Italy, TX, USA
N525TA
BELL 525
The experimental research and development helicopter was undergoing developmental flight tests before type certification. On the day of the accident, the helicopter test crew was performing a series of one-engine-inoperative (OEI) tests at increasing airspeeds with a heavy, forward center-of-gravity configuration. (For the OEI tests, the pilots used OEI special training mode software to reduce the power of both engines to a level that simulated the loss of one engine.) The crew initiated the final planned OEI test at a speed of 185 knots. After the crew engaged OEI special training mode, rotor rotation speed (Nr) decayed from 100% to about 91%, and the crew began lowering the collective to stop Nr decay and increase Nr to 103% (the target Nr for recovery). About 5.5 seconds into the test, the crew stopped lowering the collective, and Nr only recovered to about 92%. About 6 to 7 seconds into the test, the helicopter began vibrating at a frequency of 6 hertz (Hz). The vibration was evident in both rotor systems, the airframe, the pilot seats, and the control inputs; the vertical vibration amplitude at the pilot seat peaked about 3 G. (G is a unit of measurement of acceleration; 1 G is equivalent to the acceleration caused by the earth's gravity [about 32.2 ft/sec2].) Nr remained between about 90% and 92% until about 12 to 13 seconds into the test, then began fluctuating consistent with collective control inputs; subsequent collective control input increases led to further decay in Nr. Nr decayed to about 80% as the collective was raised, and the main rotor blades began to flap out of plane. About 21 seconds into the test, the main rotor blades flapped low enough to impact the tail boom, severing it and causing the in-flight breakup of the helicopter. The main rotor, tail rotor, flight controls, powerplants, and rotor drive systems exhibited no evidence of preexisting malfunction before the vibrations began. The structural wreckage did not exhibit evidence that the oscillations themselves resulted in a structural failure leading to the in-flight breakup. Examination of the wreckage revealed no indications that the helicopter had been improperly maintained. Helicopter Performance After Stop in Nr Recovery During previous OEI tests, the crew lowered the collective input to near or below 50% to allow Nr to recover. As airspeed increased during each test, the crew took longer to recover Nr to 103%. (At 102 knots, recovery time was 3.4 seconds, and at 175 knots, recovery time was 13 seconds.) However, after initiating the final OEI test at 185 knots, the crew only lowered the collective to 58% and subsequently only recovered Nr to 92%. While at 92%, the main rotor scissors mode was excited. (The main rotor scissors mode occurs when the lead-lag motions of the blades act in such a way that adjacent blades move together and apart in a scissoring motion. See the factual report for more information about the scissors mode.) The main rotor scissors mode excitation resulted in the 6-Hz airframe vertical vibration, which was transmitted to the crew seats and created a biomechanical feedback loop through the pilot-held collective control. A second feedback system, driven by the attitude and heading reference system (AHRS) inputs to the main rotor swashplate, also continued to drive the scissors mode and its resultant 6-Hz airframe vibration. Biomechanical Feedback Biomechanical feedback in the aircraft design industry refers to unintentional control inputs resulting from involuntary pilot limb motions caused by vehicle accelerations. The gain between the vertical acceleration and 6-Hz collective stick movement can be calculated by dividing stick movement by vertical acceleration. (If no biomechanical feedback existed, there would be no gain [0 inch per G].) During the accident, the collective stick moved, on average, 0.2 inch per every G of seat acceleration. The "nonzero" relationship between the control stick amplitude and the seat vibration illustrates that biomechanical feedback contributed to the helicopter's vibration. Further, a positive value of pilot gain occurred near 6 Hz, which indicates instability in the system (meaning that any input to the system will amplify as opposed to dampen). Thus, biomechanical feedback contributed to increases in vibration amplitude during this accident. Although the helicopter manufacturer's design process for biomechanical feedback included software filters in the cyclic control laws to reduce certain types of oscillatory cyclic control inputs by the pilot, no filter was designed for the collective. Thus, the 6-Hz oscillatory collective inputs by the pilot were not filtered. As a result, a control feedback loop began when the pilot-held collective stick