Safety summary

What happened

On the evening of 21 October 2016, a Eurocopter BK 117 C-2 helicopter, registered VH-SYB, departed Crookwell Medical helicopter landing site, New South Wales. The crew were returning to their home base at Orange, New South Wales, after conducting an emergency medical service (EMS) task. The flight was conducted as a night visual imaging system (NVIS) operation under night visual flight rules (NVFR), with the pilot and aircrew member (ACM) both wearing night vision goggles (NVG).

Shortly after take-off, the helicopter unexpectedly encountered low cloud, and the pilot initiated the operator’s inadvertent entry into instrument meteorological conditions (IMC) procedure. As the momentum of the helicopter’s climb reduced, the pilot lowered the helicopter’s nose to regain airspeed, but she inadvertently over corrected the pitch angle to 15° nose-down, as well as allowing a slight roll to the left. The resulting unusual attitude triggered a caution alert from the helicopter’s enhanced ground proximity warning system.

What the ATSB found

The ATSB found that the pilot had undertaken relevant recent training in inadvertent IMC recovery, and the pitch over correction was probably (at least in part) associated with the surprising nature of the event. In addition, during a high workload situation, the pilot was probably distracted by the reflection of the helicopter’s red anti-collision light reflecting off nearby cloud while wearing NVG.

In response to the distraction, the pilot asked the ACM to switch the light off. However, the ACM was not familiar (or required to be familiar) with the operation of the light switch, and inadvertently switched on the strobe light, which exposed the pilot to bright white light reflecting off cloud while wearing NVG. Exactly when the strobe light was switched on, and whether it contributed to the unusual attitude, could not be determined.

Earlier that evening, the pilot had diverted to Crookwell during a flight from Canberra to Orange due to the presence of thunderstorms and reduced visibility en route. The flight from Canberra to Orange was conducted under NVFR with NVIS, when the use of instrument flight rules (IFR) was practical and involved less risk. The ATSB identified that although the operator’s policies stated that EMS flights should be conducted under IFR where practical, this policy was not reinforced in the manual that covered NVIS operations. An IFR departure was not available for the take-off from Crookwell.

What’s been done as a result

As a result of this occurrence, the helicopter operator undertook several proactive safety actions, including clarifying its flight planning policy on IFR and NVIS operations and enhancing its training and advisory materials. The operator is also assessing the potential fitment of flight data monitoring equipment to all of its fleet.

Safety message

Although NVIS/NVG can significantly improve the quality and quantity of visual information available to pilots at night, the use of such devices also involves risk in some situations. This occurrence highlights the importance of ensuring that operators and pilots have robust processes for deciding when to conduct NVIS operations. It also serves as an example of the limitations and risks of NVIS operations when there are external light sources or reflections, and highlights the benefit of having a predetermined strategy for responding to degraded visibility conditions.

The occurrence

Overview

On the evening of 21 October 2016, a Eurocopter BK 117 C-2 helicopter, registered VH-SYB and operated by CHC Helicopter Australia, departed Crookwell Medical helicopter landing site, New South Wales (NSW). The crew were returning to their home base at Orange, NSW, after conducting an emergency medical service (EMS) task. The flight was conducted as a night visual imaging system (NVIS^[1]) operation under night visual flight rules (NVFR), with the pilot and aircrew member both wearing night vision goggles (NVG).

Shortly after take-off the helicopter unexpectedly encountered low cloud. During the recovery from inadvertent instrument meteorological conditions^[2] (IMC), the helicopter entered an unusual attitude, which triggered an enhanced ground proximity warning system (EGPWS) alert.

Previous flights to Young and Canberra

Earlier that day, at about 1627 Eastern Daylight-saving Time (EDT), the rostered crew accepted a task to transport a patient from Young, NSW to Canberra Hospital, Australian Capital Territory. The rostered day shift crew consisted of a pilot and aircrew member (ACM) and the medical team consisted of a doctor and a paramedic.

As the day shift pilot was completing the flight planning, the night shift pilot arrived at the operator’s base at Orange. Following a discussion, the night shift pilot agreed to take the task. The day shift pilot briefed the night shift pilot on the helicopter, personnel, forecast weather and flight plan.

The area forecast (ARFOR) for area 21 (which included Orange, Young and Canberra)^[3] available at that time was valid from 1600 to 0400. For flights between Orange, Young and Canberra, the forecast included:

• scattered showers and isolated thunderstorms north of Young, with isolated showers in other areas after 2000
• rain developing after 2200
• broken^[4] stratus cloud from 2,000-5,000 ft above mean sea level (AMSL) after 2000 and broken cloud in precipitation^[5]
• visibilities of 2,000 m in thunderstorms, 3,000 m in rain, and 4,000 m in showers.

The flights from Orange to Young and Young to Canberra were expected to be completed well before to 2000. The aerodrome forecasts (TAFs) for Young and Canberra for the relevant period indicated the potential for showers but did not indicate any notable problems with visibility or low cloud.

The helicopter departed Orange at 1650 and landed at Young at 1728. While on the ground at Young, the pilot checked for updated weather information using the OzRunways^[6] program on a tablet computer^[7] provided by the operator. At that time the 1600 ARFOR was still valid, and the TAF for Canberra was effectively the same.

After the doctor and paramedic boarded the patient, the helicopter departed Young at 1818 and landed at Canberra Hospital at 1855. After the doctor, paramedic and patient disembarked at the hospital, the pilot and ACM departed Canberra Hospital to position to the operator’s base near Canberra Airport to refuel the helicopter. The helicopter landed at the base at 1918.

The flights between Orange, Young, Canberra Hospital and the operator’s Canberra base (Figure 1) were planned and conducted under day visual flight rules (VFR) in visual meteorological conditions.^[8] No weather-related difficulties were reported to have occurred during those flights.

Figure 1: VH-SYB’s approximate flight path for outbound flights from Orange to Canberra (in blue) and return flights (in yellow), including diversion flight paths (yellow dots)

Source: Google maps annotated by ATSB

Preparation for the return flight from Canberra to Orange

During the late afternoon, the Bureau of Meteorology (BoM) issued a significant meteorological information (SIGMET)^[9] valid from 1850 to 2150. It provided information about thunderstorms in a squall line with hail north-west of Orange moving in an east, south-easterly direction at 35 kt.

An updated ARFOR for area 21, valid from 1900 to 1000, indicated some changes in the forecast weather conditions along the planned route from Canberra to Orange. Key details included:

• frequent thunderstorms in a squall line
• isolated thunderstorms with possible hail and scattered showers north of Young, with isolated showers in other areas
• rain developing
• broken stratus cloud from 2,000-5,000 ft AMSL after 2000 and broken cloud in precipitation
• visibilities of 2,000 m in thunderstorms, 3,000 m in rain, and 4,000 m in showers.

BoM subsequently issued an amended ARFOR valid from 1925 to 1000, which provided essentially the same information as the 1900 ARFOR.

An amended TAF for Orange was issued at 1729, but this forecast was basically the same as the previous (1200) forecast, indicating showers from 2000 and the potential for thunderstorms from 1900. The forecast also required the pilot to provide for a suitable alternate aerodrome when flight planning.^[10]

While at the operator’s Canberra base, the pilot planned the return flight to Orange. She recalled that she checked for updated weather information using the OzRunways program on the tablet computer, and the ARFOR information she saw on the computer was unchanged from that which she saw prior to departing Young. The pilot reported that she was not aware of the updated 1900 ARFOR or the amended 1925 ARFOR, and did not recall any statement about frequent thunderstorms in a squall line. She noted that, had she seen that statement, she would not have departed Canberra.

The pilot planned the return flight to Orange under NVFR as an NVIS operation. She reported that conducting the flight under NVFR rather than instrument flight rules (IFR) as it provided more diversion options along the planned route.^[11] The pilot uploaded the maximum fuel possible to cover inflight contingencies, and this included sufficient fuel to conduct an approach at Orange and, if required, divert to a suitable alternate aerodrome such as Bathurst.

At 1947, the helicopter departed from the operator’s Canberra base to Canberra Hospital to pick up the doctor and paramedic, landing at the hospital at 1953. While on the ground at the hospital, the paramedic viewed BoM weather radar information on his tablet computer and noted there was a line of thunderstorms extending from Moree (north of Orange) to Young. At about this time, the ACM received a text message from another ACM at Orange that stated there was a storm overhead Orange at that time.

The paramedic expressed concern about the situation and suggested they stay in Canberra, and the ACM agreed. The pilot recalled noting that the storms forecast for later that day appeared to have arrived earlier than expected, and that they would probably have passed through Orange by the time they arrived. She also believed the storms could be traversed safely.

The pilot advised the ACM and medical team that she had checked the weather forecast and, even after reviewing the BoM weather radar information, was happy to proceed with the flight to Orange. After a more detailed explanation of the flight planning from the pilot, including the en route alternatives and planning for contingencies, further discussion took place and the ACM and medical team agreed to conduct the flight.

Return flight from Canberra and diversion to Crookwell

The helicopter departed Canberra Hospital at 2010 for the return flight to Orange, with both the pilot and ACM wearing NVG. Shortly after departure, the pilot asked the paramedic if he still had any concerns. The paramedic recalled stating that he did, however, he would discuss it in the debrief at the end of the flight. The ACM subsequently reported to the ATSB that she also was not entirely comfortable conducting the flight due to the storms but did not voice those concerns again at this time.

While en route, the pilot and ACM continually assessed the movement of storms using the helicopter’s weather radar and ground-based weather radar using a tablet computer. The paramedic used his tablet computer to view the BoM weather radar information.

Approximately 20 NM north-west of Crookwell the pilot determined that the weather conditions were no longer suitable for continuing the flight due to closing gaps between the storms and reduced visibility ahead. After some discussion between the pilot, ACM and paramedic regarding suitable diversion sites, it was decided to divert to the Crookwell Medical helicopter landing site (HLS) (Figure 1).

At about 2100, the pilot conducted a visual approach (using NVG) to the Crookwell Medical HLS. After landing, the pilot, ACM and medical team went inside a rural fire service building to wait until the approaching storm cells passed.

Preparation for the flight from Crookwell to Orange

The pilot reported that, while on the ground at Crookwell, there was some rain but no storms passed directly overhead. The pilot monitored the weather situation using the OzRunways program and ground-based weather radar. The paramedic viewed the BoM weather radar information on his tablet computer.

The pilot later recalled that, again, the latest ARFOR displayed on the operator’s tablet computer was unchanged from that which she obtained prior to departing Young. She reported that she obtained the latest weather observations for Orange (including temperature and dew point) and Bathurst, and both indicated appropriate conditions.^[12] No recorded weather observations were available at Crookwell.

After being on the ground for some time at Crookwell, the pilot noted that the storms appeared to have passed through. She went outside to check the weather conditions, and she conducted a visual scan and then a second scan with NVG to assess the conditions. She recalled seeing lightning flashes in the distance, which lit up the local area, showing no cloud, a small amount of rain and no haze.

During initial discussions about departing Crookwell, the paramedic stated that he was not comfortable that the storms had passed so did not think they should depart. After waiting a short period and having further discussions, the pilot, ACM and medical team agreed and planned for a departure under NVFR using NVIS at about 2240. The pilot, ACM and paramedic recalled that, when walking out to the helicopter, there was light rain but no lightning or other indication of thunderstorms or low cloud observed nearby.

Departure from Crookwell

Both the pilot and the ACM were wearing NVG for the departure from Crookwell. The pilot conducted a standard, hand-flown back-up take-off procedure to a decision point of approximately 120 ft. She then transitioned the helicopter to a forward climbing profile, increasing airspeed towards the take-off safety speed^[13] (45 kt) with a positive rate of climb.

The ACM recalled that, during the initial part of the take-off, she had poor visibility due to a combination of effects from the helicopter’s searchlight and rain on the windscreen. She stated to the pilot that she had poor visibility, which the pilot noted.

The pilot recalled that the building floodlights and rain limited her vision ahead, but she had good visibility on the right-hand side, which she stated to the ACM. The pilot expected the visibility problems ahead would improve as the helicopter moved away from the building lights and the rain streamed off the windscreen with increased forward speed. Accordingly, she continued the departure.

The pilot reported that when the helicopter reached the take-off safety speed, she adjusted the helicopter attitude but the searchlight reflecting off the rain limited her forward visibility. She adjusted the searchlight down and to the right so she could see the ground more clearly, but the forward visibility did not improve enough for her to be comfortable with continuing the flight in that configuration. She then adjusted the helicopter’s attitude to slow the helicopter and help maintain visibility with the ground.

From this point, there was a rapid series of events. The pilot recalled that a flash of the anti-collision light reflecting off cloud alerted her to the presence of low cloud above. She immediately assessed that returning to land was no longer an option and called ‘inadvertent IMC’ while initiating the operator’s inadvertent IMC recovery procedure. She transitioned to flying on instruments, and commenced a rapid, vertical climb at maximum power, announcing that she was climbing to the other occupants. With the next flash of the anti-collision light, they had entered cloud.

In an attempt to reduce the distracting effect of the anti-collision light, the pilot asked the ACM to turn the light off. The ACM removed her NVG, took a few seconds to locate the switch and, not realising the switch had three positions, inadvertently moved the anti-collision/strobe light switch from the ACOL (anti-collision light) position to the ON (anti-collision and strobe light) position. This changed the reflected light from red to bright white, which the pilot subsequently reported had a blinding effect on her vision while wearing NVG. The pilot again asked the ACM to turn the light off, which the ACM then actioned.

The pilot recalled that, during this period, the momentum of the helicopter’s climb was decreasing and the airspeed was reducing, so she lowered the helicopter’s nose to increase the airspeed. She recalled that she intended to lower the pitch angle by 5° but inadvertently overcorrected to about 15° nose-down, as well as inadvertently allowing a slight roll to the left. She also noted that during this period, the flashing of the reflected external lights had been distracting, and her transition to effective instrument scanning had not been ‘tidy’.

Very soon after the helicopter entered the unusual attitude, the enhanced ground proximity warning system (EGPWS) provided an aural ‘caution terrain’ alert. The pilot called ‘climbing, climbing, climbing’ and commenced unusual attitude recovery actions (in accordance with the operator’s procedure) by adjusting the helicopter’s attitude to wings-level and initiating a climb, with the helicopter still at full power. She recalled that the transition from the unusual attitude to the recovery attitude was conducted promptly and smoothly.

The ACM recalled that, following the EGPWS alert, she looked out the front windscreen and saw terrain. She also called out ‘climb, climb, climb’. During the climb, she noted that the airspeed had reduced close to 20 kt and stated the airspeed was low to the pilot. The pilot replied that she was happy to accept low airspeed at that stage in order to clear the terrain. Soon after, she adjusted the pitch attitude to increase airspeed.

Although the ACM’s recollection of many of the events during this period was similar to that of the pilot (above), there were some notable differences. More specifically:

• When the pilot slowed the helicopter down to maintain visual reference with the ground, the ACM recalled the helicopter also descended. However, the pilot did not recall any descent at that time.
• During the initial climb as part of the inadvertent IMC recovery, the ACM recalled that the pilot asked her to turn on the helicopter’s weather radar, which resulted in the ACM looking inside the helicopter. The pilot could not recall asking the ACM to turn on the weather radar until later (see Continuation to Orange and diversion to Bathurst).
• The ACM recalled that while she was looking inside the helicopter, she observed the vertical speed indicator (VSI) indicating a rate of descent of about 3,000 ft/minute, and it was at this time the EGPWS alert occurred. The pilot recalled seeing a figure of about 100 ft/minute at the time of the EGPWS alert (see also Helicopter performance).
• The ACM recalled that the pilot first asked her to turn off the anti-collision light during the climb after the unusual attitude event and EGPWS alert occurred. The pilot and the paramedic both recalled that the initial request to turn off the anti-collision light occurred during the earlier climb as part of the inadvertent IMC recovery (and before the unusual attitude event). The initial statements made by the pilot and paramedic soon after the occurrence also indicated that the inadvertent selection of the strobe light on occurred before the unusual attitude event. However, the pilot subsequently could not recall whether the strobe light was selected on before or after the unusual attitude event.

The pilot estimated that the unusual attitude event occurred between 200 and 400 ft above ground level (AGL). The ACM estimated that it occurred at about 500 ft AGL, and that she believed the helicopter could not have descended below 200 ft AGL as she did not recall the altitude alerter (normally set at 200 ft AGL during a departure) annunciating.

Continuation to Orange and diversion to Bathurst

Following the unusual attitude recovery, the pilot engaged the autopilot as soon as possible (after the helicopter reached the required minimum airspeed limitation) and continued climbing at best rate of climb towards the area’s lowest safe altitude for IFR flight (6,100 ft). She recalled asking the ACM to turn on the weather radar at this time, and she asked the ACM to change the destination on the helicopter’s GPS to Orange and obtain the latest weather information for Orange. The pilot contacted air traffic control (ATC) at 2246 and advised they had inadvertently entered IMC and requested to change flight rules from NVFR to IFR. ATC approved the request.

While tracking to Orange, the crew obtained weather reports that indicated the weather conditions at Orange were unsuitable for landing due to low cloud (with overcast cloud at 400 ft), whereas the weather at Bathurst was suitable. The pilot confirmed there was sufficient fuel to conduct an instrument approach at Orange and then divert to Bathurst, and elected to proceed to Orange.

When the helicopter was closer to Orange, the crew obtained further weather reports, which indicated the conditions at the airport had not improved (still overcast at 400 ft) but the conditions at Bathurst were still suitable.^[14] Consequently, at 2311 the pilot requested clearance from ATC to divert to Bathurst, which was provided. The pilot diverted to Bathurst, conducted an instrument approach and landed at Bathurst Airport at 2331.

__________

1. NVIS: a self-contained binocular night vision enhancement device that is (a) helmet mounted or otherwise worn by a person, and (b) can detect and amplify light in both the visual and near infra-red bands of the electromagnetic spectrum. The devices are usually googles, which are also known as night vision goggles (NVG).
2. Instrument meteorological conditions (IMC): weather conditions that require pilots to fly primarily by reference to instruments, and therefore under IFR, rather than by outside visual reference. Typically, this means flying in cloud or limited visibility.
3. Area forecast (ARFOR): routine forecasts for designated areas and amendments when prescribed criteria were satisfied. An ARFOR provided a forecast of weather conditions for the specified area from the surface to 10,000 ft above mean sea level (AMSL). The standard validity period was 12 hours but this could vary from state to state. They were normally issued about 1 hour prior to the start of their validity period.
4. Cloud cover: in aviation, cloud cover is reported using words that denote the extent of the cover. ‘Sky clear’ (SKC) indicates no cloud, ‘few’ indicates that 1-2 oktas (or eighths) is covered, ‘scattered’ indicates 3-4 oktas is covered, ‘broken’ indicates 5-7 oktas is covered, and ‘overcast’ indicates that 8 oktas is covered.
5. The terrain along most of the route between Canberra and Orange was 2,000-3,000 ft AMSL. For example, elevations included 1,886 ft at Canberra Airport, 3,112 ft at Orange Airport and 2,840 ft at Crookwell Medical HLS.
6. OzRunways is an electronic flight bag application that provides weather, area briefings, and other flight planning information in Australia. It includes a weather manager that provides a visual representation of weather along a route, as well as significant meteorological information, weather radar and rain prediction.
7. The operator’s tablet computer was used as an electronic flight bag (EFB). An EFB is a portable information system for flight deck crew members which allows storing, updating, delivering, displaying and/or computing digital data to support flight operations or duties. The tablet computer met the requirements for an EFB under Civil Aviation Order 82.0 (Air operator’s certificates – applications for certificates and general requirements).
8. Visual meteorological conditions (VMC): an aviation flight category in which visual flight rules (VFR) flight is permitted – that is, conditions in which pilots have sufficient visibility to fly the aircraft while maintaining visual separation from terrain and other aircraft.
9. Significant meteorological information (SIGMET): a weather advisory service that provides the location, extent, expected movement and change in intensity of potentially hazardous (significant) or extreme meteorological conditions that are dangerous to most aircraft, such as thunderstorms or severe turbulence.
10. The amended TAF for Orange issued at 1729 indicated, for the period from 2000, visibility of at least 10,000 m, showers of rain, scattered cloud at 500 ft and broken cloud at 1,000 ft above ground level. It also indicated that, from 1900, there was a TEMPO with 30 per cent probability of deteriorations of up to 60 minutes due to thunderstorms with rain, as well as visibility of 2,000 m and broken cloud at 500 ft. The 1729 TAF was similar to the previous TAF issued at 1200.
11. If the flight was conducted under IFR, the pilot would have only been able to conduct planned flights to aerodromes that had an instrument approach procedure. Relevant aerodromes near the intended route between Canberra and Orange included Cowra and Bathurst. Other off-route aerodromes included Goulburn and Young.
12. The weather observation for Orange issued at 2130 was a special weather observation report (SPECI), which indicated visibility of at least 10,000 m and overcast cloud at 900 ft. However, SPECIs issued at 2200 and 2230 indicated visibility of at least 10,000 m and overcast cloud at 400 ft. For the period from 2045 to 2330 the SPECI temperature / dew point split was 1° C. The weather observations for Bathurst during this period were routine reports (METARs), which indicated visibility of at least 10,000 m and scattered cloud at 3,600 ft (2130), 2,400 ft (at 2200) and 2,000 ft (at 2230). The pilot obtained current weather reports via the aerodrome weather information service (AWIS), which may have been slightly different to the SPECIs / METARs, which are normally issued every 30 minutes.
13. V_TOSS or take off safety speed: the minimum speed at which climb shall be achieved with the critical engine inoperative, the remaining engine operating within approved operating limits.
14. An amended TAF for Orange was issued at 2326, which included visibility of at least 10,000 m, showers in rain and broken cloud at 500 ft. The 30 per cent probability of periods of up to 60 minutes with visibility 2,000 m, thunderstorms with rain and broken cloud 500 ft was still included. ATC broadcast the amended TAF at 2327.

