bent about 15- degrees upward. The spindle moved freely. Rotor blade" B" was consumed by post-crash fire, delaminated, and was separated from the spar 13 inches outboard the blade root. There was no chord or span wise scratching. The blade root was still connected to the spindle, which moved freely. The pitch horn was bent about 45- degrees upward and the pitch rod connector was fractured in a manner consistent with tension overload. The No. 3 sprocket was connected to the secondary drive system. The coupler connecting the lower shaft to the splitter gear box was consumed by post-crash fire. There was drive belt residue on the main rotor No. 3 and No.4 sprockets. Engine continuity was not verified due to thermal damage to the engine accessories and the main engine casing. Three motor mounts were present, with the fourth motor mount retaining bolt having been sheared off. Two motor mount retaining bolts on the torque side of the engine were also bent. Both engine carburetors were consumed by the post-crash fire, and were unrecognizable. The muffler was attached to the engine and was unremarkable. The primary drive belt was attached to the No. 1 and No. 2 sprockets, and was thermal damaged. Control continuity from the flight control pedals to the tail rotor pitch links was verified. Control continuity from the cyclic and collective control was not verified due to consumption of the mechanisms from the post-crash fire. The tail rotor blades were connected to their respective pitch links, and were moved freely through their range of travel. The tail rotor gear box, pitch links, and control rods were thermal damaged. The tail rotor blades were free of chord or span wise scratching. The splitter gear box was thermal damaged and the jaw couplers were unremarkable. The dampener in between the couplers was consumed by the post-crash fire. From the splitter gear box to six feet aft of the splitter gear box, the tail rotor drive shaft was either melted or thermal damaged. The three internal carrier bearings were present and thermal damaged. From the tail rotor gear box to a point located 22 ½ inches forward, a fracture of the tail rotor drive shaft, consistent with bending overload and thermal damage was present. The right landing skid was thermal damaged, and otherwise unremarkable. The left landing skid was thermal damaged on the rear left side. The forward cross bow was thermal damaged but remained connected to the "T" fittings. The forward cross cable was thermal damaged and connected to the cross bow. The rear cross bow was thermal damaged and fractured 14 inches upward from the left side of the "T" fitting. The fracture was consistent with bending overload. The rear cross cable was connected to the left side and disconnected on the right side due to thermal damage. The left, right, front, and rear cables, were thermal damaged, and connected to the rear cross bow. Both cables were disconnected from their respective front mounts as a result of the post-crash fire. SURVIVAL FACTORS INFORMATION The occupant restraint system consisted of a lap belt only. Examination of the restraint system revealed that the lap belt was latched. No shoulder harness or anti-dive strap was installed. Review of the seat design also indicated that it was not of an energy absorbing design. TESTS AND RESEARCH During an interview with a friend of the pilot the friend advised that approximately 2 months prior to the accident flight, he observed the pilot have a loss of control with the helicopter similar to what happened on the accident flight when after lifting off and flying around for about 15 minutes he came back in to land. At approximately 30 feet above ground level, the helicopter rapidly spun three times to the left. The pilot then gained altitude, regained control, and flew away from the landing site for a few minutes. The pilot then checked the controls and then came in and landed without incident. When the pilot's friend asked him what happened, the pilot advised him that after the helicopter started spinning to the left, he shut off the automatic throttle governor and was able to recover. Engine Governor and Engine Management System Examination the engine governor and engine management system could not be accomplished as the majority of the system had been consumed by the postcrash fire. Review of information provided by the manufacturer of the engine governor and engine management system revealed that it was composed of a servo module, a system control module, and a twist grip throttle module. The servo module mounted on the airframe of the helicopter and the twist grip and control modules had been mounted on the collective control arm. The servo module contained several sub-assemblies: - A non-conformally coated circuit board with a micro controller, servo driver, magnetic clutch driver, servo limit switches, and a set of dual inline package (DIP) switches for setting operational characteristics. - A servo motor. - A linear screw drive. - A magnetic clutch During operation the micro controller would measure the engine rpm and compute an "Error" signal based on the difference between the actual engine rpm, and the "Set Point" rpm. The Error signal would be sent to the servo driver circuitry to drive the servo motor in the proper direction for pulling or releasing a Bowden cable connected to the twist grip throttle control. When the system was first engaged by turning on the governor switch on the control module, the engine rpm and position