commanded an oscillatory collective pitch input (about 6 Hz) into the main rotor, increasing the 6-Hz vibration, which in turn increased the magnitude of the oscillatory (6-Hz) collective pitch input. In addition, the gain between the pilot movement and the collective control stick movement in the vertical axis was never tested on a shake table before the accident. For the cyclic control, lateral vibration was introduced on a shake table. This test was conducted specifically for the helicopter model's side-stick cyclic since the manufacturer expected a different transfer function from that of a traditional cyclic. For the collective control, no such test was conducted despite this being the first helicopter with a side-stick collective control. While it is possible that the decision to not shake test in the vertical axis to inform the pilot model could have been influenced by schedule pressure, interviews did not suggest that decisions would have been different given the lack of anticipation of scissors mode and resulting aerodynamic effect. Attitude and Heading Reference System The AHRS is designed to detect uncommanded accelerations (such as the helicopter's reaction to a gust of wind) and reduce their effects by automatically providing corrective inputs to the main rotor swashplate. The AHRS detected and responded to the 6-Hz airframe vertical vibration in a manner that sustained the main rotor scissors mode and its resultant 6-Hz vibration. Specifically, analysis of the telemetry data revealed that the AHRS responded to the 6-Hz vibration with inputs to the main rotor swashplate analogous to a "cyclic stir" (when the cyclic control stick is moved in a stirring motion). The helicopter manufacturer's assessment of the AHRS-induced cyclic stir swashplate motion was that it would exacerbate the main rotor scissors mode. The AHRS is intended to respond primarily to lower-frequency uncommanded accelerations. Because the helicopter manufacturer did not predict an excitement of the scissors mode in the accident test flight conditions, the filter design of the AHRS allowed it to respond to the 6-Hz airframe vibration. Thus, the AHRS detected and attempted to attenuate the 6-Hz airframe vertical vibration, but its response instead exacerbated the main rotor scissors mode and its resultant 6-Hz vibration, closing the AHRS feedback loop. Reasons for Crew Stop in Nr Recovery Investigators explored possible reasons why the crewmembers stopped their recovery from the initial Nr droop, including a reaction to an abnormal condition on the helicopter, distraction from the recovery task, or a conservative response due to the high airspeed. Telemetry data does not indicate the existence of an abnormal condition when the crewmembers stopped their recovery. In addition, the chase helicopter crewmembers reported seeing no distractions or abnormalities outside of the helicopter (for example, birds). Therefore, investigators focused on the crew's increasingly conservative response as the airspeed increased during the tests. During the previous OEI tests, as airspeed increased, the crew recoveries took more time to reach 103% Nr and collective response became less pronounced. During postaccident interviews, helicopter manufacturer test pilots indicated that they interpreted this trend as the tendency of the crew to be more judicious while applying collective at successively higher airspeeds to avoid recovering too fast and overspeeding the rotor or damaging the transmission. Thus, the crew may have been more conservative during recovery at the helicopter's high speed during the final test. The chief test pilot also stated that if Nr had stabilized, the pilot would not have been in a rush and was possibly initiating a slow recovery. As an experimental research and development helicopter configured to carry two pilots and with no passenger seating, the accident helicopter was not required to be equipped with either a flight data recorder (FDR) or cockpit voice recorder (CVR) under the provisions of 14 Code of Federal Regulations (CFR) 91.609. (When certified as a transport-category rotorcraft under 14 CFR Part 29, the helicopter model will be equipped with both CVR and FDR recording capabilities.) A combination CVR and FDR (CVFDR) was installed in the flight test helicopter but was not operational at the time of the accident. Although investigators were able to examine and analyze telemetry data, a properly functioning CVFDR would have recorded any discussions between the accident pilots that could have offered more information about potential abnormal conditions, distractions, or reasons for their stop in recovery after initiation of the OEI test. Additionally, cockpit image recording capability would have recorded any pilot actions and