Context

Personnel information

Pilot

The pilot held an Air Transport Pilot (Helicopter) Licence and Command Instrument Rating. This permitted her to conduct night visual flight rules (NVFR) and instrument flight rules (IFR) flights. She had a total aeronautical experience of about 5,065 hours including 602 hours on BK 117 helicopters. Her last command instrument proficiency check was conducted on 25 May 2016. Her last recurrent proficiency check (base and line check) was conducted on 27 May 2016.

The pilot was qualified for night vision imaging system (NVIS) flight and had last completed an NVIS proficiency check flight on 9 August 2016. This check included conducting recovery from inadvertent instrument meteorological conditions (IMC) exercises while wearing night vision goggles (NVG). No problems were noted on this proficiency check, or her other recent proficiency checks.

The pilot had conducted about 770 hours of night flying including about 300 hours on NVIS. She met the recency requirements to conduct both NVIS flights and IFR flights.

The pilot had a valid Class 1 Aviation Medical Certificate, and she stated that she had no health-related issues.

The pilot’s roster pattern included 2 days standby at work (0730-1730), 2 nights standby at work (1730-0730) and 4 days off. She was rostered on standby at work from 1730 to 0730 the previous night, but had not been required to conduct any flight tasks. On the day of the occurrence, the pilot’s rostered period commenced at 1730. She arrived at work early, for no particular reason, at about 1600. She stated that she did not feel tired prior to or during the occurrence flight.

The pilot reported that she had no specific pressure or reason to return to Orange that night. However, she felt a responsibility to complete the mission and to return the helicopter back to home base if possible and safe to do so, thereby allowing further emergency medical service (EMS) or search and rescue (SAR) tasking from the Orange base (see Operator information).

Aircrew member information

The aircrew member (ACM) had about 859 hours total experience, which included about 465 hours on the BK 117 C-2. She was qualified, met the recency requirements to conduct NVIS operations sitting in a control seat, and had last completed an NVIS capability check flight on 31 May 2016. She had been trained and assessed to provide assistance to pilots in deteriorating visibility conditions and in-flight recovery procedures.

In addition to their role as part of a rescue crew, an ACM’s role during EMS flights could include performing tasks at the request of and under the direct supervision of the pilot. These tasks could include operating the weather radar, radios, the GPS and the searchlight and, if required, calling normal and emergency checklists and monitoring responses.

The operator reported that initial training courses for an ACM provided a general overview of helicopter systems. However, ACMs were not specifically trained in the use of helicopter exterior light switches (such as the anti-collision/strobe light switch), and were not expected to operate such switches during flight. The ACM on the occurrence flight reported that she had not been asked to operate the anti-collision/strobe light switch during flight before.

The pilot reported that she was aware that some other ACMs she had flown with had operated the switch before. However, the pilot had not often flown with this ACM as they were normally on different rosters.

The ACM had the same type of roster pattern as the pilot. She was rostered on standby at work from 0730 to 1730 on the day of the occurrence. Although she had conducted some duty tasks during the day at the Orange base, she also had rest periods to manage potential fatigue. The ACM’s duty time associated with the patient transfer task commenced at the time of the tasking call. As a result, the delayed return flight did not exceed her maximum duty time limit of 14 hours.

The ACM recalled that, while on the ground at Crookwell, she felt tired, so she closed her eyes and rested for a period. The pilot recalled asking the ACM, prior to departing Crookwell, whether she was fit to continue the flight and the ACM agreed. However, in hindsight the ACM recalled that she was fatigued at that time as it had been a long day, and she should have stated she was fatigued and not able to complete the flight.

Aircraft information

General information

The BK 117 C-2 is a medium-sized, single main rotor and tail rotor helicopter with skid-type landing gear. It was fitted with two Turbomeca Arriel 1E2 turbine engines.

VH-SYB was manufactured in 2008. It was equipped for EMS operations and complied with the requirements of Civil Aviation Order (CAO) 82.6 (Night vision imaging system – helicopters). The model of NVG used during the occurrence flight were ITT M949 aviator night vision imaging system 9.

The helicopter was approved for single-pilot operations by day and night under VFR and IFR and was equipped with an autopilot, radio altimeter and weather radar. The helicopter could also be flown with two pilots, with both front seats fitted with appropriate flight controls and displays.

The helicopter was fitted with a Honeywell Mark XXI enhanced ground proximity warning system (EGPWS). The EGPWS was a terrain awareness and warning system (TAWS) designed for helicopters, with additional features. The EGPWS alert described by the occupants was a caution alert rather than a warning alert.

The helicopter did not have a flight data recorder fitted, nor was a recorder required under Civil Aviation Order 20.18 (Aircraft equipment – basic operational requirements).

Anti-collision/strobe light switch

The helicopter’s exterior lighting included navigation (position), anti-collision and strobe lights. The anti-collision light was a red, flashing light, mounted at the top of the vertical fin. It was normally required to be selected as ‘on’ any time the aircraft’s engines were operating. The strobe light was a white flashing light, located near the anti-collision light.

The anti-collision and strobe lights were operated by a toggle switch, located on the overhead panel above the pilot’s seat. The switch had three positions (Figure 2):

• OFF
• ACOL, which operated the helicopter’s red anti-collision light only
• ON, which operated the red anti-collision light and white strobe light.

Figure 2: VH-SYB’s anti-collision / strobe light switch

Source: CHC annotated by ATSB

Flight control display system – Primary flight display

VH-SYB was fitted with a three-screen flight control display system, which included two independent primary flight displays (PFD). Each PFD (Figure 3) provided a pilot with primary flight data including attitude, altitude, airspeed and vertical speed.

The altimeter provided pilots with information as to their height above mean sea level. It was located on the right side of the PFD. In VH-SYB, the barometric corrected altitude was displayed with a numerical value every 100 ft and graduated markers every 20 ft.

The vertical speed indicator (VSI) provided pilots with information as to the helicopter’s rate of climb or descent. It was located immediately to the left of the altimeter and had a digital value and an analog indicator.

The analog indicator’s scale was from -2,000 ft/minute to +2,000 ft/minute with a graduated mark at every 500 ft. The digital indication was displayed as a value between -9,900 ft/minute and +9,900 ft/minute and it provided pilots a rate of climb or descent information even if the analog information was out of range. The displayed figure was in 100 ft/minute units (for example, ’3‘ equalled 300 ft/minute).

The BK 117 C-2 flight manual noted a limitation of the accuracy of the VSI, which stated:

Vertical speed indication may be unreliable in gusty conditions and at airspeeds around 30 kt during fast transition from hover flight into steep descent and from level flight to hover during rapid pull up’s.

Figure 3: VH-SYB’s primary flight display

Source: CHC.

Helicopter performance

As noted in Departure from Crookwell, the ACM recalled seeing the VSI indicate a rate of descent of 3,000 ft/minute at the time of the EGPWS alert. She also recalled seeing a rate of climb of 2,900 ft/minute during the initial unusual attitude recovery. The ACM stated she read these figures from the digital VSI display.

The helicopter had an all engines operating rate of climb of less than 2,200 ft/minute with gross weights above 3,000 kg (the approximate helicopter weight at the time of the occurrence). This maximum rate of climb would be lower if the helicopter had a low airspeed.

A rate of climb of 2,900 ft/minute would be beyond the performance capability of the helicopter, particularly given the low airspeed at the time. In addition, it is unlikely that a helicopter could transition from such a significant rate of descent to a significant rate of climb in a short period of time. Furthermore, the pilot, ACM and paramedic all stated that they did not experience any significant sensory feeling of acceleration forces acting on their bodies that would normally be associated with a helicopter rapidly recovering from a 3,000 ft/minute rate of descent to a 2,900 ft/minute rate of climb in a short period of time. The pilot also advised that the rate of climb and rate of descent figures recalled by the ACM were unrealistic.

The elevation of the Crookwell Medical helicopter landing site (HLS) was 2,840 ft. Therefore, indicated altitudes during the occurrence would have been slightly higher than 2,800 ft. Given the proximity of the altimeter to the VSI on the PFD, it is possible that the values recalled by the ACM during this high workload period may have been read from the altimeter rather than the VSI.

Meteorological information

Details of relevant forecasts and weather reports are provided in The occurrence. Based on the available information, the reported weather observations en route and at Orange were consistent with the issued forecasts.

The operator used tablet computers as an electronic flight bag (EFB) to assist pilots with various flight operations tasks, including obtaining weather information. The pilot had received training by the operator in the use of the EFB in accordance with the operator’s procedures. The crew had two operator-provided tablet computers on board the helicopter to provide redundancy.

Records from the National Aeronautical Information Processing System^[15] indicated that weather information was accessed while the helicopter was on the ground in Young, Canberra and Crookwell. However, such records did not provide information about what forecasts or other weather information was accessed. The pilot reported that she used the EFB OzRunways program during the afternoon prior to the occurrence flight in the same way that she had always used it, and on no other occasions had she had a problem obtaining the latest weather information.

OzRunways advised that its program did not automatically download weather information from the Bureau of Meteorology. To obtain weather information, a pilot had to initiate a request. The OzRunways program presented users with the time of request and period of validity for the received area forecast. Received information was only stored for a maximum of 60 minutes, after which that information was no longer displayed. This feature mitigated the risk that a pilot would reference outdated weather information.

OzRunways advised the ATSB that it had not received any previous reports of a pilot not receiving updated weather information when using its program.

Helicopter landing site information

Crookwell Medical HLS (Figure 4), located to the northern side of the Crookwell Township, and was operated by the Crookwell Rural Fire Service (RFS). The HLS had no published IFR arrival or departure procedures. Its elevation was 2,840 ft and no weather information was available for the site. The concrete helipad was not equipped with standard pad lighting, although general floodlighting was available for night operations upon request to an RFS duty officer. Fuel and rest facilities were available, but there was no sleeping accommodation.

Figure 4: Crookwell Medical helicopter landing site

Source: OzRunways annotated by ATSB

Night vision imaging systems

Under visual meteorological conditions (VMC), NVIS/NVG amplify the amount of light reflected from the terrain. In most conditions and with proper implementation, NVIS/NVG provide pilots with a significant increase in the quantity and quality of visual information compared with unaided night vision. They allow a pilot to see the horizon, objects, terrain and weather more easily. Furthermore, they assist the pilot to maintain spatial orientation, to avoid hazards such as inadvertent entry into IMC, and to visually navigate.

Despite the advantages associated with NVIS/NVG, their application has limitations. Compared with optimal day vision, they are monochromatic, have a limited field of view, and a lower visual acuity. In addition, the quality of the NVG image is variable depending on the operating environment, and NVIS/NVG do not provide adequate imagery under all lighting, scene contrast, and atmospheric conditions. For example, the quality of an NVG image can vary depending on the amount of celestial illumination, the intensity of direct bright light, weather conditions, and the height above the surface and speed of the helicopter.

NVGs are not designed to be used for flight under IFR, however, it is possible to ‘see through’ areas of light moisture when using NVGs which increases the risk of inadvertently entering IMC.

CAO 82.6 (Night vision imaging system – helicopters) provided direction to NVIS operators of matters to be included in their operations manuals, including equipment standards, ongoing maintenance requirements, operating procedures and flight crew capability. Operators considering NVIS operations were required under CAO 82.6 to carry out a risk assessment prior to the commencement of operations or training with these devices.

The operator was approved by the Civil Aviation Safety Authority to conduct NVIS operations in accordance with CAO 82.6. The operator’s training and checking system ensured NVIS qualified pilots and ACMs maintained ongoing competency for NVIS flights.

Civil Aviation Advisory Publication (CAAP) 174-1(1) (Night vision googles – helicopters)^[16] provided guidance to operators and pilots for the conduct of helicopter aerial work operations using NVIS. It discussed general information on NVIS and its capabilities, and limitations including environmental considerations as well as information on known human factors and physiological limitations.

Operator information

Flight planning policy

The operator provided its flight planning policies and procedures in multiple manuals. The Operations Manual Part A (OMA General/Basic Procedures) contained the operator’s VFR/IFR policy, which stated:

The Company policy is to conduct passenger carrying flights under the IFR. Other operations described in the OMF [EMS and SAR] and the OMG [external load and similar special operations] can be performed under the IFR or the VFR, but where practical IFR procedures are to be used.

However, a departure may be flown under the VFR to a predetermined changeover point, and IFR may be cancelled before landing when the commander has visual reference to the terrain (VFR then applies)…

The OMA defined a passenger as ‘any person other than an operating crew member carried on-board an aircraft.’

The Operations Manual Part F (OMF) contained the operator’s policies and procedures for EMS and search and rescue (SAR) operations. It stated a SAR/EMS crew normally consisted of a pilot (commander), ACM (or winch operator), rescue crewman (if a crew member was intended to be lowered by the winch) and other optional crew members required for the task (such as medical personnel).

The OMF further stated that EMS/SAR flights could be planned as IFR or VFR (as long as relevant limitations were met). In addition, it stated that NVIS operations under the VFR were permitted for specific types of flights, including EMS flights and positioning flights for EMS flights. There was no specific statement regarding the planning of return flights after a patient had been transported to a hospital.^[17]

In terms of flight planning for NVIS operations under the NVFR, the OMF stated:

• no cloud was permitted up to 1,000 ft AGL within a 2 NM corridor either side of the planned track if the aircraft was IFR capable and the crew were IFR qualified (otherwise no cloud was permitted up to 2,000 ft)
• minimum visibility of 5,000 m.

The pilot interpreted the operator’s flight planning policies to mean the return flight from Canberra to Orange could be conducted under NVFR using NVIS, as the flight was still being conducted as part of an EMS mission; returning the helicopter and crew back to the base to enable their use for future tasking. The ACM reported that in her experience most, but not all, return flights were conducted under the IFR. The operator noted that its policy was for IFR to be used wherever practical for such flights. It also advised that, following the occurrence, it identified that this policy had not been applied consistently across all of its EMS/SAR bases.

The OMF stated that all NVIS flights were only permitted to be conducted in an IFR serviceable helicopter and crewed by an IFR rated and current crew (unless approved by the manager of flight operations). It also stated that one NVIS-qualified pilot was required for NVIS flights with an en route height at more than 1,000 ft AGL. For other flights, a second NVIS-qualified crew member was required (either a pilot or ACM) to assist the pilot.

Risk management of inadvertent entry into instrument metrological conditions

The OMF stated that pilots were to conduct a risk assessment for all NVIS operations as part of the pre-flight planning and briefing process. To reduce the risk of possible loss of control or controlled flight into terrain when encountering inadvertent IMC, pilots were required to include actions for inadvertent IMC in their departure and approach briefs.

The OMF included a list of specific NVIS hazards and relevant treatments (or risk controls) required by the operator and the crew to eliminate or reduce the risk to an acceptable level. The list of hazards included several related to inadvertent entry into IMC.

With regard to procedures for inadvertent entry to IMC, the operator’s OMA stated:

Loss of visual references can occur suddenly and with little warning in marginal weather conditions. Pilots should contingency plan regarding how to recover from inadvertent IMC... Consideration should be given to the elements of the missed approach briefing on an IFR approach as preparation.

An instinctive reflex to 'go back', descend rapidly or turn sharply to regain visual clues can be hazardous and may add to disorientation, possibly accompanied by loss of control and / or unintentional contact with terrain.

If the pilot has lost visual cues:

a. Follow the procedure… [for] Recovery from unusual attitude, ensuring that a heading is selected clear of known obstacles / terrain

b. Once safe flight is achieved, formulate a new plan

Inadvertent entry into IMC is a dangerous condition, and the immediate safety of the aircraft shall then take precedence over any rules relating to IFR flight plans, clearances, or normal rules for transition from VFR to IFR.

The OMF provided further details relating to loss of visual reference when using NVIS:

Despite careful preparation, the potential for inadvertent IMC penetration always exists. It is important that crews are able to recognise subtle changes to the NVIS image that occur prior to entry into instrument meteorological conditions. A loss of visual reference when conducting NVIS operations may be caused by a reduction in light levels that occur when moisture or haze is in the atmosphere. Visual reference may also be lost when terrain or obstacles obstruct the illumination from the moon or stars. Further, any build-up of moisture such as precipitation, mist, fog or cloud along with haze, smoke and dust may also lead to a loss of visual reference.

In areas where there is little lighting from built up areas, cloud cover will significantly reduce NVIS performance. Under cloud cover there is an increased likelihood of increased moisture that in the reduced light levels may go undetected until a total loss of visual reference has occurred…

When flying below LSALT the PF shall immediately apply climb power, obtain V_TOSS or V_Y^[18] as appropriate and climb to the minimum safe altitude. Depending on the situation prior to the loss of visual reference, manoeuvring over a safe area while climbing may assist in avoiding terrain, however that this will lead to a reduction in climb performance.

In the event of loss of visual reference the PF [pilot flying] shall announce "Losing visual reference – commencing IMC (intentions)".

Use of helicopter external lighting

CAAP 174-1(1) stated that exterior lighting such as an anti-collision lights and searchlights may have adverse effects on NVG. Light reflections off cloud or rain may interfere with NVG performance as the unit cannot adjust well to flickering or intermittent bright light. This causes rapid changes in the NVG image, which may be distracting and disorienting for the pilot.

Prior to NVIS operations, an operator was required to ensure that a helicopter’s exterior (and interior) lights were compatible with its NVG and met relevant standards. CAO 82.6 also stated that an NVIS operator was exempt from the requirements for using external aircraft lighting if complying with the CAO was at variance with the external lighting requirements.

The operator’s procedures stated that the use of NVIS to detect traffic was inconsistent, and that exterior lighting should be used as much as possible. The procedures also stated that:

If the helicopter's exterior lighting adversely affects NVIS performance, the commander must:

a. If he is satisfied there is no risk of collision with another aircraft: Turn off the exterior lighting, or

b. If he is satisfied there is such a risk: Immediately cease NVIS operations.

Goggle and de-goggle procedures

The OMF stated that time was required for each crew member to transition from aided flight (with NVG) to unaided flight (without NVG). It further stated that the crew was ‘not to goggle-up or de-goggle while in the hover’, and, where possible, the helicopter was to be at a safe height (generally above 1,000 ft AGL) before transitioning from (or to) NVGs. Crew members were required to transition one at a time (at the pilot’s direction), with each crew member clearly announcing when they had transitioned.

Crew resource management

The operator’s training and checking system included crew resource management (CRM) training. Pilots and ACMs were required to complete initial CRM training prior to commencing unsupervised line flying. Recurrent CRM training and assessment was then conducted annually. In addition to recurrent training, pilots and ACMs underwent in-depth CRM training during aircraft conversion and (for pilots) command upgrades.

The operator’s CRM program had a training matrix to ensure the structured delivery of major topics was achieved at least every 3 years. Human factors, threat and error management, decision making, leadership, team behaviour and communication and co-ordination inside and outside the cockpit were topics contained in the matrix. The pilot and ACM had both undertaken the operator’s CRM training program.

As previously noted, the operator had a risk management policy and procedure for all crew members involved in NVIS operations to follow prior to conducting such operations. The policy stated:

…all crew members involved in NVIS operations shall participate in the risk management of those operations. NVIS risk management involves recognition of those hazards that NVIS use may create and identifying and applying treatments to eliminate or reduce the risk to an acceptable level…

During operations, whether inflight or during the pre-flight process, crew members who believe that the planned or current NVIS operation is not in accordance with the operations manual or that risks exist that are not covered by the operations manual shall bring them to the attention of the pilot in command.

This policy was referred to by crew members as ‘All to say go, one to say no’ when referring to the decision to conduct a flight.