of the magnetic clutch magnet are unknown. The microcontroller would then detect the rpm at which the engine was turning, and then would store that rpm in memory as the Set Point rpm. It would then drive the servo motor until the electromagnetic clutch magnet reached the outer end of the drive screw and contacted the clutch armature. The microcontroller then would turn on the clutch driver causing electrical current to flow through the windings of the electromagnet and engage the clutch. Once the clutch was engaged, the servo motor would retract the clutch and the throttle cable until a slight increase in rpm was detected. The increase in rpm would then "tell" the micro controller, that the servo module now had control of the throttle. As the load on the engine varied, the servo module would attempt to hold the rpm constant by changing the throttle setting to vary the engine power. The system would work in conjunction with the helicopter's collective pitch correlator to hold the engine rpm constant as the pitch control was raised and lowered. Because of rotor blade and drive train inertia, the correlator would act before the fact and would increase or decrease the engine power before the engine actually "felt" the load change. The governor would act after the fact, depending on a change in engine rpm to sense load change. Both would work together to achieve rpm control. The governor would work by retracting and releasing the Bowden cable connecting the servo module to the engine carburetors via the twist grip module. The servo module would pull on the throttle cable against the springs in the two carburetors. It could only release the cable. It could not push the throttles closed. Normally, the pull against the carburetor springs needed to hold the throttles open was about 9 pounds. This was adequate to move the mechanisms in the throttle system, including the twist grip and servo modules provided the system moved smoothly and was free of any drag or snag points. If the mechanism did not move freely over the full range of travel, had points were drag increased, the system could allow the rpm to excessively exceed or significantly fall below the set point. According to the manufacturer, for optimal performance of the engine speed governor it was imperative that: - The engine was tuned to produce rated power in the operational environment where it was being used. - The correlator was adjusted for maximum throttle change for collective control movement. - The throttle mechanism was set up for smooth, drag and snag free operation. DIP Switches There were 4 groups of switches that needed to be set for the engine speed controller to function properly with the engine in the helicopter: - Gain/Sensitivity switches. - Engine Tachometer switches. - Engine Responsiveness switches. - Factory Setting switches. The gain/sensitivity switches determined how sensitive the system was to errors between the actual engine rpm and the set point rpm. The engine speed controller would constantly compare that value to the real-time rpm to adjust the throttle appropriately. If the sensitivity was too high the system would constantly make throttle adjustments. The engine tachometer switches were used to set the range of pulses per minute that would be received from the engine and used to establish the real-time engine rpm. The system would get this information by capturing the trigger pulses from the number one capacitor discharge ignition (CDI). The CDIs were triggered 2 times per crankshaft revolution. At 6,000 rpm this equated to 12,000 pulses per minute. The engine tachometer switches were set to use 13,000 pulses per minute as the high end of the expected operational range. The engine responsiveness switches were used to set up the function that dampened out fluctuations in the error curve. This ensured that large error differences caused for example by a rapid pull up on the collective control would be responded to by a large, fast, throttle adjustment, whereas, small load increases, and adjustments to rpm as the engine approached the setpoint rpm would be smaller and less rapid. Together, the gain/sensitivity switches, and the engine responsiveness switches were used to optimize, or tune, the system to minimize hunting, that is overshooting while trying to settle on the exact rpm while also minimizing undershooting (failing to adjust engine rpm close to the setpoint rpm). All settings were based on 1/8th inch of slack in the Bowden cable. More slack would cause the system to surge or lose speed on set. Loss of Tail Rotor Effectiveness Review of the accident flight and the flight two months earlier revealed that both flights also displayed evidence of a loss of tail rotor effectiveness (LTE) which is a critical low-speed aerodynamic flight characteristic that is not related to a maintenance malfunction and is not necessarily the result of a control margin deficiency. It is characterized by an uncommanded, rapid yaw towards the advancing blade which does not subside of its own accord. It can result in the loss of the helicopter if left unchecked. It is caused by an aerodynamic interaction between the main rotor and tail rotor. Some helicopter types are a
The pilot’s loss of helicopter control while hovering for reasons that could not be determined during postaccident examination of the helicopter, which was limited due to postcrash fire damage.
Source: NTSB Aviation Accident Database
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