interactions with the aircraft systems including avionics button presses, warning acknowledgements, and any other physical response to the aircraft. Cockpit audio and imagery could have provided insight into when the crewmembers first felt or detected the 6-Hz vibration, how they may have verbalized their assessment of an observed anomaly, and whether they attempted any specific corrective action because of the vibration. Thus, the lack of cockpit audio or image data precluded access to data needed to fully determine why the crew may have momentarily stopped the collective pitch reduction to recover Nr and any corrective actions the crew may have attempted as a result of the 6-Hz vibration. Regardless of why the crew stopped recovery of Nr at 92%, other helicopter test pilots suggested in postaccident interviews that continuous flight in the 92% to 93% Nr range was not abnormal for an OEI maneuver (in this model helicopter and another model in the helicopter manufacturer's test program). This is further supported by another model in the helicopter manufacturer's test program during which extended flight occurred in the low 90% Nr range. (The other helicopter model did not encounter any unusual behavior [rotor mode/vibration] during the test points with the extended recovery time, and the pilots did not receive negative feedback on the recovery time.) The lack of any negative feedback on extended flight in the low 90% Nr range may have reinforced that flight through that range was appropriate. On the pilot displays (specifically, the power situation indicator [PSI]) in the accident helicopter model, 90% to 100% Nr is depicted as a green range or arc. The decision to fly continuously in the 92% to 93% Nr range is consistent with typical pilot association of green arcs with flight regimes that are appropriate for continuous flight. The company's flight technology specialist stated that the colors (green arc) presented on the PSI were a precedent taken from the other helicopter model test program, which suggests that it was likely not reevaluated for appropriateness given the accident helicopter's operating limitations. In addition, flight testing was only conducted for continuous flight at 103% and 100% Nr with all engines operative; however, no testing of Nr continuously between 90% to 100% while in an OEI condition was conducted. Extended flight in the low 90% Nr range during previous testing of another helicopter model and the depiction of 90% to 100% Nr in a green arc on the PSI may have contributed to the pilots' decision to stop in the 92% range during the recovery from the OEI maneuver, which resulted in the onset and increase of the 6-Hz vibration. Crew Response to Low Nr and Vibration Interviews with the helicopter manufacturer test pilots and engineers suggest that there were two ways for the pilots to exit the low Nr and, correspondingly, the vibration condition: (1) lower/reduce the collective to increase Nr or (2) exit OEI training mode, which would increase power available from the engines. About 1.5 to 2 seconds passed between the stop at 58% collective and the onset of the vibration. Had the pilots recovered Nr to 100%, it is possible that the main rotor scissors mode would have subsided and the airframe vibrations would have dampened. Lowering the Collective One option for recovering from the low Nr and vibration condition was to lower the collective to increase Nr. The investigation could not determine if the pilots' fluctuating collective inputs were deliberate when the 6-Hz vibration was dominant. Because the crew needed to be aware of low Nr to respond appropriately, investigators considered the available visual, aural, and tactile cues regarding Nr in the vibration environment. The visual cues available to the crew included the crew alerting system (CAS) text "ROTOR RPM LO," PSI numeric display, warning flag, warning push button annunciator (PBA), and the change of the PSI Nr display from a bar to an arc. The CAS text, warning flag, and warning PBA would have been flashing until acknowledged by the crew. Because the telemetry did not record crew button presses, it is not possible to know if the crew acknowledged these alerts. Studies indicate that visual acuity is negatively affected by vertical vibration, particularly in the 5- to 7-Hz frequency range (Lewis and Griffin 1980a; Lewis and Griffin 1980b). Results indicated that reading speed and accuracy degraded for amplitudes as low as 0.1 G (McLeod and Griffin 1989; Griffin and Hayward 1994). Further studies show that visual performance decreases with increasing vibration amplitude (Shoenberger 1972; Griffin 1975; Griffin 2012). The vertical vibration amplitude at the pilot seat rose above 1 G from 10 seconds into the test until the end of the test, with