15. The National Aeronautical Information Processing System (NAIPS) provides a central database of meteorological, NOTAM and chart information. The system is used by Airservices Australia to provide pre-flight and in-flight briefings and to accept and distribute flight notifications.
16. CAAP 174-1(1), issued October 2007, was removed from the CASA website in July 2017. CASA subsequently issued CAAP 174-01 v2.1 in October 2017.
17. A section in the OMF stated that a designated EMS call sign shall be applicable ‘for the full duration of the aircraft mission from call-out until return back to base and back on ‘standby’.’
18. V_y is the best rate of climb speed.

Safety analysis

Introduction

The flight from the Crookwell Medical helicopter landing site (HLS) to Orange was planned as a night vision imaging system (NVIS) operation under night visual flight rules (NVFR) with the pilot and aircrew member (ACM) both using night vision googles (NVG). During the take-off, the helicopter entered cloud, and soon after the helicopter entered an unusual attitude.

As far as could be determined, the unusual attitude occurred and was recovered more than 200 ft above ground level. Nevertheless, it triggered an enhanced ground proximity warning system (EGPWS) terrain alert, and an unusual attitude at a relatively low height and low airspeed has the potential to result in serious consequences.

This analysis first discusses flight-planning aspects, both in terms of the return flight from Canberra to Orange, which diverted to Crookwell, and then from Crookwell, as well as related crew resource management (CRM) aspects. It then discusses factors potentially associated with the unusual attitude, including the operation of the helicopter’s external lighting in cloud.

Flight planning aspects

Flight planning aspects prior to departing Canberra

When the pilot was on the ground in Canberra planning the flight from Canberra to Orange, there had been a significant meteorological information (SIGMET) issued, indicating thunderstorms in a squall line with hail. In addition, the current area forecast (ARFOR), valid from 1925, stated there were frequent thunderstorms in a squall line, as well as isolated thunderstorms with hail. The pilot reported that she checked the weather information using the operator’s tablet computer, but did not see the SIGMET or a statement in the ARFOR about frequent thunderstorms in a squall line. Instead, she recalled that the information in the ARFOR was the same as she had seen prior to departing Orange (which was issued at 1600 and included isolated thunderstorms).

Based on the available information, it appears likely that the latest ARFOR was available each time the pilot accessed weather information on the tablet computer. The time between the access at Young and Canberra was longer than the 60-minute time that the OzRunways program stored previously received information. The reason why the pilot did not notice the SIGMET or the updated ARFOR while on the ground at Canberra (or subsequently at Crookwell) could not be determined.

In addition to thunderstorms, the ARFORs issued at 1600, 1900 and 1925 all indicated broken cloud from 2,000-5,000 ft above mean sea level (AMSL), broken cloud in precipitation and scattered showers. The forecast for Orange also indicated there was showers, scattered cloud at 500 ft above ground level (AGL) and a TEMPO associated with reduced visibility, low cloud and thunderstorms with rain. Overall, these forecasts indicated there was a reasonable potential for encountering cloud and/or reduced visibility during the flight.

The pilot believed that, based on reviewing weather radar information, the forecast storms had passed through the area earlier than expected. She believed that conducting the flight under NVFR with NVIS offered more diversion options along the planned route, and she had ensured that there was sufficient fuel to arrive at Orange and then divert to a suitable aerodrome.

Pilots flying under visual flight rules (VFR) use visual information to avoid weather, obstacles and terrain, and to maintain control of an aircraft’s orientation. The amount of external visual information available to a pilot is significantly degraded during night operations. Although the use of NVIS normally increases the amount of visual cues available, this improvement is limited, particularly when there is a significant amount of cloud and/or precipitation. Overall, conducting the flight under instrument flight rules (IFR) was both possible and practical, and would have minimised the potential risk relative to doing the flight under NVFR.

The pilot reported that she felt no specific pressure to conduct the return flight to Orange, but believed there was an obligation to complete the mission by returning the helicopter to its home base to allow possible further emergency medical service (EMS)/search and rescue (SAR) tasking. However, conducting the flight under IFR would have provided a similar potential for reaching Orange, even though the flight may have had to divert further from the planned flight path.

The pilot also believed that under the operator’s flight planning policy, NVFR with NVIS was a permitted flight planning option for EMS positioning flights. Although the operator’s operations manual (Part A) clearly stated that IFR was to be used ‘where practical’ for all EMS/SAR flights, the procedures for EMS/SAR flights (in Part F) did not reinforce that policy or state a clear preference for when NVFR with NVIS versus IFR should be used. This created potential ambiguity regarding when VFR with NVIS could be used, and the available evidence indicates that other pilots also had not fully understood the operator’s intended policy.

If the flight from Canberra was conducted under the IFR, it is still likely that the pilot would have had to divert during the flight, due to thunderstorms along the planned flight path, with the most likely option being a return to Canberra. However, the decisions prior to departing Canberra were not considered to be contributing factors to the unusual attitude and EGPWS alert that occurred when departing Crookwell. Departures at night from locations such as Crookwell were well within the normal range of operations conducted by the operator.

Flight planning aspects prior to departing Crookwell

Crookwell Medical HLS was not equipped with a published instrument flight rules (IFR) departure procedure and there were no recorded weather observations available. Therefore, when planning the NVFR departure, the pilot had to rely on a visual assessment of the immediate weather conditions in conjunction with the ARFOR and recorded observations for Orange.

Although the thunderstorm hazard had passed, other hazards potentially remained so the pilot assessed the local weather conditions using a visual scan. She also conducted a second scan of the area using NVG. Although light rain was observed, the pilot’s scans of the local area did not detect any potentially associated low cloud.

The decision about whether to depart Crookwell was not straightforward, as departure under the IFR was not available and it was difficult to get detailed information about the weather conditions. In addition, if the crew and medical personnel remained at Crookwell, there were limited rest facilities available, and the helicopter would have been unavailable for further tasking.

Ultimately, the pilot obtained all the information that was available, and based on this information determined the local weather conditions were compliant with the requirements for a NVFR with NVIS departure. The ACM and paramedic’s recollections of the weather conditions at Crookwell were consistent with the pilot’s assessment, and they did not report having any concerns with the conditions prior to departure.

Crew resource management

The effective use of CRM and/or threat and error management will minimise risk in abnormal and emergency situations through identifying and discussing threats and hazards, ensuring all crew have the same understanding of the situation, involving all crew in key decisions and minimising workload.

To ensure CRM will be effective, pilots, ACMs and medical teams need appropriate training and guidance. The operator provided CRM training as well as a documented risk management policy and procedure for all crew members involved in NVIS operations that incorporated CRM aspects.

Prior to departing Canberra, the ACM and paramedic had significant reservations about the approaching line of thunderstorms, and they stated their concerns to the pilot. The pilot briefed the ACM and medical team about the weather conditions, advised she had uploaded sufficient fuel for en route diversions and holding, and that they had multiple diversion options on the route. The pilot had experience flying operations in thunderstorm conditions. It is likely that her confidence flying in similar weather conditions moderated the ACM’s and paramedic’s reservations, and led to them agreeing to conduct the flight even though they both still had concerns. However, with a more forceful relay of concern from either the ACM or medical team, it is likely the pilot would have decided to remain in Canberra.

Prior to departing from Crookwell, the paramedic initially stated to the pilot that he was not comfortable with departing as he was not confident the storms had passed, and the departure was delayed until all the parties were comfortable that the storms had passed. This was an effective use of the operator’s policy, which was described by crew members as ‘All to say go, one to say no’ when referring to the decision to conduct a flight.

The ACM subsequently reported that, in hindsight she was fatigued prior to the flight from Crookwell and should have stated to the pilot that she was fatigued. Based on the available information, it is difficult to make a conclusion about the extent to which the ACM was fatigued, and the extent to which her perception may have been affected by subsequent events.

Undesired attitude and recovery

Avoidance of instrument meteorological conditions (IMC) is more problematic at night because it is difficult to distinguish clear air from cloud in darkness. As a result, loss of already limited visual references can occur suddenly and with little warning in marginal weather conditions. Two main risks associated with flying in limited visibility are:

• spatial disorientation,^[19] leading to loss of control of an aircraft and an uncontrolled flight into terrain
• inability to see and avoid obstacles while remaining under control, potentially leading to controlled flight into terrain.

As VH-SYB was not fitted with a flight data recorder it was not possible to determine the exact sequence of events or flight path of the helicopter during the departure from Crookwell. Based on the available evidence, the helicopter reached the decision point (120 ft) and the pilot transitioned the helicopter to a forward climbing profile, increasing airspeed towards the take-off safety speed (45 kt) with a positive rate of climb.

Shortly after this time, the weather conditions unexpectedly deteriorated and the pilot was unable to continue flight by visual reference using NVG. Soon after, the pilot identified that the helicopter was about to enter cloud, and she promptly commenced the operator’s inadvertent IMC recovery procedure, which involved a rapid, vertical climb at maximum power.

As the momentum of the helicopter’s climb deceased, the pilot intended to lower the pitch angle by 5° but she recalled that she inadvertently overcorrected to about 15° nose-down, as well as inadvertently allowing a slight roll to the left. This subsequently led to the helicopter entering an undesired attitude at low level, which resulted in the EGPWS alert. The pilot effectively identified the undesired attitude and successfully recovered the helicopter to normal flight by following the operator’s unusual attitude recovery procedures. The descent rate during the unusual attitude event could not be reliably determined.

In terms of the reasons for the unusual attitude, there was no known problems with the helicopter’s serviceability and there was no windshear or other environmental factors that directly affected the helicopter’s flight path. The unusual attitude appeared to result from the pilot’s overcorrection of the pitch attitude during her transition from visual flight (with her vision mainly focussed on external visual cues) to instrument flight (focussed mainly on the helicopter’s flight instruments), but the control difficulties did not appear to be related to pilot fatigue, medical or physiological problems. Other potential explanations involved instrument flying proficiency, expectancy and the distraction and other effects of the helicopter’s external lighting when the pilot was using NVG while flying in cloud (see next section).

The pilot met relevant recency requirements for IFR flight, and had conducted an IFR proficiency check 4 months prior to the occurrence and another proficiency check including inadvertent IMC recovery procedures 2 months prior to the occurrence. Research has shown that helicopter pilots with an instrument rating and recent experience are more successful in dealing with an unexpected entry into IMC than pilots who are not rated or do not have recent experience. Nevertheless, they may still encounter difficulties.

For example, one study found that commercial helicopter pilots who did not meet relevant instrument rating proficiency requirements were significantly more likely to lose control (67 per cent) than pilots who did meet the requirements (15 per cent) when unexpectedly entering IMC (Wuerz and O’Neal 1997). Another study showed that instrument-rated helicopter pilots’ performance (including control of pitch and bank) was significantly better in visual metrological conditions (VMC) than after they inadvertently entered IMC, and their performance in dealing with inadvertent entry to IMC in subsequent trials improved with recent practice (Crognale and Krebs 2011).

Other research has shown that airline pilots who are provided with unexpected emergency situations in a simulator do not manage these situations as well as when the emergencies are expected, with increased variability in pilot performance. This effect has been shown to occur with emergencies such as upset recoveries (Landman and others 2017) and stalls (Casner and others 2013), and it occurs with pilots who have successfully managed the same types of emergencies successfully many times before.

Overall, the pilot’s difficulty with controlling the helicopter’s pitch on this occasion was probably (at least in part) associated with the surprising or unexpected nature of the event. No problems were identified with the amount or recency of the proficiency checking undertaken by the pilot in the period prior to the occurrence.

Operation of helicopter’s external lighting

Workload refers to the interaction between an individual and the demands associated with the tasks they are performing. It varies as a function of the number and complexity of task demands and the capacity of the individual to meet those demands. High workload leads to a reduction in the number of information sources an individual will search, and the frequency or amount of time these sources are checked (Staal 2004). It can result in an individual’s performance on some tasks degrading, tasks being performed with simpler or less comprehensive strategies, or tasks being shed completely (Wickens and Hollands 2000).

When using NVG, helicopter external lighting can have various effects on a pilot’s vision and/or attention if the helicopter is near or in cloud. In this case, the helicopter’s red anti-collision light was selected as ‘on’ on prior to departing Crookwell, consistent with the operator’s normal procedures for night operations. The pilot reported that the reflection off cloud of the red flashing anti-collision light was distracting. Accordingly, she wanted the light switched off.

During the inadvertent IMC recovery, the pilot would have been experiencing a high workload and required both hands to control the helicopter. This prevented her capacity to de-goggle and manipulate the light switch. Therefore, it is understandable that the pilot instructed the ACM manipulate the light switching in this situation. However, the operator’s ACMs were not trained or expected by the operator to perform this cockpit function. Accordingly, when under higher workload, it is not surprising that the ACM inadvertently positioned the anti-collision light to the anti-collision and strobe light position. This resulted in a much brighter, white flashing light reflecting off the cloud, and the pilot reported that this had a significant, ‘blinding’ effect on her vision when using NVG.

In addition to impaired vision, exposure to an intense bright light could potentially lead to a range of other short-term performance effects when using NVG, including surprise and potentially a ‘startle’ response. It is therefore critically important to minimise the risk of pilots using NVGs to be exposed to such light, particularly during a critical stage of flight. The sudden onset of the bright light from the strobe light reflecting off cloud would certainly provide a viable explanation for the pilot’s difficulty managing the helicopter’s pitch attitude during the recovery from inadvertent IMC.

However, the extent to which the helicopter’s external lighting contributed to the pilot’s control of the helicopter was difficult to determine due to the inconsistency in the recall of the sequence of events by the pilot, ACM and paramedic. The recollection of the pilot and paramedic was that the pilot asked the ACM to switch off the anti-collision light prior to the unusual attitude event, whereas the ACM recalled that the unusual attitude event occurred first. A person’s memory about a sequence of events during a serious incident or accident can be affected by a range of factors, including the person’s workload, the complexity of the events, the pace at which the events occur and interference from other sequences of events that may occur before and after the sequence of interest. Even if many of the events are recalled, the memory of the sequence in which they occurred may not be accurate (Davis 2001).

Based on the available evidence, the ATSB concluded that the description provided by the pilot and paramedic was viable and more likely to be correct. Nevertheless, although the pilot probably asked the ACM to switch the anti-collision light off prior to the unusual attitude, this action would have taken some time to action and it is unclear whether the ACM inadvertently selected the strobe light on prior to the unusual attitude. Nevertheless, the flashing of the red anti-collision light off cloud, and the process of asking the ACM to select the light off and the ACM commencing this task, provided some level of distraction during a critical time, and probably influenced the pilot’s ability to effectively transition to instrument flight.

If a pre take-off risk assessment identified a realistic potential for entering cloud, then it would be reasonable to expect that the pilot would not undertake a flight using NVG. Alternatively, if some potential risk was identified, then the pilot had the authority to select the external helicopter lights off prior to departure (assuming there is no identified collision risk with other aircraft).

Once the pilot had encountered the unexpected situation, it would have been ideal if the pilot had been able to have the anti-collision light selected as ‘off’ reliably and quickly. However, the ATSB understands that it would not be a reasonable expectation for an operator to ensure that a non-pilot had received training and proficiency checking for such tasks so that they could successfully operate them in a high workload or emergency situation. In addition, there is also the associated risk of either of a front seat occupant de-goggling at a low altitude in such a situation.

It is difficult to estimate the extent of the distraction the anti-collision light provided on this occasion, however the effect would have been minimised if the pilot’s vision was focussed on the helicopter’s flight instruments. Recovery from inadvertent entry into IMC at low altitude is an emergency procedure, and therefore it would generally be appropriate for a pilot to focus on ensuring the helicopter was recovered to a safe height and flight path prior to dealing with distractions wherever possible.

__________

19. Spatial disorientation occurs when a pilot does not correctly sense the position, motion and attitude of an aircraft relative to the surface of the Earth. Further information about spatial disorientation is provided in the ATSB report AO2011-102, VFR flight into dark night involving Aérospatiale AS355F2 VH-NTV, 5 km north of Marree, South Australia, 18 August 2011, available at www.atsb.gov.au.

Findings

From the evidence available, the following findings are made with respect to the terrain awareness warning system alert involving Eurocopter BK 117 C-2, registered VH-SYB, near Crookwell, New South Wales on 21 October 2016. These findings should not be read as apportioning blame or liability to any particular organisation or individual.

Contributing factors

• During the night visual flight rules take-off from Crookwell, the helicopter entered unexpected low cloud.
• The helicopter’s anti-collision light reflecting against nearby cloud distracted the pilot while she was using night vision goggles.
• When responding to the inadvertent entry into instrument meteorological conditions, the pilot over-controlled the helicopter nose down. The resulting unusual attitude at low altitude triggered the enhanced ground proximity warning system to provide a caution alert.

Other factors that increased risk

• Although the pilot had checked weather information for updates prior to departing Canberra, she did not identify significant changes in updated area forecasts pertaining to the planned route from Canberra to Orange. The reason the changes were not identified could not be determined.
• The flight from Canberra to Orange was planned to be conducted under night visual flight rules with the use of night vision goggles, when the use of instrument flight rules was practical and involved less risk, given the forecast and actual weather conditions.
• Although CHC Helicopter Australia’s operations manual stated that emergency medical service flights should be conducted under instrument flight rules (IFR) ‘where practical’, its procedures for night visual flight rules (NVFR) operations using night vision goggles did not clearly state when IFR rather than NVFR should be used. [Safety issue]
• The aircrew member and paramedic had significant concerns regarding the expected weather conditions for the flight from Canberra to Orange, but these concerns were not effectively communicated to, and/or resolved with, the pilot prior to them agreeing to depart Canberra.
• Due to the distraction of the red anti-collision light reflecting off cloud soon after take-off from Crookwell, and high workload, the pilot requested the aircrew member (ACM) to turn off the light. The operation of the light switch was not part of the ACM’s defined and trained duties, therefore creating potential for the task to not be conducted accurately or promptly.
• The aircrew member inadvertently positioned the anti-collision / strobe light switch to the anti-collision and strobe lights position, which resulted in the pilot being exposed to bright white light (reflecting off cloud) while using night vision goggles.

Safety issues and actions

The safety issues identified during this investigation are listed in the Findings and Safety issues and actions sections of this report. The Australian Transport Safety Bureau (ATSB) expects that all safety issues identified by the investigation should be addressed by the relevant organisation(s). In addressing those issues, the ATSB prefers to encourage relevant organisation(s) to proactively initiate safety action, rather than to issue formal safety recommendations or safety advisory notices.

All of the directly involved parties were provided with a draft report and invited to provide submissions. As part of that process, each organisation was asked to communicate what safety actions, if any, they had carried out or were planning to carry out in relation to each safety issue relevant to their organisation.

The initial public version of these safety issues and actions are repeated separately on the ATSB website to facilitate monitoring by interested parties. Where relevant the safety issues and actions will be updated on the ATSB website as information comes to hand.

Policies and procedures for planning IFR and NVIS flights

Safety issue: AO-2016-160-SI-01

Safety issue description: Although CHC Helicopter Australia’s operations manual stated that emergency medical service flights should be conducted under instrument flight rules (IFR) ‘where practical’, its procedures for night visual flight rules (NVFR) operations using night vision goggles did not clearly state when IFR rather than NVFR should be used.

Additional safety action

Whether or not the ATSB identifies safety issues in the course of an investigation, relevant organisations may proactively initiate safety action in order to reduce their safety risk. The ATSB has been advised of the following safety action in response to this occurrence.

CHC Helicopter Australia

As a result of this occurrence, the helicopter operator advised the ATSB it had undertaken the following safety actions:

• integrated its incident investigation and associated learnings into its crew resource management training package
• conducted a review of its simulator training program to ensure inadvertent instrument meteorological conditions (IMC) and unusual attitude procedures were adequately covered
• developed an accident prevention publication regarding inadvertent IMC scenarios
• conducted a review to assess the safety benefit associated with fitment of flight data monitoring equipment for those aircraft within the fleet that did not have it.

Sources and submissions

Sources of information

The sources of information during the investigation included:

• the pilot and crew
• CHC Helicopter Australia
• New South Wales Health Emergency and Aeromedical Services
• OzRunways Pty Ltd
• Bureau of Meteorology
• Airservices Australia.

References

Casner, SM Geven, RW & Williams, RT 2013, ‘The effectiveness of airline pilot training for abnormal events’, Human Factors: The Journal of the Human Factors and Ergonomics Society, vol. 55, pp.477-485.

Crognale, MA & Krebs, WK 2011, ‘Performance of helicopter pilots during inadvertent flight into instrument meteorological conditions’, The International Journal of Aviation Psychology, vol. 21, pp. 235-253.