peaks as high as about 3 G. Given the sensitivity of the human body to vibration frequencies near 6 Hz and the extreme amplitude of the vibration environment, the displays were likely unreadable to the crew (although the colors of the warning text, flag, and PBA may have been discernable). In addition, the change of the Nr display on the PSI from bar to arc may have been recognizable; however, reading of the needle would likely not have been possible in the vibration environment. Thus, the crew was likely unable to read visual information that provided specific low Nr information, although they may have had a generalized cue that Nr was low. Aural cues available to the crew regarding low Nr included the master warning annunciation and the sound of decreasing Nr. The master warning aural tone would have annunciated at 12.5 seconds and 16.8 seconds (continuing until acknowledged by the crew). However, this tone was associated with at least 21 other warning messages and was not unique to the "ROTOR RPM LO" message despite a technical standard that requires that low Nr have a unique tone associated with it. The master aural tone annunciating continuously was chosen for test flight because audio files had not yet been developed; the helicopter manufacturer pilots and test team had decided that some aural annunciation of low Nr would be enough to proceed with test flights but that the distinct tone for low Nr was not immediately needed for flight test. Aural cues can be used for redundancy if visual information is unavailable. The accident pilots were aware that a unique tone did not exist for low Nr; however, they likely were not able to retrieve unambiguous visual information to confirm the warning, outside of a change in shape of the rpm display. Had a unique aural warning tone been implemented in the helicopter, it could have provided a salient, unambiguous cue to the crewmem
HISTORY OF FLIGHTOn July 6, 2016, about 1148 central daylight time, an experimental research and development Bell 525 helicopter, N525TA, broke up in flight and impacted terrain near Italy, Texas. The two test pilots received fatal injuries, and the helicopter was destroyed. The helicopter, which was owned by Bell Helicopter Textron, Inc., was being operated under the provisions of 14 Code of Federal Regulations (CFR) Part 91 as a developmental flight test. Visual meteorological conditions prevailed at the time of the accident. The flight originated from Arlington Municipal Airport, Arlington, Texas. About 0630 on the morning of the accident, the two test pilots, flight test engineers, and a chase helicopter flight crew briefed the planned flight. The brief detailed that the accident helicopter, accompanied by a chase helicopter, would proceed to the Arlington Initial Experimental Test Area (about 30 miles south of Arlington Municipal Airport) to perform the in-flight portion of the tests. The purpose of the flight was to evaluate engine loads at maximum continuous power, two-to-one-engine simulated engine failures, longitudinal roll oscillations, and run-on landings in the heavy, forward center-of-gravity configuration. The test card for the two-to-one-engine simulated engine failure detailed that the pilots would simulate the loss of engine power from one engine while keeping both engines operating by using one-engine-inoperative (OEI) special training mode software, which reduced the power output of both engines to represent the maximum power that can be produced by one engine. When the OEI special training mode was engaged and a loss of power was simulated, the pilot would monitor rotor rotation speed (Nr) and intentionally delay his response by about 1 second before recovering from the maneuver by lowering the collective to reduce the power demanded by the rotor (and increase Nr). The lowest allowable Nr limit was identified as 86%; if Nr went below 86%, the test would be halted, and the crew would recover Nr to 103%, exit OEI special training mode, and return to steady level flight. A Bell structural engineer stated that flight below 86% Nr would result in the helicopter returning to base. During test flights, flight test engineers monitor real-time telemetry data from the helicopter under the oversight of the flight test director, who was in direct radio communications with both the test helicopter pilots and the chase helicopter pilots. About 0959, weather conditions were determined to be acceptable for the flight, and about 1038, the helicopter departed for the test area, followed by the chase helicopter. About 1048, the pilots established the helicopter's maximum level flight airspeed (Vh) at 4,000 ft density altitude (DA) as 148 knots calibrated airspeed (KCAS). After performing steady-heading sideslips, the pilots performed a series