Davis, D 2001, ‘Foibles of witness memory for traumatic / high profile events’, Journal of Air Law and Commerce, Vol. 66, pp. 1421-1549.

Landman, A Groen, EL van Passen, MM Bronkhorts, AW & Mulder, M 2017, ‘The influence of surprise on upset recovery performance in airline pilots’, The International Journal of Aerospace Psychology, vol. 27, pp. 2-14.

Wuerz, RC & O’Neal, R 1997, ‘Role of pilot instrument proficiency in the safety of helicopter emergency medical services’, Academic Emergency Medicine, vol. 4, pp. 972-975.

Submissions

Under Part 4, Division 2 (Investigation Reports), Section 26 of the Transport Safety Investigation Act 2003 (the Act), the Australian Transport Safety Bureau (ATSB) may provide a draft report, on a confidential basis, to any person whom the ATSB considers appropriate. Section 26 (1) (a) of the Act allows a person receiving a draft report to make submissions to the ATSB about the draft report.

A draft of this report was provided to the operator, pilot, aircrew member (ACM) and medical team on VH-SYB, NSW Health Emergency and Aeromedical Services, OzRunways and the Civil Aviation Safety Authority (CASA).

Submissions were received from CASA, the ACM, the paramedic and OzRunways. The submissions were reviewed and, where considered appropriate, the text of the report was amended accordingly.

An amended draft report was sent to the operator, pilot, ACM and medical team on VH-SYB, NSW Health Emergency and Aeromedical Services, OzRunways and CASA.

Submissions were received from the operator and the ACM. The submissions were reviewed and, where considered appropriate, the text of the report was amended accordingly.

Purpose of safety investigations The objective of a safety investigation is to enhance transport safety. This is done through:  identifying safety issues and facilitating safety action to address those issues providing information about occurrences and their associated safety factors to facilitate learning within the transport industry. It is not a function of the ATSB to apportion blame or provide a means for determining liability. At the same time, an investigation report must include factual material of sufficient weight to support the analysis and findings. At all times the ATSB endeavours to balance the use of material that could imply adverse comment with the need to properly explain what happened, and why, in a fair and unbiased manner. The ATSB does not investigate for the purpose of taking administrative, regulatory or criminal action. Terminology An explanation of terminology used in ATSB investigation reports is available here. This includes terms such as occurrence, contributing factor, other factor that increased risk, and safety issue. Publishing information Released in accordance with section 25 of the Transport Safety Investigation Act 2003 Published by: Australian Transport Safety Bureau © Commonwealth of Australia 2018 Ownership of intellectual property rights in this publication Unless otherwise noted, copyright (and any other intellectual property rights, if any) in this report publication is owned by the Commonwealth of Australia. Creative Commons licence With the exception of the Coat of Arms, ATSB logo, and photos and graphics in which a third party holds copyright, this publication is licensed under a Creative Commons Attribution 3.0 Australia licence. Creative Commons Attribution 3.0 Australia Licence is a standard form licence agreement that allows you to copy, distribute, transmit and adapt this publication provided that you attribute the work. The ATSB’s preference is that you attribute this publication (and any material sourced from it) using the following wording: Source: Australian Transport Safety Bureau Copyright in material obtained from other agencies, private individuals or organisations, belongs to those agencies, individuals or organisations. Where you wish to use their material, you will need to contact them directly.

What happened

On 10 November 2016, a Sikorsky S-92A helicopter, registered VH-ZUQ (ZUQ), was scheduled to fly from Broome Airport, WA, to an offshore facility. The flight crew consisted of a pilot flying (PF) in the right crew seat, and pilot monitoring (PM) in the left crew seat^[1]. The PM was undergoing conversion training on the S-92A.

During the start procedure, the flight crew reported feeling abnormal vibrations in the airframe. The crew believed this might have been the result of a crosswind or recent maintenance work performed on the aircraft, and the PF attempted to minimise it by adjusting the cyclic control.

At approximately 1215 Western Standard Time (WST), a member of the ground crew showed the pilots an anti-flap stop that had broken off the main rotor head. Figure 1 shows the recovered anti-flap stop and securing hardware. Figure 2 shows the anti-flap stop attached to the main rotor hub. The helicopter was subsequently shut down. Further inspection revealed that two anti-flap stops had been sheared off the main rotor head during start-up. One of the stops landed next to the aircraft, while the other struck one of the main rotor blades and narrowly missed a member of the ground crew. It was recovered 45 metres from the aircraft.

Figure 1: Anti-flap stop and securing hardware

Source: Operator

Figure 2: Anti-flap stop secured to the main rotor hub

Source: Operator

The damage to the anti-flap stops was found to have been the result of the PM not lowering the collective control lever at the appropriate time during the engine start. The raised collective resulted in the blades lifting upward, placing abnormal stresses on the anti-flap stops (which are designed to limit upward movement of the blades). This behaviour resulted in the vibrations experienced by the crew.

Events leading up to the broken anti-flap stops

The flight crew had flown together on each of the three days leading up to the serious incident. In all three flights, the pilots were acting in the opposite roles compared with the incident flight. The PM on the day of the incident had never flown an S-92 in that role before.

On the day of the serious incident, the crew arrived approximately two hours before the scheduled take-off. This was in accordance with base instructions that required at least 90 minutes for flight planning and pre-flight inspections. However, the PM was required to make a phone call to the company’s flight operations department and the crew were then involved in a discussion regarding the PM’s further training requirements.

The flight crew were also required to ‘shadow plan’^[2] and perform the pre-flight inspection on a second aircraft, to ensure the flight could proceed in the event that ZUQ was unable to fly. According to the operator’s daily flying roster, the base pilot-in-command (PIC) was originally rostered on to be part of the backup crew. However, client obligations on the base prevented the PIC from being available in the event of the backup crew being required. The PF and PM were not aware of this prior to arriving that day, so they had not given themselves additional time to prepare a second aircraft before departure.

Prior to departure, the PF was tasked with pre-flight inspections of ZUQ as well as the backup aircraft, but they were delayed as maintenance activities were still being conducted on ZUQ. In addition, flight data from ZUQ’s previous flight had not yet been downloaded and analysed, further delaying the PF.

Flight crew comments

Both the PF and PM felt time pressure compounding from the morning’s events. Given that, in accordance with the base rules, 90 minutes is required to prepare for a single flight and considering the number of distractions and delays encountered that morning, the crew felt that more time was required to adequately plan and prepare for both flights. As a result of feeling rushed and to ensure that they were fully prepared for the flight, they completed a second threat and error review, rechecked the flight plan together, rechecked the helicopter’s technical log and walked to the helicopter together.

The flight crew perceived a significant amount of pressure from the operator’s client for this flight. They believed that this client, more than any other they were aware of, required flights to adhere to strict schedules.

Engine start procedures

For engine starts with the rotor brake off, as were normally performed by this operator, the collective must be lowered as part of the start sequence. According to the operator’s standard operating procedures, the PF starts the number 1 engine first, and brings the throttle to idle. When the rotational speed of the main rotor is over 20%, and the hydraulic pressure reaches an appropriate level, the procedures state that the collective is to be moved to the full down position. The PF then starts the number 2 engine.

The S-92A was the only helicopter in the operator’s fleet that required the collective to be lowered as part of the start sequence. It was also the only helicopter where responsibility for controls was split between the two pilots. The operator’s procedures did not designate the task of lowering the collective to either the PM or PF. However, the operator’s parent company uses procedures that specifically assign the role to the PM. Despite not being contained in the operator’s procedures, both pilots understood that it was the role of the PM to lower the collective during engine start and this had been briefed prior to the engine start.

Operator’s investigation

Immediately after the serious incident, the operator commenced its own safety investigation. It identified that the PM had not lowered the collective at the appropriate time during the start procedure. A number of factors were identified. Some of these are listed below:

• Within the operator’s fleet, the requirement that the PM, rather than the PF, lower the collective is unique to the S-92A.
• During the S-92A start procedure, there is no documented requirement for a call out and challenged response between the flight crew to ensure the collective is lowered.
• The PM had been serving as PF for the last three flights the crew carried out together.

Safety analysis

Almost immediately after they arrived at work in the morning, the flight crew began experiencing steadily increasing time pressure. The crew had more work than expected which resulted in less time in which to plan their flight. This pressure was great enough that the flight crew discussed it prior to the flight.

On the incident flight, the PM was flying for their first time in that particular role on a S-92A. This meant that the PM had never been tasked with lowering the collective on a S-92A during start-up. This was the first helicopter that the PM had flown where the responsibility for controls (collective, cyclic and anti-torque pedals) was split between the two pilots. In addition, there were no other helicopters in the operator’s fleet in which the collective must be lowered during the start-up procedure. The PM’s lack of experience in this particular role may have contributed to the collective not being lowered at the appropriate time.

The operator’s standard operating procedures required the collective to be lowered during the start-up procedure. However, there was no requirement on either pilot to “call out” in order to check/verify the collective position. If a call out had been required by the procedures, it is possible that the PF would have noticed it had not been lowered and the incident might have been avoided.

There was no division of labour specified between the PF and PM in the operator’s documentation. In this particular instance, both the PF and PM were aware of their responsibilities prior to start-up, so this likely did not contribute to the incident. However, this ambiguity in the operator’s procedure has the potential to cause similar problems in future.

Findings

These findings should not be read as apportioning blame or liability on any organisation or individual.

• A series of events resulted in the flight crew having more tasks to complete than originally planned, and less time in which to complete them. This resulted in time pressure on the crew.
• The PM's unfamiliarity with the unique starting procedures in the S-92A compared with other helicopters contributed to the collective not being lowered in time.
• The operator's standard operating procedures did not require any call outs for the lowering of the collective, which probably contributed to the collective not being lowered in time.
• With regard to start-up, the operator's standard operating procedures do not provide an explicit division of tasks between the PF and PM.

Safety action

Whether or not the ATSB identifies safety issues in the course of an investigation, relevant organisations may proactively initiate safety action in order to reduce their safety risk. The ATSB has been advised of the following proactive safety action in response to this occurrence.

The Operator

The operator’s investigation into this incident produced 17 recommendations, some of which include:

• a review of S-92A start procedures, addressing as a minimum:
  • flight crew division of duties with specific regard for the collective control lever
  • standardised calls relating to the position of the collective control lever
  • the use of the rotor brake for the start procedure^[3]
• changes to the engine start procedures, the section on lowering the collective is currently written in a ’Note’, this will be changed to a ‘Caution’ to highlight its importance to flight crew
• a review of the policy and procedures used for shadow planning be undertaken, addressing as a minimum all hazards and risks associated with shadow planning inclusive of whether flight crew undergoing training should be exposed to the requirement.

Safety message

This incident is an example of what can occur when pressure associated with on-time departures is compounded with the absence of clear operating procedures. It is important that pilots have detailed and specific instructions on which to fall back when they feel pressure beginning to build.

Part of Aviation Short Investigations Bulletin - Issue 60

About this report Decisions regarding whether to conduct an investigation, and the scope of an investigation, are based on many factors, including the level of safety benefit likely to be obtained from an investigation. For this occurrence, a limited-scope, fact-gathering investigation was conducted in order to produce a short summary report, and allow for greater industry awareness of potential safety issues and possible safety actions. Publishing information  Released in accordance with section 25 of the Transport Safety Investigation Act 2003 Published by: Australian Transport Safety Bureau © Commonwealth of Australia 2017 Ownership of intellectual property rights in this publication Unless otherwise noted, copyright (and any other intellectual property rights, if any) in this report publication is owned by the Commonwealth of Australia. Creative Commons licence With the exception of the Coat of Arms, ATSB logo, and photos and graphics in which a third party holds copyright, this publication is licensed under a Creative Commons Attribution 3.0 Australia licence. Creative Commons Attribution 3.0 Australia Licence is a standard form licence agreement that allows you to copy, distribute, transmit and adapt this publication provided that you attribute the work. The ATSB’s preference is that you attribute this publication (and any material sourced from it) using the following wording: Source: Australian Transport Safety Bureau Copyright in material obtained from other agencies, private individuals or organisations, belongs to those agencies, individuals or organisations. Where you wish to use their material, you will need to contact them directly.

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1. While roles may vary between operations, the pilot flying is primarily responsible for handling the aircraft. The pilot monitoring monitors the pilot flying and is usually responsible for other tasks such as radio calls.
2. Shadow planning refers to creating a completely separate flight plan for another aircraft.
3. Rotor brake starts make it easier for a single pilot to operate all controls, as the engines can be started before control input is required.

What happened

On 21 November 2016, at about 0730 Eastern Daylight-saving Time (EDT), a Bell 206B helicopter, registered VH-CHO, took off from a property about 30 km south of Bathurst, New South Wales. The pilot was conducting an aerial inspection of the property, with the farm manager on board as a passenger.

After overflying a flat area on high ground, the pilot raised the collective^[1] and turned the helicopter to follow down-sloping terrain. At about 75 ft above ground level and an airspeed of about 40 kt, the pilot and passenger heard and felt a bang. The pilot looked outside to see if there was any obvious damage to the spray booms on the helicopter and in the back seat to see whether anything had fallen onto the floor, and assessed that the helicopter was too high to have collided with anything outside.

The pilot initially decided to land as soon as possible in order to check the helicopter to determine the cause of the bang, and started to slow it down. As the airspeed decreased, the helicopter started to yaw rapidly to the right and the pilot, unable to arrest the rotation with left anti-torque pedal, realised they had lost tail rotor authority. The pilot immediately rolled the throttle to the ground idle detent and as the helicopter stopped yawing, lowered the collective. The pilot saw that the rotor rpm had dropped to about 80 per cent and prepared for a hard landing. The pilot cushioned the landing by pulling back on the cyclic,^[2] but the helicopter landed heavily. The belly tank (for spraying) absorbed some of the impact of the landing. The spray booms (fitted to the aircraft for spraying operations) probably helped prevent the helicopter rolling over.

The electronic locator beacon activated on impact. The pilot and passenger sustained minor injuries and the helicopter was substantially damaged (Figure 1).

Figure 1: Accident site showing damage to VH-CHO

Source: Aircraft operator

Post-accident inspection

The tail rotor driveshaft had fractured at the No. 2 bearing (Figure 2).

The tail rotor blades were undamaged in the impact, which indicated that they were probably not rotating when the helicopter collided with the ground. There were scrapes inside the tail rotor driveshaft cowling, which indicates that the shaft was probably rotating when it fractured.

There was no evidence of oil leakage, overheating, or vibration of the tail rotor driveshaft system.

Figure 2: Fractured tail rotor driveshaft

Source: Aircraft engineer

Maintenance history

The helicopter had serial number 714, and was fitted with a long tail rotor driveshaft. The manufacturer required the single long driveshaft to be replaced with a segmented shaft in serial numbers 1252 and above.

The helicopter was fitted with a data augmentation monitoring system, which did not show any abnormalities.

The tail rotor bearings were ‘on condition’ items, that is, were required to be replaced if worn on inspection. The driveshaft was inspected every 1,200 hours and was in good condition apart from the fracture after the accident.

On 18 August 2016, the No. 1 and No. 3 tail rotor driveshaft bearings and tail rotor gearbox were replaced. On 29 September 2016, the No. 1 tail rotor bearing and bearing hanger were replaced. Post-accident inspection did not reveal any abnormalities.

Manufacturer comments

The helicopter manufacturer (Bell Helicopter) reported that the single long tail rotor driveshaft (P/N 206-040-330-001) was replaced with the segmented shafts in Bell 206B helicopters about 45 years ago. At serial number 1252, all Bell 206B helicopters and follow-on Bell 206B3 helicopters were equipped with segmented shafts. The change was made for ease of maintenance in replacing hanger bearings. In addition, if a segmented shaft was damaged, only it would have to be replaced and not the entire long shaft. The long shaft is still procurable through Bell Helicopter. The manufacturer reported that the long tail rotor driveshaft has not been a safety concern with no recent failures recalled.

Manufacturer investigation

The helicopter manufacturer inspected the tail rotor driveshaft and found that it fractured as a result of fatigue cracking. The origin of the fatigue was associated with an area of corrosion pitting, intergranular corrosion and associated cracking on the outside surface of the driveshaft tube, around both sides of the second hanger bearing inner ring location. The driveshaft was otherwise found to comply with the manufacturer’s requirements.

Pilot comments

The pilot provided the following comments:

• When the pilot heard the bang, the helicopter was at a very high-power setting with the blades highly pitched. Although the pilot rolled the throttle off, it took a couple of seconds before the yawing stopped, then they put the collective down and by that time the rotor rpm had dropped.
• There were about 30 gallons (114 litres) of fuel on board and the helicopter was loaded well within weight and balance limitations.
• The temperature was 18.5 °C, the wind light and variable, and the elevation about 3,200 ft above mean sea level. The conditions were well within the performance limitations of the helicopter.
• If the pilot had been able to get into a position where they could do a run-on landing, there may not have been any damage.
• The pilot did not know the tail rotor authority was lost until it was too late to do anything other than land immediately.

Safety analysis

The pilot and passenger heard a bang, which was probably the tail rotor driveshaft failing in flight, as indicated by the scraping inside the tail rotor driveshaft cowling. This resulted in a loss of authority of the tail rotor, and the lack of tail rotor blade damage suggests the blades were (near) stationary at the time of ground impact. The manufacturer’s examination found that the driveshaft tube fractured as a result of fatigue cracking that was associated with an area of corrosion pitting on the outside surface of tube, adjacent to the location of the second hanger bearing. The pilot slowed the helicopter to make an approach to land and check out the cause of the bang, which caused the helicopter to start yawing to the right. The pilot then realised they had lost tail rotor authority, rolled off the throttle to stop the yaw, and as the rotor rpm had dropped they were then committed to an immediate landing.

Findings

These findings should not be read as apportioning blame or liability to any particular organisation or individual.

• The tail rotor driveshaft tube fractured as a result of fatigue cracking that likely originated from corrosion pitting on its outside surface. This resulted in a loss of tail rotor authority.
• The helicopter was low and slow when tail rotor authority was lost, limiting the time available for the pilot to assess the circumstances and manage the forced landing.

Safety message

This incident highlights the importance of robust and current training in emergency procedures. Being able to identify a problem and react quickly can reduce the severity of damage and injuries.

About this report Decisions regarding whether to conduct an investigation, and the scope of an investigation, are based on many factors, including the level of safety benefit likely to be obtained from an investigation. For this occurrence, a limited-scope, fact-gathering investigation was conducted in order to produce a short summary report, and allow for greater industry awareness of potential safety issues and possible safety actions. Publishing information Released in accordance with section 25 of the Transport Safety Investigation Act 2003 Published by: Australian Transport Safety Bureau © Commonwealth of Australia 2017 Ownership of intellectual property rights in this publication Unless otherwise noted, copyright (and any other intellectual property rights, if any) in this report publication is owned by the Commonwealth of Australia. Creative Commons licence With the exception of the Coat of Arms, ATSB logo, and photos and graphics in which a third party holds copyright, this publication is licensed under a Creative Commons Attribution 3.0 Australia licence. Creative Commons Attribution 3.0 Australia Licence is a standard form licence agreement that allows you to copy, distribute, transmit and adapt this publication provided that you attribute the work. The ATSB’s preference is that you attribute this publication (and any material sourced from it) using the following wording: Source: Australian Transport Safety Bureau Copyright in material obtained from other agencies, private individuals or organisations, belongs to those agencies, individuals or organisations. Where you wish to use their material, you will need to contact them directly.

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1. Collective: a primary helicopter flight control that simultaneously affects the pitch of all blades of a lifting rotor. Collective input is the main control for vertical velocity.
2. Cyclic: a primary helicopter flight control that is similar to an aircraft control column. Cyclic input tilts the main rotor disc, varying the attitude of the helicopter and hence the lateral direction.

What happened

On 18 November 2016, a Robinson R44 helicopter, registered VH-HHZ, was operating to support fire-fighting personnel near Sleisbeck, Northern Territory.

At about 1600 Central Standard Time (CST), the pilot conducted an approach to a landing site they had already landed at twice that day. The landing site was a flat rocky surface, but one side had a slight downwards slope. The pilot confirmed the wind direction from smoke nearby and approached the landing site into wind. As the pilot lowered the collective^[1] and the helicopter’s skids touched down, the helicopter started to slide to the right. The pilot attempted to correct the sideways movement, but the main rotor blade struck a rock, and the helicopter started vibrating. The pilot rolled off the throttle and applied right pedal, but the helicopter rotated to the left and the horizontal stabiliser struck a rock. The helicopter sustained substantial damage (Figure 1). The pilot and two passengers were not injured.

Figure 1: Accident site showing damage to VH-HHZ

Source: Helicopter operator

Findings

These findings should not be read as apportioning blame or liability to any particular organisation or individual.

• The helicopter landed on a portion of rock, which had a slight downwards slope. The actions taken by the pilot, when the helicopter started to slide, did not prevent the main rotor blades and the horizontal stabiliser from striking a rock.

Safety action

Whether or not the ATSB identifies safety issues in the course of an investigation, relevant organisations may proactively initiate safety action in order to reduce their safety risk. The ATSB has been advised of the following safety action in response to this occurrence.

Helicopter operator

As a result of this occurrence, the helicopter operator has advised the ATSB that they are taking the following safety actions:

The company issued a notice to flight crew emphasising the importance of conducting a thorough aerial assessment before committing to landing at any remote location including confined areas.

The company also reminded pilots that in hot humid conditions, fatigue can occur a lot sooner that during the cooler months. They should make every effort to remain hydrated and inform the chief pilot immediately if they feel adversely affected.

Safety message

Civil Aviation Advisory Publication 92-2(2), Guidelines for the establishment and operation of onshore Helicopter Landing Sites, sets out factors that may be used to assess the suitability of a site for helicopters to land and take off. The guidelines include the recommendation that helicopter operators conduct thorough risk and hazard assessments for a basic helicopter landing site and implement controls to manage identified hazards.

Aviation Short Investigations Bulletin - Issue 57

About this report Decisions regarding whether to conduct an investigation, and the scope of an investigation, are based on many factors, including the level of safety benefit likely to be obtained from an investigation. For this occurrence, a limited-scope, fact-gathering investigation was conducted in order to produce a short summary report, and allow for greater industry awareness of potential safety issues and possible safety actions. Publishing information  Released in accordance with section 25 of the Transport Safety Investigation Act 2003 Published by: Australian Transport Safety Bureau © Commonwealth of Australia 2017 Ownership of intellectual property rights in this publication Unless otherwise noted, copyright (and any other intellectual property rights, if any) in this report publication is owned by the Commonwealth of Australia. Creative Commons licence With the exception of the Coat of Arms, ATSB logo, and photos and graphics in which a third party holds copyright, this publication is licensed under a Creative Commons Attribution 3.0 Australia licence. Creative Commons Attribution 3.0 Australia Licence is a standard form licence agreement that allows you to copy, distribute, transmit and adapt this publication provided that you attribute the work. The ATSB’s preference is that you attribute this publication (and any material sourced from it) using the following wording: Source: Australian Transport Safety Bureau Copyright in material obtained from other agencies, private individuals or organisations, belongs to those agencies, individuals or organisations. Where you wish to use their material, you will need to contact them directly.

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1. Collective: a primary helicopter flight control that simultaneously affects the pitch of all blades of a lifting rotor. Collective input is the main control for vertical velocity.

Safety summary

What happened

On 21 November 2016 at about 0730 Eastern Daylight‑saving Time, the pilot of an Air Tractor Inc. AT-802A, registered VH-NIA, was conducting aerial spraying activities from Trangie airfield near Narromine, New South Wales. The planned activities included spraying a small area of crop on a property about 30 km west of Narromine. The crop spraying was a continuation of the previous day’s activities that had been discontinued due to the weather becoming unsuitable for spraying conditions.

The property owner reported observing the aircraft arrive and that the pilot appeared to be experienced in the way he was manoeuvring the aircraft while spraying the crop. The property owner assumed that the pilot had completed spraying the crop as, after about 30 minutes, the aircraft departed in the direction of Trangie airfield.

At about 0810, witnesses briefly observed the aircraft to be in a nose-down attitude before impacting the ground, resulting in an intense fuel‑fed fire. The accident site was located about 5 km from the spray area. The pilot was fatally injured, and the aircraft was destroyed.

What the ATSB found

The ATSB found that the aircraft departed controlled flight, from which the pilot was unable to recover, leading to the collision with terrain. Based on the available evidence, it was not possible to determine the reasons for the loss of control.

The ATSB identified a number of observed incidents or potentially unsafe aircraft operations involving the accident pilot that were not reported to the operator’s chief pilot. This decreased the opportunity for the operator to identify and address risks that could affect the safety of operations.

What's been done as a result

The operator advised that meetings with staff have been beneficial in highlighting the importance of reporting incidents and accidents despite any concerns about an employee’s seniority or role in the company. Additionally, induction processes and documented safety reporting procedures have been reinforced.

Safety message

Operators must ensure that all personnel have an understanding of the importance of timely reporting of events that increase safety risk. A good safety reporting culture can assist operators with monitoring trends, identifying operational issues, and providing a timely response to reduce risk within the operating environment.

Air Tractor AT-802A, VH-NIA

Source:  Jayden Laing

The occurrence

On 21 November 2016, the pilot of an Air Tractor Inc. AT-802A, registered VH-NIA (NIA), was preparing to conduct aerial crop spraying activities from Trangie airfield near Narromine, New South Wales. The planned crop spraying was a continuation of the previous day’s spraying activities that were conducted on a property about 30 km west of Narromine (Figure 1). The previous day’s spraying had not been completed because of deteriorating weather conditions.

The operations manager recalled meeting the pilot at Trangie airfield at 0630 Eastern Daylight‑saving Time^[1] where about 1890 L of chemical solution was loaded into the aircraft’s hopper. The operations manager reported that the pilot considered this adequate to complete the job as there was only a small spray area remaining. A loader reported that the aircraft’s fuel tanks were likely to have been full before take-off as they were routinely filled at the end of each day’s flying.

The property owner reported observing the aircraft arrive and commence spraying the remaining paddocks at about 0730. He reported that, judging by the way the aircraft was manoeuvred while spraying the crop, the pilot appeared to be experienced and there did not appear to be any difficulties with the operation of the aircraft. The property owner observed the aircraft leave the spray area after about 30 minutes. They assumed that the pilot had completed the spraying, as the aircraft was flying in the direction of Trangie.

Figure 1: Accident location

Source: Google Earth, modified by the ATSB

At about 0800, a contractor driving away from the property witnessed the pilot spraying alongside Doonside airstrip (Figure 2), which was bordered by a crop thought to have been sprayed on the previous day. The contractor reported observing the aircraft flying beside the airstrip with some up and down movements, before the spray stopped and the aircraft commenced a climb and continued north.

A farmer located between the airstrip and a short distance (2 km) from the accident site reported seeing the aircraft fly overhead at an altitude ‘that was quite high for an agricultural aircraft’. The farmer estimated the aircraft to have been at about 150 ft (45 m). Although the farmer could not clearly hear the aircraft because of machinery noise, the aircraft flew overhead in a level attitude and appeared to be in normal, controlled flight.

Figure 2: Location of witnesses and approximate flight path

Source: Google Earth, modified by the ATSB

At about 0810, witnesses driving along a nearby road observed the aircraft above a tree line in a steep, almost vertical, nose-down attitude. Some witnesses reported the aircraft also having a slow, right to left roll, as it quickly descended. The aircraft was destroyed by the impact and a subsequent post-impact fire. The pilot was fatally injured.

Personnel

Pilot

The pilot was appropriately qualified for the flight, holding a Commercial Pilot Licence with Aeroplane and Helicopter category ratings, an Aerial Application Rating (Aeroplane and Helicopter) and a Low-Level Rating. The pilot also held an Aerobatic endorsement. The pilot’s most recent logbook recorded a total aeronautical experience in excess of 23,967 hours, with the majority of these hours accumulated conducting agricultural flying activities.

The pilot was endorsed on the Air Tractor Inc. AT-802 (AT-802). The most recent logbook, which commenced in 2011, did not record any hours in this aircraft type. The operator records show that the pilot advised of having accrued about 1,500 hours of flying experience on the AT-802 and 700 hours on the Air Tractor Inc. AT-502/504 (AT-502). Further, the operator had employed the pilot to fly the AT-502/AT-802 on a casual basis over many years and considered him to be a highly experienced and proficient agricultural pilot.

The operator reported that proficiency checks were normally conducted yearly with the company pilots. An additional proficiency check was not required as the pilot’s logbook recorded that a proficiency check had been conducted by another operator on the 30 August 2016 flying an AT-502. The operator that conducted the pilot’s last proficiency check recorded that the pilot had no difficulty operating the AT-502.

The pilot held a valid class 1 medical certificate. Highlighted in the medical records was a pre‑existing medical condition that was medically managed through a CASA-approved medical officer. There was good evidence that the pilot was taking the prescribed medication used to manage the condition. Reports by work colleagues, friends, and relatives indicated that the pilot appeared well on the morning and in the days leading up to the accident.

Operations

The operator was approved to conduct aerial spraying activities in accordance with an air operators certificate issued by CASA.The air operators certificate authorised the operator to engage in aerial work operations, and the requirements of Civil Aviation Order 82.0 and Civil Aviation Safety Regulations Part 137 were applicable. The operator had not implemented, nor was it required to implement, a safety management system.

In accordance with the requirements of their air operators certificate, the operator was required to provide an operations manual for the use and guidance of their personnel. In the performance of their duties, personnel were required to comply with all instructions in the operations manual.The operations manual included a requirement for company personnel to report accidents and incidents. This included the reporting of any occurrence associated with the operation of the aircraft that affects, or had the potential to affect the safety of the operation. To assist with the management of accident and incidents reported to the supervisor, the operator used a software database. Completed reports were to be forwarded to the chief pilot for investigation and follow up corrective action as considered necessary.

Incidents and accidents

During the investigation, it was identified that the pilot in command of the accident flight had been involved in a number of operational events in the six days leading up to the accident flight. These events all occurred while the pilot in command was performing his flying duties in the AT‑802.

Most of the identified operational events had the potential to compromise safety, and met the operator’s definition of an incident included in the company operations manual. The events observed by company personnel included:

• A runway excursion during landing that required the aircraft to be towed from the runway edge drain
• A subsequent landing incident that resulted in the aircraft deviating from the centreline with the pilot regaining control prior to the aircraft again departing the runway
• An airborne event that involved a tight/aggressive 180-degree turn that, according to a witness who was also a pilot, resulted in a possible stall with the pilot recovering control at a very low altitude
• A downwind take-off approaching maximum take-off weight, which resulted in the aircraft narrowly missing trees at the departure end of the runway
• During a spray run, the aircraft’s right wing spray equipment was damaged due to contacting the crop being sprayed. The pilot was reported to have removed the vegetation from the spray equipment, conducted a repair, and recommenced spraying operations.

The ATSB requested all of the operator’s accident and incident reports recorded in the company’s safety database in the 12 months prior to the VH-NIA accident. There was one unrelated incident report recorded in the database.

Despite being a company requirement, none of the incidents detailed above were reported to the chief pilot or management. Witnesses to the incidents had either discussed the events with other pilots, or were intending to discuss these matters with management at the first available opportunity. In the absence of incident reports associated with these events, there was no ability for the operator to undertake an investigation and subsequently determine if corrective action was required.

Spraying activities

The operator’s fleet primarily consisted of AT-502 and AT-802 aircraft. It was reported by the operator that each aircraft would have a specific pilot allocated to each aircraft.The recent spraying season, however, had been busy and the operator employed additional experienced pilots on a casual basis. The use of casual pilots allowed the permanent pilots the opportunity to rest and have a break from flying activities, in preparation for the next aerial spraying and fire season.

Loaders supported pilots conducting aerial spraying and were generally located at an airfield central to spraying operations. Primarily, the loaders were responsible for refilling the aircraft’s spray tank (hopper) and refuelling the aircraft. A loading truck was used to refuel the aircraft in addition to mixing and transferring chemicals into the aircraft’s hopper.

The spraying operation on the day before the accident required a loader to be at the Doonside airstrip (Figure 2), and the accident pilot to reposition NIA from Trangie airfield, to load the first quantity of chemicals. Spraying commenced at about 0730, and a number of loads were applied during the morning’s operations. The loader reported that the pilot uplifted a full hopper of about 3,000 L of chemical mix, and that each load took about 30 minutes to spray.

The loader reported that during the morning, the pilot advised the spraying conditions had deteriorated and a remaining load would need to be applied the following day. The pilot subsequently completed the spray activities and returned to Trangie airfield at about 1030. In accordance with the reported standard company practice, the loader completely filled the aircraft’s fuel tanks after returning with the loading truck to Trangie airfield.

Meteorological information

On-site evidence, and data recorded at a nearby agricultural recording weather station indicated a 9 kt (16 km/h) wind from the north-east, and a temperature of 26 °C at about the time of the accident.

Aircraft information

The Air Tractor Inc. AT-802A aircraft was of tail wheel, fixed landing gear design. The aircraft was powered by a Pratt & Whitney Canada PT6A-67A turboprop engine and was purpose-built for use in aerial agricultural applications and fire control operations.

VH-NIA was manufactured in 2003, and had a current special certificate of airworthiness, certificate of registration, and maintenance release. The last maintenance inspection was conducted about two weeks and 50 flight hours prior to the accident. The maintenance release was identified at the accident site, and indicated that the aircraft had a total time in service of 2812.9 flight hours before the accident flight. There were no outstanding maintenance requirements or defect endorsements entered on the maintenance release. Examination of the aircraft’s maintenance documentation did not identify any issues that would have been detrimental to the operation of the aircraft.

Site and wreckage

Site examination

The accident site was located adjacent to a road on cleared flat farmland, about 33 km west of Narromine, NSW. The wreckage trail was about 70 m long, towards the south-west. The initial ground impact marks were from the upper part of the vertical stabiliser, left and right wings, upper cockpit area, engine and propeller (Figure 3). Those marks and items such as navigation lights, tail and cockpit components indicated that the aircraft impacted with terrain inverted, with the left wing striking the ground first, oriented on the right side of the wreckage trail.

Figure 3: Accident site, showing ground impact marks and the main wreckage in the background

Source: ATSB

Wreckage examination

A post-impact fuel-fed fire began during the accident sequence and consumed the majority of the aircraft wreckage. Figure 4 shows the remaining sections of the main wreckage, which were inverted. The level of disruption and fire damage significantly reduced the amount of evidence available to be examined.

Figure 4: Main wreckage inverted with the nose (engine and propeller) in the foreground

Source: ATSB

Inspection of the remaining wreckage indicated that:

• there was no evidence of impact with trees, powerlines, or birds
• all of the aircraft’s main structural components were in the immediate area of the accident site
• the main support structure had no identified pre-impact defects
• the left wing had bending damage to the main and rear spars that was greater than that observed on the right wing structure, indicating that the left wing most likely struck the ground first
• the vertical stabiliser had separated from the empennage due to downward and back bending forces, which was a further indication of an inverted impact with terrain
• sections of the upper cockpit area, vertical stabiliser, and wing secondary structure had separated from the fuselage. 

The aircraft’s approximate angle of entry was calculated using the position of aircraft components and angled crush damage to the tail section of the aircraft (Figure 5). The aircraft most likely impacted the ground inverted, at an angle of about 30-40° nose down.

Figure 5: Side view of an AT-802 aircraft with superimposed tail and rudder showing angled crush damage

Source: Air Tractor Inc., modified by the ATSB

Flight controls

Examination of the flight control surfaces, control cables and push rods did not identify any pre-impact defects. The fire damaged flap actuator was located within the wreckage, and its attachment points were destroyed by fire. A measurement was taken on the threaded portion of the actuator to ascertain flap position. That measurement indicated that the flaps were in the fully extended position at the time of impact with terrain. Based on the trim actuator and trim control surface positions, the trim position was calculated as being almost fully nose up. Given the wreckage disruption and trim cable disconnection from the actuator, the trim position prior to impact could not be confirmed.

Engine

An external examination of the engine did not identify any pre-impact defects (Figure 6). The engine first stage compressor was inspected through the inlet and the second stage power turbine through the exhaust outlet. No pre-impact defects that would indicate an internal failure were identified.

Figure 6: Engine and propeller assembly

Source: ATSB

The engine reduction gearbox was removed from the accident site for further examination of the engine to propeller drive components, in an area of what appeared to be overload failure (Figure 7).

Figure 7: Reduction gearbox drive section viewed from the rear with three of the ten fracture points arrowed

Source: ATSB

Detailed examination of the drive components showed that:

• the fractures observed were overstress, due to high torsional loads
• the direction of overstress failure was consistent with a sudden propeller stoppage while the engine was driving the propeller (Figure 8).

The engine manufacturer confirmed that the damage observed in the drive components was indicative of the engine producing power at the time of impact.

Figure 8: Reduction gearbox drive section showing fracture points, damage to bearing sleeves and an illustration showing direction of failure.

Source: ATSB and P&WC, modified by the ATSB

Propeller

The propeller was inspected on-site. Two of the five blades had been liberated from the hub and were located in the vicinity of the first impact point. The propeller was removed from the accident site, disassembled and examined at a propeller overhaul facility, under the supervision of the ATSB. The examination revealed that:

• the internal components of the propeller hub did not have any pre-impact defects
• only one blade did not have its pitch link broken at impact. Its pitch angle was calculated to be 43 degrees, which was reported by the propeller overhaul facility staff as being in the cruise power pitch range
• all other blades either had broken pitch linkages or were liberated from the hub, which precluded an accurate measurement of their positions
• two of the propeller blades showed signs of either double bending (forward at the tip and rearward through the mid-section) or bending in the opposite direction to rotation. That indicated that the engine was driving the propeller on impact with terrain.

Recorded data

The aircraft was not fitted with a flight data or cockpit voice recorder, nor was it required to be according to Australian regulations.

The aircraft was equipped with a satellite-based Global Positioning System (GPS) SATLOC AirStar system (SATLOC) to provide guidance for aerial spraying operations. The system also recorded position and spray information. Unfortunately, the data logging card that stores the recorded information was destroyed by the post-impact fire.

Wreckage inspection summary

Although there were no pre-impact defects identified during the wreckage examination, the possibility of an in-flight mechanical failure could not be discounted due to the level of disruption and fire damage.

__________

1. Eastern Daylight-saving Time (EDT): Coordinated Universal Time (UTC) + 11 hours.

Safety analysis

The occurrence

Examination of the aircraft and accident site concluded that the aircraft impacted terrain in an inverted and uncontrolled state. The ATSB considered several scenarios that might explain why the aircraft departed from controlled flight after the pilot had appeared to have completed all low-level spraying activities.

There was no evidence of any mechanical defect or failure within the aircraft or engine that may have contributed to the accident.The level of impact and fire damage, however precluded a detailed examination of all of the aircraft systems. Therefore, a mechanical failure could not be discounted.

A review of the probable environmental conditions indicated that it was unlikely that the weather had an adverse effect on the operation of the aircraft. In addition, there was no evidence of a birdstrike or wirestrike.

A review of the pilot’s recent AT-802 flying experience and incidents prior to the accident indicated that the pilot was not as proficient flying the AT-802 as the AT-502. Although both aircraft are similar, the flying characteristics of the AT-802 are different. The manual tail wheel lock mechanism, increased weight, increased inertia, and reported slower manoeuvring characteristics would all require some degree of pilot adaptation. These differences may not have been fully appreciated by the pilot, and were likely manifested in the way the aircraft was flown on previous flights. Despite this, the pilot conducted the morning’s planned spraying activities successfully, with nearby witnesses reporting that this was done with no apparent difficulty.

After completing the planned spraying and a spray run along a paddock near the Doonside airstrip, the pilot climbed the aircraft to a higher altitude and flew away from the spray area. A farmer located close to the accident site reported the pilot appeared to be in control of the aircraft, and was maintaining level flight moments before the accident.

No conclusive evidence was available to determine how the aircraft went from what appeared to be controlled, level flight, at a reasonable altitude above terrain, to an apparent loss of control and a steep nose down attitude prior to impact with terrain.

Incident reporting

Despite the operator implementing a safety reporting system, the ATSB became aware that a number of incidents and concerns by pilots and loaders about the accident pilot were not reported to the operator using the prescribed procedure.

The inclusion of reporting requirements in the operations manual was intended to assist the operator and management personnel to manage safety outcomes. Company personnel were made aware of these requirements through an induction program, however the investigation noted that in relation to events concerning the accident pilot, these requirements were not followed.

While it could not be established if that affected the outcome in this accident, the accurate and timely reporting of incidents and accidents is essential for organisations to be able to manage safety outcomes. Such reports from operational personnel directly involved in operational activities enable management to take action as necessary to manage risk.

The establishment of a reporting system is only one aspect of effective safety reporting within an organisation. It establishes the platform which enables reporting to occur, but does not provide assurance that personnel will comply with the requirements. Effective reporting systems require integration into the broader management systems of an organisation.

As such, the reporting system of the operator was not effective on its own in ensuring that hazards or perceived risks that existed in the operational environment were reported to management. It is probable that the effectiveness of the established reporting system was decreased by a lack of a systemic approach to its management. As it is possible that concerns held by some operating personnel of the pilot’s flying were perceived by them as being unduly critical, they may not have reported out of a sense of fairness to the pilot. A focus on the need to report, despite common obstacles such as time, distance from the office and ease of reporting, along with stated management support for this process will help to ensure it is followed.