of level turns and then began the two-to-one-engine simulated engine failures. About 1108, the pilots set the OEI training mode shaft horsepower to a value predetermined by the flight engineers. The first three tests were performed in level flight at 102 KCAS, 131 KCAS, and 145 KCAS. The pilots then performed tests at 155 knots true airspeed (KTAS), 160 KTAS, 165 KTAS, and 175 KTAS, which required the helicopter to be in a shallow descent to achieve the required airspeed. These OEI tests had resulted in a rotor speed decay of 5 to 13% Nr. During these tests, to allow Nr to recover to 97% or greater, the crew lowered the collective input to near or below 50%. (100% is the full-up collective position, and 0% is the full-down collective position.) Data recorded on the helicopter's flight test recorder system, which was typically downloaded after each test flight and also transmitted via a telemetry stream to Bell's flight-test facility for real-time analysis and recording, indicate the build-up tests and recovery time required (see table 1). (Record 45 was a void record, and record 49 was aborted because of two engine torque spikes typical of wind gust encounters.) Table 1. Build-up tests and recovery time required. During the build up to the final test, the flight test engineers received warning and alert notifications, most of which related to main rotor and tail rotor pitch link loads, pylon loads, and tail boom loads. These alerts and warnings were expected as the airspeed increased and the dynamic loads on the rotor system and airframe also increased. During most of the OEI transitions, the pilot responded by lowering the collective between 1 and 2 seconds after the simulated loss of engine power. However, with each increase in airspeed, the time the crew took to recover Nr to the target value of 103% was longer. Bell test pilots indicated that they interpreted this trend as the tendency of the crew to be more judicious while applying collective at successively higher airspeeds in order to avoid recovering too fast and overspeeding the rotor or damaging the transmission. About 1148, the final test was performed at 185 KTAS, which was the helicopter's never-to-exceed speed (Vne) at the time of the test flight; the set up and entry were the same as the previous tests. OEI was engaged, and Nr drooped to about 91% within 1.5 seconds. The Nr decay was stopped by the pilot's reduction of collective, and Nr began to recover and leveled out around 92%. The crew stopped lowering the collective at the 58% collective stick position. About 7 seconds after arresting the Nr decay (about 12 seconds into the test), the structural dynamics engineer noticed increased engine vibrations, at which point he called "knock-it-off." The test director radioed to the Bell 525 pilots to "knock-it-off," while other engineers in the telemetry room were receiving warnings and alerts and were reinforcing the "knock-it-off" call. The crew of the chase helicopter, which was positioned about 100 ft above and on the right side of the Bell 525 about 3 to 4 rotor diameters away, heard the test director call "knock-it-off" about the same time they observed the 525's rotor blades flying high and the rotor looking wobbly and slow. The chase helicopter crew radioed, "Hey, you're flapping pretty good," but the 525 pilots did not respond. About 21 seconds into the test, the main rotor severed the tail boom, and the telemetry signal was lost. The chase helicopter crew observed the helicopter's tail and fuselage jack-knife and debris separate from the helicopter. The chase helicopter crew radioed to the test director, "We've had a major accident," and landed near the wreckage to attempt assistance. PERSONNEL INFORMATIONThe pilot held a letter of authorization (LOA) from the Federal Aviation Administration (FAA) dated December 2, 2015, authorizing him to act as pilot-in-command (PIC) of the Bell experimental helicopter designated model 525. He completed crew resource management (CRM) training on January 12, 2015. The pilot graduated from the United States Naval Test Pilot School (USNTPS) in 2010. He then worked on numerous flight test projects involving the Bell AH-1W (SuperCobra, a twin-engine attack helicopter) and UH-1Y (Venom/Super Huey, a twin-engine utility helicopter). On September 23, 2013, he was hired by the Bell Helicopter flight test department as a pilot for the Bell 525 program. The copilot held an LOA from the FAA dated December 2, 2015, authorizing him to act as PIC of the Bell experimental helicopter designated model 525. He completed CRM training on January 12, 2015. The copilot completed US Navy flight training in 2000 and graduated from the USNTPS in 2006. He then worked on numerous AH-1W and UH-1Y test programs. On August 2, 2010, he was