While the lack of reporting of safety events represented a missed opportunity to improve safety outcomes, given the unknown reason for the accident and limited time between the events and the accident, it is not possible to determine if better reporting would have prevented this accident.

Findings

From the evidence available, the following findings are made with respect to the departure from controlled flight and collision with terrain of Air Tractor Inc. AT-802A, registered VH-NIA, that occurred 33 km west of Narromine, New South Wales, on 21 November 2016. These findings should not be read as apportioning blame or liability to any particular organisation or individual.

Safety issues, or system problems, are highlighted in bold to emphasise their importance. A safety issue is an event or condition that increases safety risk and (a) can reasonably be regarded as having the potential to adversely affect the safety of future operations, and (b) is a characteristic of an organisation or a system, rather than a characteristic of a specific individual, or characteristic of an operating environment at a specific point in time.

Contributing factors

• Shortly after departing the spray area, and for reasons that could not be determined, a loss of control occurred from which the pilot was unable to recover before impacting terrain.

Other factors that increased risk

• Several aircraft incidents involving the accident pilot were not reported to the operator’s management as per the documented procedure. This limited the operator’s awareness of potential operational risks.
• The operator's documented procedure for company personnel to report accidents and incidents was in itself not sufficient to ensure that occurrences that had affected, or had the potential to affect safety, were reported to management. This decreased the opportunity for the operator to identify potential operational risks and take appropriate action to minimise them [Safety issue].

Safety issues and actions

The safety issue identified during this investigation is listed in the Findings and Safety issues and actions sections of this report. The Australian Transport Safety Bureau (ATSB) expects that all safety issues identified by the investigation should be addressed by the relevant organisation(s). In addressing those issues, the ATSB prefers to encourage relevant organisation(s) to proactively initiate safety action, rather than to issue formal safety recommendations or safety advisory notices.

All of the directly involved parties were provided with a draft report and invited to provide submissions. As part of that process, each organisation was asked to communicate what safety actions, if any, they had carried out or were planning to carry out in relation to each safety issue relevant to their organisation.

Descriptions of each safety issue, and any associated safety recommendations, are detailed below. Click the link to read the full safety issue description, including the issue status and any safety action/s taken. Safety issues and actions are updated on this website when safety issue owners provide further information concerning the implementation of safety action.

Safety reporting culture and procedures

Safety issue number: AO-2016-157-SI-01

The operator's documented procedure for company personnel to report accidents and incidents was in itself not sufficient to ensure that occurrences that had affected, or had the potential to affect safety, were reported to management. This decreased the opportunity for the operator to identify potential operational risks and take appropriate action to minimise them.

Sources and submissions

Sources of information

The sources of information during the investigation included the:

• operator
• maintenance provider
• witnesses
• engine manufacturer
• aircraft manufacturer
• Civil Aviation Safety Authority (CASA).

Submissions

Under Part 4, Division 2 (Investigation Reports), Section 26 of the Transport Safety Investigation Act 2003 (the Act), the Australian Transport Safety Bureau (ATSB) may provide a draft report, on a confidential basis, to any person whom the ATSB considers appropriate. Section 26 (1) (a) of the Act allows a person receiving a draft report to make submissions to the ATSB about the draft report.

A draft of this report was provided to the aircraft and engine manufacturer, Rebel Ag, the maintenance provider, National Transportation Safety Board, Transportation Safety Board of Canada and the CASA.

A submission was received from CASA. The submission was reviewed and where considered appropriate, the text of the report was amended accordingly.

Purpose of safety investigations The objective of a safety investigation is to enhance transport safety. This is done through:  identifying safety issues and facilitating safety action to address those issues providing information about occurrences and their associated safety factors to facilitate learning within the transport industry. It is not a function of the ATSB to apportion blame or provide a means for determining liability. At the same time, an investigation report must include factual material of sufficient weight to support the analysis and findings. At all times the ATSB endeavours to balance the use of material that could imply adverse comment with the need to properly explain what happened, and why, in a fair and unbiased manner. The ATSB does not investigate for the purpose of taking administrative, regulatory or criminal action. Terminology An explanation of terminology used in ATSB investigation reports is available here. This includes terms such as occurrence, contributing factor, other factor that increased risk, and safety issue. Publishing information Released in accordance with section 25 of the Transport Safety Investigation Act 2003 Published by: Australian Transport Safety Bureau © Commonwealth of Australia 2018 Ownership of intellectual property rights in this publication Unless otherwise noted, copyright (and any other intellectual property rights, if any) in this report publication is owned by the Commonwealth of Australia. Creative Commons licence With the exception of the Coat of Arms, ATSB logo, and photos and graphics in which a third party holds copyright, this publication is licensed under a Creative Commons Attribution 3.0 Australia licence. Creative Commons Attribution 3.0 Australia Licence is a standard form licence agreement that allows you to copy, distribute, transmit and adapt this publication provided that you attribute the work. The ATSB’s preference is that you attribute this publication (and any material sourced from it) using the following wording: Source: Australian Transport Safety Bureau Copyright in material obtained from other agencies, private individuals or organisations, belongs to those agencies, individuals or organisations. Where you wish to use their material, you will need to contact them directly.

Safety summary

What happened

On 18 November 2016, at about 1244 Eastern Standard Time, a Robinson R44 II helicopter, registered VH-ZNZ, broke-up in the Mount Windsor National Park, about 41 km north-west of Mossman, Queensland. The helicopter was on a charter flight with one pilot and one passenger on board. Following impact with the ground, the passenger was fatally injured, and the pilot was seriously injured. The helicopter was destroyed.

What the ATSB found

The ATSB found that one of the main rotor blades struck and separated a section of the tailcone of the helicopter from the airframe, which resulted in a near vertical descent of the helicopter main body through the forest canopy. The wreckage indicated there was low engine power and rotor speed at the time of the strike, which was likely the result of a main rotor blade stall event. The ATSB was unable to determine what precipitated the blade stall event.

The ATSB also found that the helicopter was likely operating in at least moderate turbulent flight conditions, which were not forecast. The turbulence was associated with a wind strength of 15‑25 kt in undulating terrain at a high-density altitude (reduction in ambient air density). The Pilot Operating Handbook did not prohibit flight in either the forecast or actual conditions, but the presence of moderate turbulence potentially placed the helicopter in an environment for which it had not been flight tested.

Safety message

This accident highlighted the importance of impact-activated emergency locator transmitters. The activation of the transmitter on impact was the trigger to start the search and rescue operation, and recover the pilot, who may otherwise have not survived. ATSB research report AR-2012-128 provides guidance to owners and operators on how they can maximise the reliability and effectiveness of emergency locator transmitters.

Although it could not be determined if the wind and associated turbulence contributed to the accident, it is important for pilots to consider the effect of the terrain on the weather forecast, which could result in their helicopter not achieving its predicted performance.

The occurrence

On 18 November 2016, at about 1244 Eastern Standard Time,^[1] a Robinson R44 II helicopter, registered VH-ZNZ, broke-up in the Mount Windsor National Park, about 41 km north-west of Mossman, Queensland. The helicopter was on a charter flight with one pilot and one passenger on board. Following impact with terrain, the passenger was fatally injured and the pilot was seriously injured. The helicopter was destroyed.

The passenger was a contractor providing maintenance services (gas and plumbing inspections) for the Queensland Park and Wildlife Service (QPWS) Northern Estate facilities. He contracted the pilot’s company to provide the transportation services for the site inspections and they were on their fourth day of flying together. The pilot provided the service for the first 2 days with a Bell 206 helicopter. The Bell 206 became unserviceable the day prior to the accident and was replaced with the R44 for 17 and 18 November.

On 18 November, the helicopter departed from Cardwell and stopped at Mareeba Airport, where it was refuelled with 140 L (100 kg) of aviation gasoline at about 1006. It then proceeded to Mossman, where the passenger met QPWS staff. The QPWS staff provided the passenger with the geographic coordinates to locate two facilities for gas appliance inspections in the Mount Windsor National Park. They were the forestry barracks facility and station house facility.

Recorded track data from OzRunways^[2] showed that the helicopter departed from the vicinity of the QPWS Mossman facility at about 1134 and tracked towards the Mount Windsor National Park. The last recorded track was at about 1144, at which time the helicopter was about 18 km to the south-east of the barracks facility and on a direct track towards it.^[3]

At about 1244, the Australian Maritime Safety Authority’s Joint Rescue Coordination Centre (JRCC) detected a signal from the helicopter’s emergency locator transmitter. At about 1432, the helicopter wreckage was located in the dense rainforest of Mount Windsor National Park, about 500–600 m south-east of the barracks facility (refer to Figure 4 in the section titled Wreckage and impact information).

The helicopter was found to have broken into multiple fragments and the main body was subject to a post-impact fire. The pilot required hospitalisation for a number of weeks, which included five days under an induced coma, and subsequently had no recollection of the events on the day of the accident.

Figure 1 depicts the known flight path from recorded track data and the location of the accident site with key event timings. The end of the recorded track is consistent with an expected loss of coverage in that area at low level.^[4]

Figure 1: VH-ZNZ track and accident site

Recorded track data for the flight from Mossman to the Mount Windsor National Park, which is the forested area surrounding the accident site. The start of the track is in the vicinity of the Mossman QPWS office. Source: Google earth and OzRunways, annotated by ATSB.

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1. Eastern Standard Time (EST): Coordinated Universal Time (UTC) + 10 hours.
2. OzRunways is an approved data provider for Australian pilots for flight planning and in-flight navigation.
3. The station house facility was located about 15 km north-west of the forestry barracks facility.
4. Track data is transmitted to OzRunways server using the mobile phone network, which is affected by terrain shielding at lower levels in the area of the accident.

Context

Pilot information

The pilot held a commercial helicopter pilot licence and had accumulated over 5,700 hours of flight time. His licence included the class rating for single-engine helicopters and the type rating for the Robinson R44.^[5] The pilot initially qualified on the Robinson R22, R44 and Bell 206 helicopters in 2007. His last flight review was certified as taking place on 29 May 2016 in an R44. The pilot had flown VH-ZNZ on six occasions since the helicopter’s last periodic maintenance inspection on 7 September 2016, which included the day prior to the accident.

The pilot’s last aviation medical examination was on 19 January 2016. He was issued with a Class 1 Aviation Medical Certificate without restrictions. The pilot reported that prior to the accident flight he had slept about 18 hours in the last 48 hours. From the pilot’s logbook, there was no flying recorded for 12 and 13 November. The accident day, 18 November, was his fifth consecutive day of flying. He recorded 12 hours 20 minutes of flying for the previous four days. The time of the accident was not in the circadian low period and did not include an extended period of duty.

Helicopter information

General

VH-ZNZ was a Robinson Helicopter Company (RHC) R44 II helicopter, serial number 11954, powered by a 6-cylinder horizontally opposed Textron Lycoming IO-540-AE1A5 engine (Figure 2). It was a four-seat helicopter, certified by the United States (US) Federal Aviation Administration (FAA) and accepted by the Civil Aviation Safety Authority (CASA) in the normal category rotorcraft.^[6] RHC manufactured the helicopter in 2007 and it was added to the CASA aircraft register in November of that year.

Figure 2: Robinson R44 II helicopter VH-ZNZ

Source: Ian McDonell

Engine power is transmitted via four V-belts and a clutch to a shaft, which transmits power forward to the main rotor and aft to the tail rotor, via respective gearboxes. An engine governor senses changes in engine speed and applies corrective throttle inputs to maintain engine speed and therefore driveshaft and rotor speed within the normal operating limits. The throttle also opens or closes in response to the pilot raising or lowering the collective lever,^[7] or if the pilot manually adjusts the throttle lever (located on the collective) to override the governor.

The main rotor consists of two all-metal blades mounted to the hub by coning hinges.^[8] The hub is mounted to the main rotor shaft by a teeter hinge and the main rotor head is known as a teetering, or semi-rigid, rotor head. The three-hinged body teetering rotor head is a unique design feature to Robinson helicopters. The tail rotor has two all-metal blades mounted to a teetering hub. Pitch links transmit flight control movements to the rotor systems.

The flight control system operation is conventional and incorporates a centre-mounted^[9] cyclic^[10] stick to control the attitude^[11] of the main rotor disc, and collective lever to control main rotor thrust. Tail rotor pitch and thrust is controlled by pilot movement of the tail rotor pedals. Tail rotor thrust is used to control the helicopter heading while in the hover, and balance while in forward flight.

Maintenance history

The helicopter log book statement indicated that the helicopter was to be maintained in accordance with the Robinson R44 maintenance manual, the pilots operating handbook (POH) for the daily/pre-flight inspection, Civil Aviation Safety Regulation 1998 Part 39 for airworthiness directives (ADs), and Civil Aviation Order 100.5 (general requirements in respect of maintenance of Australian aircraft). The logbook statement indicated the operational category of ‘charter’.

The last periodic maintenance inspection was completed on 7 September 2016, at 1,709 aircraft hours, in accordance with the Robinson R44 maintenance manual. Maintenance performed on the flight controls during the last periodic inspection was included in the wreckage examination and no anomalies were found. The current maintenance release^[12] was not located and likely destroyed in the accident fire. Airworthiness directives were tracked and certified at the periodic inspections. The helicopter was modified with a bladder fuel tank in April 2013.

Weight and balance

The ATSB performed weight and balance calculations using the Robinson R44 II POH and estimated weights, provided by the pilot and the passenger’s next-of-kin. Calculations were made with full fuel (128.9 kg) and empty fuel. Both calculations were within the centre of gravity limits. The maximum gross weight limit for the R44 II is 1,134 kg. The estimated weight when it departed Mareeba with a full fuel load was about 1,075 kg. The helicopter would have consumed about 25 kg of fuel on arrival at the Mount Windsor National Park, providing an estimated operating weight of about 1,050 kg.

Performance

The helicopter performance for hovering in-ground effect and out-of-ground effect^[13] was calculated using the environmental conditions reported by the rescue helicopter pilot (described below). This resulted in a density altitude equivalent to about 5,900 ft in a standard atmosphere. Using the estimated take-off weight from Mareeba Airport of about 1,075 kg, the in‑ground effect and out-of-ground effect performance figures were within limits. Maximum weight to hover out-of-ground effect at the accident site in nil wind was about 1,110 kg, which was slightly above the estimated operating weight of 1,050 kg on arrival at the park.

Meteorological information

Forecast conditions

The Bureau of Meteorology weather forecast for the area of operation, valid from 1100 until 2100 on 18 November, divided Area 45 into north-east and south-west subdivisions. The accident site was located in the north-east subdivision, where the wind was forecast as from 120 degrees at 25 kt at the altitudes of 2,000 ft and 5,000 ft above mean sea level. Moderate turbulence was also forecast below 8,000 ft on and about 30 NM lee of the eastern ranges north of Cooktown. The accident site in Mount Windsor National Park was about 50 NM south of Cooktown (outside of the forecast area for moderate turbulence), but located on the eastern ranges and in the same subdivision as the area to the north of Cooktown (refer to Appendix A).

Actual conditions

The rescue helicopter pilot reported the weather conditions on-site in the hover at the time of the search and rescue were a temperature of 26–28 °C at 3,500 ft, with gusting and varying wind at 15–25 kt and mostly clear. The rescue helicopter was a larger and more powerful helicopter than VH‑ZNZ and the rescue pilot reported that he had been ‘working hard’ to maintain a hover position over the accident site in ‘very turbulent’ conditions. When asked about the likely flying conditions for the accident helicopter, the rescue pilot commented that it ‘would have been working hard’.

One of the first responders to the report of the accident was from QPWS. He reported that the Mount Windsor National Park area is susceptible to strong and gusting wind conditions, and that those conditions were present on the day of the accident. Figure 3 depicts the local terrain about 500 ft below the elevation of the accident site with a view towards the south-east, the direction of the forecast wind.

Figure 3: Image of the local terrain with a view towards the south-east

View towards the south-east, the direction of the forecast wind, at about 3,000 ft elevation. The helicopter wreckage was located further inland behind the photographer at an elevation of about 3,500 ft, with surrounding peaks of about 4,500 ft. Source: ATSB

Table 1 lists the recorded wind conditions at 1230, about 14 minutes prior to the activation of the emergency locator beacon, from the surrounding Bureau of Meteorology observation sites.

Table 1: Surrounding weather observations

Mountain and lee wave activity

According to Underdown and Standen (2003), the following conditions are conducive to the formation of mountain and lee wave activity:

• Wind at right angles (or within +/- 30°) to a continuous mountain range
• Little change of wind direction with height
• Wind speed 15 kt^[14] or more at the mountain summit and increasing with height
• Very stable layer several thousand feet thick just above the mountains, with less stable air above and below the stable layer.

Immediately downwind from the mountain, rotors form with strong downdraughts and updraughts. Rotor streaming, which differs from rotor zones under mountain waves, may occur when there is a deep layer of strong winds across high ground with lighter winds above. Here, the turbulence in the rotor streaming occurs downwind of and level with the top of the high ground and for a considerable height above.

Underdown and Standen (2003) and the Bureau of Meteorology Aviation weather services both report that aircraft may encounter severe turbulence in mountain wave systems. Moderate turbulence indicates aircraft G-load^[15] variations of +/- 0.50–0.99 and severe turbulence indicates variations of +/- 1.0–1.99. The aircraft reaction to a moderate level of turbulence is described as: ‘Appreciable changes in attitude and/or altitude. Pilot remains in control at all times. Rapid bumps or jolts’. A severe level of turbulence is described as ‘Large abrupt changes in attitude and/or altitude. Momentary loss of control’.

Bureau of Meteorology comments

The Bureau of Meteorology reported that the winds on the area forecast are averaged over the 15-hour period of the forecast and over the subdivision area. They reported that ‘it is possible that there were other wind changes (direction and strength) not depicted in the forecast, which could have resulted in the forecast of turbulence north of Cooktown only’.

Wreckage and impact information

The ATSB did not conduct an initial on-site examination of the wreckage, but later conducted two targeted on-site visits following analysis of the evidence provided by the Queensland Police Service and QPWS staff. The first visit focused on examining the main rotor head, and the main and tail rotor systems. The second visit focused on the flight controls, engine and drive train.

Barracks facility

The accident pilot reported that he would not land the helicopter at the facility to be inspected if he assessed the location as unsuitable. In this situation, he would survey the local area for an alternative landing site nearby and the passenger would then hike to/from the facility to perform his inspections. The barracks facility was located a short walk east of a concrete causeway, which was large enough for an R44 to land on (Figure 4). The causeway was oriented east-west at a point where the river was oriented north-south. The accident site was in an upwind position relative to the barracks facility and causeway, consistent with the expected direction of travel if departing from the barracks.^[16] The gas compliance plate at the barracks facility was not annotated for the visit, but the Queensland Petroleum and Gas Inspectorate reported that it was not required to be annotated for an inspection only. The compliance plate is only annotated for installation or modification to an existing installation.

Figure 4: Accident site relative to barracks facility and causeway

Forecast wind direction indicating the accident site was located about 500–600 m upwind of the barracks facility and causeway.  Source: Google earth, annotated by ATSB

Site and distribution of wreckage

The accident site was at an elevation of about 3,500 ft and towards the south-east edge of a bowl with surrounding terrain up to about 4,500 ft. This placed it on the lee side of high terrain and 500‑600 m south-east of the Mount Windsor barracks facility. There was very dense vegetation with a forest canopy height of about 100 ft above ground level throughout the Mount Windsor National Park. The park is used for scientific research and closed to the public.

The helicopter was found to have broken into multiple fragments separated over an area of about 40 m x 50 m (refer to Appendix B). A separated section of tailcone, the empennage, tail rotor driveshaft sections and tail rotor blade debris were distributed to the west and south of the main wreckage. A post‑impact fire had consumed the main body of the helicopter, but the fragments separated from the main body were undamaged by fire. There were several tall trees surrounding the main wreckage, which had evidence of fire damage, but little impact damage. The damage to the forest canopy was concentrated directly above the main body of the helicopter, which indicated a near vertical final descent with little main rotor rotational energy (Figure 5).

Figure 5: Approach to canopy entry directly above wreckage site

Rescue helicopter view of the damaged canopy. Main body of the helicopter is located on the forest floor, about 100 ft directly below the canopy damage.  Source: Search and rescue helicopter service (courtesy Queensland Police Service), annotated by ATSB

The main wreckage included the helicopter frame assembly, engine, main rotor drive train and the main rotor. The helicopter frame was oriented left nose down in a south-west direction with high ground on the port (left) side. It showed evidence of significant burning as a result of a post-impact fuel-fed fire (Figure 6). Immediately beyond the main wreckage site, helicopter debris and flora were unburnt. The fire destroyed the main body and forward section of the tailcone. The aluminium and bladder fuel tanks had perished in the fire. The left landing gear skid tube extender was located in a position consistent with a left nose down impact.

Figure 6: Main wreckage

Source: Queensland Police Service, annotated by ATSB

Engine

The engine was examined in the impact position and then the main frame section was rotated so that the engine could be examined from the underside. An external examination of the engine found no evidence of pre-existing mechanical defects. The throttle and mixture linkages were connected from the cockpit through to the engine controls. The fuel control unit was examined and the mixture noted to be in the full rich position and the throttle full open (full power position).

The ATSB inspected the engine’s external components for evidence of rotational impact damage to determine if the engine was operating when the helicopter impacted the ground. One area of identified contact was between the flywheel and port oil cooler. RHC reported that flexing of the airframe on impact with the ground may push the port or starboard oil cooler into the flywheel, and that such damage did not necessarily indicate the flywheel was rotating. The on-site examination could not determine if the damage was cutting from rotation and therefore this evidence was considered inconclusive.

The majority of the fuel system was consumed by fire. However, both fuel tank caps were identified in the wreckage. They were found secured to the top sections of their respective fuel tank inlets. The fuel shut-off valve was not found. Remnants of the bladder fuel cells were found throughout the wreckage.

Tailcone and tail rotor system

The rear half of the tailcone was found about 25 m south of the main wreckage with evidence of a main rotor blade strike mark and yellow paint transfer from the rotor (Figure 7). The forward separation point of the tailcone had a tension failure (pulled apart at the rivet joint). RHC reported that in power-on situations, the main rotor blades will cut through the tailcone and driveshaft. In power-off situations, and/or rotor speed reduced, it will ‘smash’ the tailcone, pulling the driveshaft out, which was consistent with the wreckage.

Figure 7: Separated tailcone (left) and empennage (right)

Left: the rear tailcone assembly with strike mark and yellow paint transfer in and around the strike was found to the south of the main wreckage. Right: the empennage was found adjacent to the main wreckage but outside the fire zone with both tail rotor blades separated. Source: Queensland Police Service, annotated by ATSB

An onsite representative from the Queensland Police Service crash forensics unit noted the site had a dense canopy with a lot of trees between the separated tailcone and main body of the helicopter. This suggested to him that the tailcone had fallen through the canopy after separation, rather than being liberated from the main body inside the canopy. The empennage assembly (tail rotor gearbox with vertical and horizontal stabiliser) had torn laterally from the tailcone and was found adjacent to the main wreckage, but outside the fire zone, with both tail rotor blades separated near the hub.

There were no indications of a failure of the tail rotor gearbox, gearbox mounts or pitch links. Fracture and separation of the tail rotor blades was by overstress. Tail rotor debris was distributed on an arc of a radius of about 15 m around the main wreckage in an anti-clockwise direction from west-north-west through to the south. One tail rotor blade exhibited impact damage consistent with a main rotor blade strike (Figure 8).^[17] The other tail rotor blade tip cap was recovered and found to exhibit bending and yellow paint transfer that was consistent with main rotor blade contact (Figure 9). The outboard leading edges of the tail rotor blades were eroded and had split open. Wood fibres were found embedded within debris from both tail rotor blades.^[18]

Figure 8: Tail rotor blade impact damage

Tail rotor blade 1 with strike damage consistent with the leading edge of a main rotor blade. Source: ATSB

Figure 9: Tail rotor blade tip

Tail rotor blade 2 tip cap with yellow paint transfer from the main rotor and bending deformation. Source: ATSB

Main rotor system

The main rotor and mast were inspected for evidence of in-flight damage, including mast bumping^[19] and coning.^[20] One of the teeter stops^[21] was not present, likely to have been consumed in the post-accident fire. There was no evidence of damage to the mast where the teeter stop was previously located. The second teeter stop was in place, but exhibited fire damage. If mast bumping had occurred, it was not of a severity to result in visible mast damage. However, given the extent of the fire damage, the presence of mast bumping was inconclusive.

Some erosion of the main rotor blades leading edges was visible, but within limits. Both main rotor blades exhibited impact and fire damage. One main rotor blade and its associated pitch link assembly exhibited significant back-bending. RHC reported that deformation of the pitch link is typical following a main rotor blade strike. When a main rotor blade is bent and/or folded and still rotating, the loads on the pitch links and swashplate change dramatically, resulting in subsequent damage to main rotor head parts.

Neither of the main rotor blades exhibited coning (upward bending) damage, but both exhibited compression wrinkling of their upper surface. RHC reported that the angled creases from the leading edge to the trailing edge were indicative of backward bending of the rotor blades. The blade skin will retain the creases irrespective of whether the bending is elastic or plastic. Two marks on the upper surface of one of the main rotor blades indicated possible tail rotor blade strikes. Minor debonding of the main rotor blades was considered to be the result of the fire as there was no indication of associated bending from in-flight aerodynamic forces.

The main rotor blades, hub and shaft were found correctly assembled with no evidence of a pre‑existing defect.

Flight controls

The post-impact fire precluded examination and testing of the flight control system in its entirety. However, all of the flight control rod ends, attaching hardware and other steel components relating to the cyclic, collective and tail rotor controls were found intact and secure.

Most of the tail rotor pitch change controls were outside of the fire zone and were able to be examined. The tail rotor pitch link controls within the tailcone had fractured in overstress, likely as a result of the impact. The pitch links on the tail rotor itself moved freely and functioned as designed.

The dual controls for the cyclic, collective and tail rotor pedals, were not fitted at the time of the accident and were located in an area to the rear of the right rear passenger seat where they were likely to have been stored.^[22]

A detailed inspection of the available flight control system components did not identify any defects that may have contributed to the accident.

Drive train

The engine to main rotor transmission drive shaft was examined and no pre-accident defects were identified. The main rotor gearbox casing was consumed by the fire, but the single-stage spiral‑bevel gear and mast roller bearings were visible and did not indicate any evidence of a catastrophic failure. The forward flex plate bolts exhibited minor bending on the main gearbox side, but not on the engine side (Figure 10). RHC reported that the deformation of the forward flex plate is consistent with damage from extreme angles between the yokes while being pulled apart. As the main rotor blades contact solid objects the mast will move fore, aft, left and right, which changes the angle of the input yoke of the gearbox. The entire airframe will flex, resulting in the yokes at all three flex couplings moving apart. They noted that in this case the forward flex plate bending was relatively minor.

Figure 10: Forward flex plate (left) and intermediate flex plate (right)

Left: forward flex plate with visible bending of bolts on the main gearbox side, but not on the engine side. Right: intermediate flex plate exhibited tension failure, but no significant torsion. The V-belt drive from the engine was located between the two flex plates. Source: ATSB

The tail rotor drive shaft had fractured into five sections between the intermediate and aft flex plate. Both flex plates had failed in tension (pulled apart) and the driveshaft pulled through the damper assembly. The separation points of the tail rotor drive shaft were reviewed and it was found that they displayed bending failures, with the exception of a section that was seized within the separated tailcone.

The section within the tailcone exhibited a torsional failure forward (facing towards engine drive) and a bending failure aft (facing towards the tail rotor gearbox). No paint transfer was found on any of the tail rotor driveshaft sections and the flex plates did not exhibit any significant torsional damage associated with their failure in tension. RHC reported there was very little indication of rotational scoring on the tail rotor driveshaft sections, which indicated to them that there was low drive speed at the time the flex plates were pulled apart.

Continuity of the drive train could not be established due to fire damage to the V-belts and clutch assembly. However, the material available did not exhibit any pre-existing defects that may have contributed to the accident.

Figure 11 depicts the approximate position of the separation points. Of note, the main rotor cyclic rigging setting is 13.5–14.25°. The blade strike mark was at an angle of about 17–18°.

Figure 11: Approximate positions of tailcone separation

Main rotor blade strike and associated tailcone separation points. 
Source: Robinson Helicopter Company, annotated by ATSB

Survival aspects

The pilot was seated in the front right seat and the passenger in the front left seat. The orientation of the wreckage was left nose down with the left skid buried into the ground slightly aft of the airframe. This suggested a left nose down impact with sufficient force to separate the left skid from the airframe. The post-mortem results for the passenger determined that the likely cause of death was crash‑associated multiple injuries. The results of testing for drugs and poisons, including alcohol, were negative.

Crashworthiness

The certification standard for the R44 was based upon providing the occupant(s) with a reasonable chance of escaping serious injury in a minor crash based upon the rotorcraft absorbing the landing loads with an ultimate descent velocity of five feet per second. The following acceleration limits applied:

• upward – 1.5 G
• forward – 4.0 G
• sideward – 2.0 G
• downward – 4.0 G.^[23]

The landing gear limit load drop test requirements for the helicopter landing gear were as follows:

(1) 13 inches from the lowest point of the landing gear to the ground; or

(2) Any lesser height, not less than eight inches, resulting in a drop contact velocity equal to the greatest probable sinking speed likely to occur at ground contact in normal power-off landings.^[24]

The fuel tank drop test requirements were as for the occupant acceleration limits. The descent path of the helicopter and orientation of the wreckage indicated that the ground impact was likely outside the certification limits, which severely compromised the liveable volume^[25] for the front left seat occupant and the integrity of the fuel tanks. The use of seat belts could not be determined from the fire damage, but are not considered likely to have influenced the outcome for the passenger.

Search and rescue

The ATSB found no evidence that a distress call was made by the pilot prior to the detection of the emergency locator transmitter (ELT). The JRCC first detected the helicopter’s ELT at 1244. Following detection of the ELT, search and rescue aircraft were directed to the vicinity of the signal. The last signal from the ELT was detected by the JRCC at 1351.^[26] When the rescue helicopter arrived, the rescue helicopter pilot observed smoke, which was used to locate the main wreckage site. A rescue wire-person was winched down to the site and confirmed they had located the accident site. The confirmation was received by the JRCC at 1432.

The rescue wire-person found the pilot during a brief search of the area and confirmed his identification with his aviation security identification card. The pilot was initially conscious and reported that he was alone and had dropped his passenger off. However, his condition deteriorated and a doctor was winched down to the site to conduct an assessment. The pilot, wire‑person and doctor were recovered to the helicopter and the pilot was transferred to Cairns hospital. A search for the passenger in the surrounding area by air and ground continued until last light without success. The following day, 19 November, a police forensic team was winched down to the accident site to locate and retrieve the deceased passenger from within the main wreckage.

In-flight break-up research

There have been a number of main rotor divergence and in-flight break-up accidents involving the Robinson family of helicopters. This includes the R22, R44 and R66^[27] helicopters. Previous R22 and R44 main rotor divergence accidents were examined by the United States National Transportation Safety Board in a 1996 special investigation report.^[28] The ATSB reviewed their report and researched previous R44 and R66 main rotor divergence accidents, which resulted in an in-flight break-up with no evidence of a pre-existing defect.^[29] This was for the purpose of a comparative analysis with the damage found to VH-ZNZ and resulted in a review of 12 historical accidents from Germany, the United States, France, Canada and New Zealand.

All historical cases reported evidence of mast bumping. Several cases included mast bending (3 of 12) and/or mast sheared (5 of 12).^[30] The Transport Accident Investigation Commission of New Zealand’s report AO-2013-003 included comments from a metallurgical and fractographic^[31] examination of the mast failure as follows:

Failure occurred by overload and the nature of the fracture indicates that significant bending and torsional loads were applied…The direction of torsional loading was consistent with power being applied to the rotor while the rotor was abruptly decelerated, perhaps through contact with the airframe or some other object.

The inspection of VH-ZNZ found no evidence of mast bending. Due to the fire damage to the teeter stops, evidence of mast bumping was inconclusive. There was no evidence of mast damage where the fire had consumed one teeter stop and therefore no evidence of excessive mast bumping.

Review of the rotor strikes to the airframes were limited by the amount of detail provided in the investigation reports and the various forms of description provided. Some reports included the description of tailcone ‘severed in-flight’, while others referred to it as ‘separated’. In some cases, the main rotors struck and cut through the cabin in addition to, or instead of the tailcone. While VH-ZNZ did exhibit a tailcone strike, the main rotors did not shear (cut through) the tailcone and there was no indication that the main rotors struck the cabin.

The historical cases were all fatal with no survivors. In several cases, the reports concluded that the accident was associated with pilot over-control in a low-G flight condition, which is described by RHC in their POH safety notice 11.^[32] In other cases the reports concluded that the reason for the main rotor divergence could not be determined.

Additional information

Pilot operating handbook

The Robinson R44 II POH section 10: Safety tips and notices, included the following safety notice information of interest to the investigation:

Safety notice SN-10: Fatal accidents caused by low RPM rotor stall

Power available from the engine is directly proportional to RPM. If the RPM drops 10%, there is 10% less power. With less power, the helicopter will start to settle, and if the collective is raised to stop it from settling, the RPM will be pulled down even lower, causing the ship to settle even faster. If the pilot not only fails to lower collective, but instead pulls up on the collective to keep the ship from going down, the rotor will stall almost immediately. When it stalls, the blades will either “blow back” and cut off the tailcone or it will just stop flying, allowing the helicopter to fall at an extreme rate.

Safety notice SN-24: Low RPM rotor stall can be fatal

Rotor stall due to low RPM causes a very high percentage of helicopter accidents, both fatal and non-fatal… As the RPM of the rotor gets lower, the angle of attack^[33] of the rotor blades must be higher to generate the lift required to support the weight of the helicopter. As with the airplane wing, the blade airfoil will stall at a critical angle, resulting in a sudden loss of lift and a large increase in drag.

When the rotor stalls, it does not do so symmetrically because any forward airspeed of the helicopter will produce a higher airflow on the advancing blade than on the retreating blade. This causes the retreating blade to stall first, allowing it to dive as it goes aft… Also, as the helicopter begins to fall, the upward flow of air under the tail surfaces tends to pitch the aircraft nose-down. These two effects, combined with aft cyclic by the pilot attempting to keep the nose from dropping, will frequently allow the rotor blades to blow back and chop off the tailboom as the stalled helicopter falls.

Safety notice SN-32: High winds or turbulence

Flying in high winds or turbulence should be avoided.

A pilot’s improper application of control inputs in response to turbulence can increase the likelihood of a mast bumping accident. If turbulence is encountered, the following procedures are recommended:

1. Reduce power and use a slower than normal cruise speed. Mast bumping is less likely at lower airspeeds.

2. For significant turbulence, reduce airspeed to 60–70 knots.

3. Tighten seat belt and rest right forearm on right leg to minimize unintentional control inputs. Some pilots may choose to apply a small amount of cyclic friction to further minimize unintentional inputs.

4. Do not overcontrol. Allow aircraft to go with turbulence, then restore level flight with smooth, gentle control inputs. Momentary airspeed, heading, altitude, and RPM excursions are to be expected.

5. Avoid flying on the downwind side of hills, ridges, or tall buildings where turbulence will likely be most severe.

The helicopter is more susceptible to turbulence at light weight. Reduce speed and use caution when flying solo or lightly loaded.

Safety notice 10 described the risk of a main rotor blade stall associated with an overpitching event, which is discussed further in the Over-pitching section of the report below. Safety notice 24 provided a more generic description of low rotor speed stall, which is discussed further under Loss of drive power below.

Safety notice 32 replaced an earlier airworthiness directive, issued by the FAA in 1995 for the R44 helicopter. The directive prohibited flight in surface winds greater than 25 kt, gusting winds greater than 15 kt, and in moderate, severe or extreme turbulence. These limitations have remained in place for low-experience R22 pilots (less than 200 flight hours in helicopters and less than 50 in the R22).

The ATSB enquired as to whether a handling assessment in moderate turbulence was ever conducted in the R44 and, if performed, what rating the helicopter received.^[34] In reply, RHC reported that no additional assessment with quantitative scoring was performed besides the requirements for certification.

Over-pitching

The International Civil Aviation Organization (ICAO) manual of aircraft accident and incident investigation, chapter 15: Helicopter investigation, described over-pitching as a phenomena that happens when collective pitch is increased to a point where the main rotor blade angle of attack creates so much drag that all available engine power cannot maintain or restore normal operation rotor speed. At low rotor speed, the rotor blades bend upwards and drag increases. The high inflow angles and rotor drag quickly decay main rotor speed, which may decrease to the point where the main rotor blades stall as described in R44 safety notice 10.

The ICAO investigation manual noted that over-pitching could occur in any phase of powered flight and that high weight, high density altitude and high temperatures are contributing factors. The manual also noted that over-pitching to low rotor speed and blade stall is greatest in small piston engine helicopters with low rotor blade inertia, and that with a piston engine, the engine speed will also decrease and ‘may cease operation during over-pitching’.^[35] RHC reported that they were aware that low rotor speed could result in the engine stalling. They reported it was more likely in the R22, but in the right conditions (high altitude, high temperature and high loading) the R44 can be over‑pitched enough to stall the engine.

Previous over-pitching events

ATSB investigation 200600979: Collision with terrain 10 km west of Gunpowder Mine, Qld, 21 February 2006, VH-HBS Robinson Helicopter Company R44, concluded that the accident was the result of insufficient main rotor thrust that was a consequence of low main rotor speed. The accident site was about 3,800 ft density altitude (assuming standard atmospheric pressure) with moderate thermal turbulence below 8,000 ft forecast.

ATSB investigation AO-2008-062: Collision with terrain – Robinson Helicopter R44 Raven, VH‑RIO, 6 km NE Purnululu ALA, Western Australia, 14 September 2008, concluded that the most likely scenario was that the engine power required exceeded the engine power available, resulting in main rotor speed decay. The helicopter was operating at about 3,500 ft density altitude and the investigation found that moderate thermal turbulence was likely present below 9,000 ft.

Loss of drive power

The ICAO investigation manual states that, in the event of a sudden engine stoppage, pilot reaction time is a significant factor to prevent a rapid decay of main rotor speed, and that a low inertia rotor system will decrease speed more rapidly than high inertia rotors. Lowering the collective lever enables the pilot to maintain rotor speed on entry to an autorotation,^[36] but the last 100 ft are equally critical.

In the last 100 ft the pilot must reduce the forward speed and rate of descent to minimise the horizontal and vertical accelerations for landing. This manoeuvre is performed by the pilot flaring the helicopter (pitching the nose up) and raising the collective lever, which increases the rotor thrust to reduce the rate of descent and forward speed. Raising the collective lever without drive power will also rapidly decay the rotor speed. If it decays too low prior to landing, a low rotor speed blade stall as described in R44 safety notice 24 may occur. Drive power for the R44 is dependent on the performance of the engine and its governor system, and the continuity of the drive train, including the V-belts, clutch, gearboxes and driveshafts. The examination of the wreckage found no pre-existing defects with the gearboxes or driveshafts, but the impact and fire damage prevented examinations of the other components.

Previous loss of drive power events (powerplant or drive train faults)

A search of the ATSB’s database for engine failure or malfunction occurrences involving the R44 II found 10 events between 2012 and 2017.^[37] A similar search for transmission and gearbox issues identified three occurrences. The results from the two searches were mutually exclusive.

The reasons for the transmission faults included one undetermined (ATSB investigation AO-2016-172), one required a clutch unit replacement, and one revealed the V-belts were loose. For AO‑2016-172, the helicopter’s rotor speed decayed in-flight, but engine speed was maintained. Following recovery of the helicopter, no pre-existing defect was found and the fault could not be reproduced.