hired by the Bell Helicopter flight test department as a pilot for the Bell 525 program. AIRCRAFT INFORMATIONThe accident helicopter was a conventional main rotor and tail rotor design (see figure 1). On April 25, 2016, the helicopter received its latest experimental research and development airworthiness certificate from the FAA. The helicopter was a manufacturing prototype being developed for certification as a transport-category helicopter in compliance with 14 CFR Part 29. As part of the airworthiness certificate, the FAA issued an operating limitations document (also dated April 25, 2016) that specified the following: pilots operating the helicopter must hold a temporary LOA issued by an FAA flight standards operations inspector to act as PIC, the helicopter must be maintained by an FAA-approved inspection program, day visual flight rules flight operations are authorized, and all flights must be conducted within the Arlington Initial Experimental Test Area. The helicopter was estimated to weigh about 19,975 lbs at the time of the accident. Figure 1. Accident helicopter (Bell 525, N525TA). Source: Bell Helicopter The Bell 525 helicopter had a five-bladed main rotor that provided helicopter lift and thrust and rotated in a counterclockwise direction when viewed from above. The main rotor was a fully articulated system that used elastomeric bearings to accommodate blade feathering, flapping, and lead-lag motions. Fluid-elastic dampers moderated lead-lag motion of the blades. The five main rotor blades were identified by colored stickers, presented in order of advancing rotation (when seated in the pilot seat and observing the blades pass from right to left): blue, orange, red, green, and white. The Bell 525 also had a four-bladed, fully articulated, canted tail rotor that provided thrust to counteract main rotor torque effect, control helicopter yaw, and provide lift. The four tail rotor blades were identified by colored stickers, presented in order of advancing rotation: blue, orange, red, and green. The helicopter was equipped with two General Electric (GE) CT7-2F1 turboshaft engines, mounted aft of the main transmission, and one Honeywell RE100BR auxiliary power unit (APU), mounted between the two engines at the aft end of the engine deck. The helicopter was equipped with a triple-redundant fly-by-wire flight control system with a triplex hydraulic system. Additionally, the helicopter was equipped with retractable tricycle landing gear. The cockpit was configured for two pilots in a side-by-side seated position and a center console between them. Each pilot had a cyclic side-stick controller forward of the seat's right armrest, a collective side-stick controller immediately forward of the seat's left arm rest, and a set of pedals forward of their feet. The instrument panel consisted of four identical primary flight display (PFD)/multifunction display (MFD) panels. The center console had two Garmin Touch Control (GTC) panels, the landing gear handle, the Nav/Com panel, and the flight test switch panel, which included some controls for the OEI special training mode software. Directly above the GTCs were the engine control COSIF (crank, off, start, idle, fly) knobs. Each pilot had an additional pilot display unit that provided real-time flight test instrumentation parameters such as DA, boom airspeed, mast airspeed, engine torque, load factor, pitch/yaw/roll rates, slip angle, and main rotor and tail rotor flapping angles. OEI Training Mode OEI training mode is a specific GE software-driven capability that permits simulation of a single-engine failure without actually rolling back or shutting down an engine in flight. When the flight crew engages the OEI training mode, both engines reduce power to represent the power available from a single engine. Consistent with normal operations and depending on the flight conditions, if the power demanded by the rotor exceeds the power available, Nr will droop. If single-engine power is insufficient to sustain the forward speed, the pilot must reduce the power demand by lowering the collective control, applying aft cyclic (to reduce speed), or using a combination of both. Nr increases to 103% when the power required matches the single-engine power available. To engage OEI training, the pilot or copilot navigates to the OEI training page on the GTC and selects the engine to fail on the touch screen. Once selected, a green bar appears on the failed engine button to signal that OEI training mode was engaged (see figure 2). When OEI training mode is engaged, the pilot's side (right-seat) PFD displays simulated OEI engine values, and the copilot's side (left-seat) PFD displays the actual all-engines-operative (AEO) data. Figure 2. OEI training page on the GTC. Source: Bell Helicopter The OEI special training mode that Bell used for the accident flight test did not incorporate an automatic disengagement of OEI training mode for low Nr. Bell modified the production version of OEI training mode software, originally created by GE, to eliminate a safeguard that automatically exited the OEI training mode when Nr fell below 90%. According to Bell, automatic disengagement at 90% Nr was not low enough to allow development and demonstration of OEI recovery across the flight envelope during testing, and a lower Nr value for automatic disengagement was deemed unnecessary due to the highly controlled test environment. To manually exit OEI training mode, the pilot could (1) press the engine fail button on the GTC (the same button used to engage OEI training mode), (2) exit the OEI training page on the GTC (using the BACK button), or (3) move the COSIF switch to a position other than "Fly" and then return the switch to "Fly." The Bell 525 lead test pilot indicated in a postaccident interview that the options to exit OEI training mode were not discussed formally with all the test pilots but were specifically discussed with the accident test pilot. Bell 525 test pilots interviewed said that they almost always press the engine fail button on the GTC to exit OEI training mode; some Bell pilots were aware of the other methods to exit OEI training mode while other test pilots were not. Disengaging OEI training mode would make both engines available to provide full power to restore the reference Nr to 100% if the rotor was in a drooped state. The production OEI training mode, which will be used in Bell 525 production helicopters, includes an automatic disengagement of OEI training if Nr decays below 90% (pending validation via testing). In the production OEI training mode, automatic exit would occur in the following circumstances: -Loss of an engine. -Torques of the two engines are not within ~30 ft-lb of each other. -There are any significant engine failures (any fault that would cause local channel degraded on any of the 4 channels). If the enable bit for training is set (bit 20) AND both engine request bits are set (bit 21 and 22). To engage training only one-engine request bit can be set. -Power turbine speed (Np) is 5% below the reference value (having previously been within 1% of the reference while in training) or to a value below 90%. -Np is above 106%. -Real engine gas producer turbine speed is above 106%. -Real engine measured gas temperature is above 1934.3° F/ 1056.8° C. -Real single-engine torque is above 521 ft-lb (67.7%). -Real engine oil temperature is above 148.89° C. -Low oil pressure switch is tripped. OEI training mode flight test risk analysis worksheets documented planned operational risk mitigation for OEI training. A worksheet approved on June 29, 2015, included a discussion of the risk of low Nr, and a worksheet approved on April 1, 2016, included a discussion about engine overtorquing. Power Situation Indicator (PSI) The PSI was located in the bottom left corner of the PFD for each pilot. The bars in the bottom right corner of the PSI represented Np for the number 1 engine, Nr, and Np for the number 2 engine, respectively. The arc in the center of the display depicted the percentage of engine value compared to its limit (see figure 3). Figure 3. Example of PSI on the Bell 525. Source: Bell Helicopter Indications of Low Rotor Rpm in the Bell 525 Power Situation Indicator The PSI displayed Nr as a vertical scale (center bar in lower right indicator) when Nr was above 90%, as shown in figure 3. If Nr dropped below 90%, the display changed to an analog needle that displayed a green arc for Nr between 100 and 90%, a yellow arc for Nr between
A severe vibration of the helicopter that led to the crew's inability to maintain sufficient rotor rotation speed (Nr), leading to excessive main rotor blade flapping, subsequent main rotor blade contact with the tail boom, and the resultant in-flight breakup. Contributing to the severity and sustainment of the vibration, which was not predicted during development, were (1) the collective biomechanical feedback and (2) the attitude and heading reference system response, both of which occurred due to the lack of protections in the flight control laws against the sustainment and growth of adverse feedback loops when the 6-hertz airframe vibration initiated. Contributing to the crew's inability to maintain sufficient Nr in the severe vibration environment were (1) the lack of an automated safeguard in the modified one-engine-inoperative software used during flight testing to exit at a critical Nr threshold and (2) the lack of distinct and unambiguous cues for low Nr.
Source: NTSB Aviation Accident Database
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