__________

5. The R44 was a type rating in the initial introduction of Civil Aviation Safety Regulation 1998 Part 61, but was later moved to the single-engine class rating in January 2015.
6. Normal category applies for rotorcraft with maximum weights of 7,000 pounds or less and nine or less passenger seats.
7. Collective: a primary helicopter flight control that simultaneously affects the pitch of all blades of a lifting rotor. Collective input is the main control for vertical velocity.
8. Coning of main rotor blades: the upwards movement of the main rotor blades while they are rotating. This is usually in response to an increase in aerodynamic force as a result of a control input from the pilot. It is more pronounced at high weights and/or low main rotor speed.
9. The centre-mounted cyclic stick is a unique design feature to the Robinson helicopter family.
10. Cyclic: a primary helicopter flight control that is similar to an aircraft control column. Cyclic input tilts the main rotor disc, varying the attitude of the helicopter and hence the lateral direction.
11. Attitude refers to the comparison of the rotor disc or helicopter axes with the local horizon.
12. Maintenance release: an official document, issued by an authorised person as described in Regulations, which is required to be carried on an aircraft as an ongoing record of its time in service (TIS) and airworthiness status. Subject to conditions, a maintenance release is valid for a set period, nominally 100 hours TIS or 12 months from issue.
13. Helicopters require less power to hover when ‘in ground effect’ than when ‘out- of ground effect’ due to the cushioning effect created by the main rotor downwash striking the ground. The height of ‘ground effect’ is usually defined as less than one main rotor diameter above the surface, which is about 33 ft for the R44 helicopter.
14. A higher value of 25 kt may be reported by other publications.
15. G-load: the nominal value for acceleration. In flight, g load represent the combined effects of flight manoeuvring loads and turbulence and can have a positive or negative value.
16. A departure into wind would improve the helicopter’s performance by reducing the power required and ground speed, and improving the climb performance, when compared with a tailwind. This is more critical as the environmental conditions become more adverse, such as stronger winds and higher density altitudes.
17. RHC reported that the damage to the tail rotor blade in Figure 8 is more consistent with the tail rotor blade striking a tree than a main rotor blade striking the tail rotor blade. However, their response did not account for the tip bending and paint transfer evident in Figure 9, and the ATSB noted a scratch in the yellow paint of a main rotor blade, near the main rotor blade tip, consistent with contact with a tail rotor blade. Footnote 19 refers to an unsuccessful attempt to verify if the tail rotor blades did strike the same tree.
18. The wood fibres were sent to the University of Adelaide, School of Biological Sciences, for genetic testing to determine if both tail rotor blades struck the same tree. However, there was insufficient material for a successful test. Therefore, no conclusion as to the nature of the break-up has been made from their presence.
19. Mast bumping: contact between the main rotor hub and the rotor mast which, if excessive, could severely damage the mast, or result in the separation of the main rotor system from the helicopter. Damage from mast bumping is indicative of excessive blade flapping and/or excessive tilt of the main rotor disc relative to the mast.
20. Coning damage may include upward bending of the main spar and/or compression wrinkling damage to the upper surface of the main rotor blade.
21. Two elastomeric stops are fitted at the top of the main rotor mast to protect the mast from direct contact with the main rotor blades. Excessive teetering of the main rotor blades in-flight will result in the main rotor blades striking the teeter stops.
22. In accordance with the R44 pilot operating handbook, the left seat controls must be removed if the person in that seat is not a rated helicopter pilot.
23. US Code of Federal Regulations Part 27.561 (2 October 1964) – Emergency Landing Conditions.
24. US Code of Federal Regulations Part 27.725 (2 October 1964) – Landing Gear.
25. Flight Safety Foundation 1989, Helicopter Safety: Helicopter crashworthiness – part one, Vol. 15 No. 6.
26. The emergency locator transmitter was installed in the main body of the helicopter, which was subject to the post-crash fire.
27. The Robinson R66 is the five seat turbine-engine helicopter using the same T-bar cyclic and teetering main rotor hub design features as the R22 and R44, approved for normal category rotorcraft on 25 October 2010.
28. United States National Transportation Safety Board 1996, Special investigation report: Robinson Helicopter Company R22 loss of main rotor control accidents, NTSB, Washington.
29. These accidents were found to have occurred under normal power and rotor speed conditions.
30. Results included both R44 and R66 helicopters.
31. Fractography is the examination of the cause of a material failure by studying the characteristics of the fracture surface.
32. R44 safety notice SN-11: Low-G pushovers – extremely dangerous: Pushing the cyclic forward following a pull-up or rapid climb, or even from level flight, produces a low-G (weightless) flight condition. If the helicopter is still pitching forward when the pilot applies aft cyclic to reload the rotor, the rotor disc may tilt aft relative to the fuselage before it is reloaded… With no lift from the main rotor…mast bumping can occur. Severe in-flight mast bumping usually results in main rotor shaft separation and/or main rotor contact with the fuselage… Always use great care to avoid any manoeuvre which could result in a low-G condition. Low-G mast bumping accidents are almost always fatal.
33. The angular difference between the chord of the rotor blade and the relative airflow. The chord is the straight line between the rotor blade leading edge and trailing edge.
34. Harper RP & Cooper GE 1984, Handling qualities and pilot evaluation, manuscript prepared for the 1984 Wright Brothers Lectureship in Aeronautics. The manuscript provided a recommended process for pilots to evaluate and rate aircraft handling qualities and was the reference source for the enquiry.
35. As engine speed decreases at full throttle, a maximum torque value will be passed. The area below the maximum torque value is referred to as the ‘low end’ torque zone, or ‘unstable torque region’. An increase in resistance in this area will result in a drop in engine torque and engine stall.
36. Descent with power-off. Air flowing in the reverse direction upwards through the rotors produces the driving force for the rotors.
37. The results prior to 2012 were excluded as some did not differentiate between the different models of R44 helicopters.

Safety analysis

Introduction

While operating in the Mount Windsor National Park on a charter flight to provide gas and plumbing inspection services, the Robinson R44 helicopter broke-up after the main rotor struck the helicopter’s tailcone. As a result of the in-flight break-up and impact, the passenger was fatally injured and the pilot sustained serious injuries.

No evidence was found of a pre-existing mechanical defect with the helicopter. However, no tests or internal inspections were performed on the engine or clutch, which were subject to impact and fire damage in a remote location. In addition, the drive train V-belts had perished, which prevented the verification of drive train continuity.

The accident flight was within the published weight and balance limits for the helicopter, which had sufficient power to hover out-of-ground effect, in nil wind. Turbulence may have reduced the power margin, which is discussed further under: Possible scenarios precipitating the in-flight break-up.

This analysis will discuss the weather conditions on the day, in particular, the possibility of turbulence, which was not forecast, and its potential effect on the flight. It also considers several scenarios for the in-flight break-up and highlights the benefits of having an emergency locator transmitter fitted for survivability.

Weather conditions

Moderate turbulence was forecast to the north of the Mount Windsor National Park for the time of the accident flight, but not for the park area. However, the park was subject to the same forecast wind strength of 25 kt as the area to the north. Given the wind direction and strength on the day, combined with the undulating terrain and tree canopy, the conditions in the area were conducive to terrain‑induced turbulence. This was consistent with the reports from personnel who responded to the accident. While the degree of turbulence could not be determined, in consideration of the reports provided and the Bureau of Meteorology’s turbulence intensity descriptions, it was likely to have been at least moderate.

In the absence of a forecast for turbulence the pilot may not have had a complete appreciation of how the weather would affect flying conditions in an R44 in the national park environment. However, as the pilot could not recall the events on the day of the accident, his understanding of the conditions could not be established.

The Robinson R44 safety notice SN-32 advised pilots to avoid flying in high winds or turbulence, which could result in fluctuating G-conditions and mast bumping, as described in safety notice SN‑11. This notice, SN-32, replaced an earlier airworthiness directive, which prohibited flight under such conditions. Those conditions were likely present at the time of the accident. Although the airworthiness directive had been cancelled, RHC reported that they had no handling qualities assessment for the helicopter in conditions of moderate turbulence (nor were they required to for certification). Therefore, the likely presence of at least moderate turbulence potentially placed the helicopter in an environment for which it had not been flight tested. Despite this, the ATSB could not determine if the weather contributed to the accident.

In-flight break-up sequence

The break-up of the helicopter likely started with a main rotor blade striking and liberating the tailcone from the airframe. This likely occurred above the forest canopy due to the separation of the tailcone from the main body wreckage on the forest floor and the density of the forest between them. The strike occurred while the tail rotor driveshaft was rotating, as indicated by a section of driveshaft with a torsional failure. The bending failures of the tail rotor driveshaft in lieu of shearing, and absence of torsional damage to the flex plates and rotational scoring to the driveshaft, indicated there was likely low engine power and rotor speed at the time of the tailcone strike.

Following the tailcone strike, the tail rotor driveshaft was pulled apart at the intermediate and aft flex plates and the empennage was torn laterally from the aft portion of tailcone due to its own inertia. This presented the tail rotor blades within striking distance of the main rotor disc path, which struck and liberated both tail rotor blades. The helicopter then descended on a near vertical trajectory through the forest canopy and impacted the forest floor left nose down relative to the ground.

The distribution of the tail fragments to the west and south of the main body was consistent with the helicopter moving on an easterly track at the time of the break-up. The elapsed time between the helicopter’s track towards the park and activation of the emergency locator beacon, in association with the location of the accident site and distribution of wreckage, indicated the accident more likely than not occurred during an upwind departure from the barracks facility. The relatively small distribution area of the wreckage and survival of the pilot indicated that the break‑up likely occurred close to the forest canopy without a high rate of descent.

A review of previous R44 and R66 helicopter in-flight break-ups at normal operating rotor speeds and powers revealed notable differences with this accident. Specifically, the tailcone strike damage and absence of any significant mast damage did not support a mast bumping scenario from the pilot overcontrolling the cyclic at normal engine power and rotor speed. Historically, these accidents have resulted in significant damage to the main rotor mast, including bending and fracture, which was not exhibited by this accident. The low engine power and rotor speed, and absence of any significant mast damage, indicated the strike was likely the result of a low rotor speed blade stall, which is discussed below.

Possible scenarios precipitating the in-flight break-up

In consideration of the environmental conditions and helicopter damage, the ATSB considered two possible scenarios, which could result in a low rotor speed blade stall.

Loss of drive power (powerplant or drive train fault)

As the helicopter climbed from near sea level into the Mount Windsor National Park it was exposed to an increasing density altitude (reduction in ambient air density). As a result of the increasing density altitude the margin between the power available and power required would have reduced with high rotor drag. The R44 pilot operating handbook indicated that the throttle is frequently wide open when operating at altitudes above 3,000‑4,000 ft. This would result in a rapid reduction of rotor speed in the event of a loss of drive power.

A loss of drive power at high power, low height and low speed, such as on take-off and initial climb, would result in very little energy available to trade-off, in order to recover from a loss of rotor speed. A significant loss of rotor speed while attempting to arrest the rate of descent before entering the forest canopy could lead to a main rotor blade stall and tailcone strike as described in the R44 pilot operating handbook safety notice 24. Of note, a loss of rotor speed due to a drive train fault with the V-belts or clutch could result in a low power tailcone strike in-flight followed by flywheel rotational damage to the port oil cooler on impact with the ground.

Overpitching

The accident site was located on the lee side of a bowl surrounded by higher terrain. The lee side of terrain is the highest risk area for encountering a downdraught in strong wind conditions as described in R44 safety notice 32. If mountain or lee wave activity was present, then tracking into wind towards rising terrain at a high-density altitude could potentially result in the helicopter flying into a downdraught, which exceeds the power available. This can result in an overpitching event followed by a blade stall and tailcone strike as described in the R44 pilot operating handbook safety notice 10.

A blade stall and tailcone strike while overpitching is expected to exhibit significant engine power at the time of the strike. However, if the engine stalled while overpitching before a main rotor blade stalled, then this would result in a tailcone strike, which exhibited low power and low rotor speed.

There was no evidence of upward bending (coning) of the main rotor blades indicative of overpitching. Further, the compression wrinkling present on the main rotor blades exhibited backward bending from a strike, rather than upward bending from coning. Therefore, although the reported environmental conditions of high-density altitude combined with turbulence indicated that the helicopter was susceptible to overpitching, the damage found to the helicopter did not provide conclusive support for this scenario.

Summary

The absence of any significant mast bumping damage combined with the apparent low power tailcone strike indicated a likely main rotor blade stall strike. This would be consistent with the pilot attempting to arrest the helicopter’s rate of descent before entering the forest canopy, which is itself consistent with the pilot surviving the accident.

A loss of rotor speed presents several scenarios, including overpitching, a loss of engine power (including an engine stall), or a loss of drive continuity. None of these scenarios could be verified. Therefore, the scenario, which resulted in the main rotor blade stall, could not be determined.

Emergency locator transmitter

Historical accidents of this nature, involving an in-flight break-up, generally result in fatal injuries for all occupants on board. The survival of the pilot in this accident was a rare occurrence and was likely due to the reduced rate of descent as the helicopter entered the canopy combined with the orientation of the airframe at impact. However, for the passenger in the front left seat, the orientation of the airframe at impact resulted in high energy attenuation in this location and was therefore considered not to be survivable.

The three key elements for the rescue of the pilot were the automatic activation of the emergency locator transmitter, the response of the Australian Maritime Safety Authority’s Joint Rescue Coordination Centre and the availability of the search and rescue helicopter. The activation of the emergency locator transmitter on impact was the trigger to start the search and rescue. This resulted in the discovery of the accident site by the search and rescue helicopter about 1 hour and 48 minutes after the initial detection of the emergency locator transmitter. The pilot may otherwise have not survived.

ATSB investigation AO-2017-033: Collision with terrain involving Agusta AB206, VH-DPU, found that the rescue of the occupants took about 39 hours. They were carrying a manually activated emergency position indicating radio beacon, which they were physically unable to retrieve and activate after the accident. They were eventually located with the assistance of OzRunways recorded track data.

Limitations on the effectiveness of emergency locator transmitters in aviation accidents were reviewed by the ATSB in research report AR-2012-128. The report identified several performance issues, which could affect their operational reliability, but still credited them for saving an average of four lives per year. While some of the reliability issues were considered beyond the control of the individual pilot, such as damage during impact, the report identified several measures within the control of the individual to improve their performance. In this accident, smoke emitted from the post-accident fire enabled the rescue helicopter to locate the accident site. However, according to the report (AR-2012-128), a global positioning system enabled transmitter can improve the accuracy of detection from about 5 km to 120 m.

Findings

From the evidence available, the following findings are made with respect to the in-flight break-up involving a Robinson R44, registered VH-ZNZ, 41 km north-west of Mossman, Queensland, on 18 November 2016. These findings should not be read as apportioning blame or liability to any particular organisation or individual.

Contributing factors

• The helicopter main rotor diverged from the normal plane of rotation and struck the tailcone, which resulted in the in-flight break-up of the helicopter.
• The main rotor strike to the tailcone was likely the result of a low main rotor speed blade stall for reasons undetermined.

Other safety factors

• The helicopter likely encountered at least moderate turbulence, which was not forecast. It was approved to operate in the prevailing weather, but was not flight tested for those conditions.

Other findings

• The fitment and registration of the emergency locator transmitter to the helicopter assisted the search and rescue of the pilot.
• The helicopter was within the weight and balance, and performance limits for the planned flight.
• Although the post-impact fire precluded examination of the helicopter in its entirety, a detailed inspection did not identify any pre-existing defects that may have contributed to the accident.

Sources and submissions

Sources of information

The sources of information during the investigation included the:

• Australian Maritime Safety Authority (JRCC)
• Bureau of Meteorology
• Civil Aviation Safety Authority
• OzRunways
• Pilot (also the operator)
• Queensland Parks and Wildlife Service
• Queensland Petroleum and Gas Inspectorate
• Queensland Police Service
• Rescue helicopter pilot
• Robinson Helicopter Company
• University of Adelaide (School of Biological Sciences).

References

Australian Transport Safety Bureau 2013, AR-2012-128: A review of the effectiveness of emergency locator transmitters in aviation accidents, ATSB, Canberra.

Bureau of Meteorology aviation weather services. Hazardous weather phenomena: Turbulence; retrieved from www.bom.gov.au on 3 August 2017.

Flight Safety Foundation 1989, Helicopter Safety: Helicopter crashworthiness – part one, Vol. 15 No. 6.

Harper RP & Cooper GE 1984, Handling qualities and pilot evaluation, manuscript prepared for the 1984 Wright Brothers Lectureship in Aeronautics.

International Civil Aviation Organization 2011, Manual of aircraft accident and incident investigation Part III: Investigation, Doc 9756, ICAO, Montréal.

Underdown RB & Standen J 2003, Meteorology (3rd Ed.), Blackwell Publishing, Oxford.

United States National Transportation Safety Board 1996, Special investigation report: Robinson Helicopter Company R22 loss of main rotor control accidents, NTSB, Washington.

Transport Accident Investigation Commission (New Zealand) 2016, Aviation inquiry AO-2013-003: Robinson R66, ZK-IHU, Mast bump and in-flight break-up, Kaweka Range, 9 March 2013, TAIC, Wellington.

Submissions

Under Part 4, Division 2 (Investigation Reports), Section 26 of the Transport Safety Investigation Act 2003 (the Act), the Australian Transport Safety Bureau (ATSB) may provide a draft report, on a confidential basis, to any person whom the ATSB considers appropriate. Section 26 (1) (a) of the Act allows a person receiving a draft report to make submissions to the ATSB about the draft report.

A draft of this report was provided to the Australian Maritime Safety Authority, Bureau of Meteorology, Civil Aviation Safety Authority, the pilot, Queensland Parks and Wildlife Service, Queensland Petroleum and Gas Inspectorate, Queensland Police Service, rescue helicopter pilot, Robinson Helicopter Company, and the United States National Transportation Safety Board.

The submissions from those parties were reviewed and where considered appropriate, the text of the draft report was amended accordingly.

Pilot and aircraft details

Pilot details

Aircraft details

Appendices

Appendix A – Planning chart with weather subdivisions

Appendix B – Distribution of the helicopter wreckage (North-Up)

Terminology used in this report

Occurrence: accident or incident.

Safety factor: an event or condition that increases safety risk. In other words, it is something that, if it occurred in the future, would increase the likelihood of an occurrence, and/or the severity of the adverse consequences associated with an occurrence. Safety factors include the occurrence events (e.g. engine failure, signal passed at danger, grounding), individual actions (e.g. errors and violations), local conditions, current risk controls and organisational influences.

Contributing factor: a safety factor that, had it not occurred or existed at the time of an occurrence, then either: (a) the occurrence would probably not have occurred; or (b) the adverse consequences associated with the occurrence would probably not have occurred or have been as serious, or (c) another contributing factor would probably not have occurred or existed.

Other factors that increased risk: a safety factor identified during an occurrence investigation, which did not meet the definition of contributing factor but was still considered to be important to communicate in an investigation report in the interests of improved transport safety.

Other findings: any finding, other than that associated with safety factors, considered important to include in an investigation report. Such findings may resolve ambiguity or controversy, describe possible scenarios or safety factors when firm safety factor findings were not able to be made, or note events or conditions which ‘saved the day’ or played an important role in reducing the risk associated with an occurrence.

Safety issue: a safety factor that (a) can reasonably be regarded as having the potential to adversely affect the safety of future operations, and (b) is a characteristic of an organisation or a system, rather than a characteristic of a specific individual, or characteristic of an operational environment at a specific point in time.

Safety action: the steps taken or proposed to be taken by a person, organisation or agency in response to a safety issue.

Purpose of safety investigations The objective of a safety investigation is to enhance transport safety. This is done through:  identifying safety issues and facilitating safety action to address those issues providing information about occurrences and their associated safety factors to facilitate learning within the transport industry. It is not a function of the ATSB to apportion blame or provide a means for determining liability. At the same time, an investigation report must include factual material of sufficient weight to support the analysis and findings. At all times the ATSB endeavours to balance the use of material that could imply adverse comment with the need to properly explain what happened, and why, in a fair and unbiased manner. The ATSB does not investigate for the purpose of taking administrative, regulatory or criminal action. Terminology An explanation of terminology used in ATSB investigation reports is available here. This includes terms such as occurrence, contributing factor, other factor that increased risk, and safety issue. Publishing information Released in accordance with section 25 of the Transport Safety Investigation Act 2003 Published by: Australian Transport Safety Bureau © Commonwealth of Australia 2019 Ownership of intellectual property rights in this publication Unless otherwise noted, copyright (and any other intellectual property rights, if any) in this report publication is owned by the Commonwealth of Australia. Creative Commons licence With the exception of the Coat of Arms, ATSB logo, and photos and graphics in which a third party holds copyright, this publication is licensed under a Creative Commons Attribution 3.0 Australia licence. Creative Commons Attribution 3.0 Australia Licence is a standard form licence agreement that allows you to copy, distribute, transmit and adapt this publication provided that you attribute the work. The ATSB’s preference is that you attribute this publication (and any material sourced from it) using the following wording: Source: Australian Transport Safety Bureau Copyright in material obtained from other agencies, private individuals or organisations, belongs to those agencies, individuals or organisations. Where you wish to use their material, you will need to contact them directly.

What happened

On 5 October 2016, a Qantas-operated Boeing 747-438 aircraft, registered VH-OJT (Figure 1) was undergoing routine maintenance at a contracted repair and overhaul facility in the Hong Kong Special Administrative Region (SAR) in the People’s Republic of China. Inspection of the No. 2^[1] engine pylon identified cracking of four outboard strut ribs from within the torque box of the pylon (Figure 2).

Following technical advice from Boeing, the No. 2 engine was removed from its wing installation to allow access to the pylon and removal of the cracked strut ribs. The aircraft was repaired and, after completing the maintenance visit, it was returned to regular passenger service. The cracked ribs were sent to Boeing’s facilities in the United States for specialist metallurgical examination.

Figure 1: Qantas Boeing 747-438 aircraft, registered VH-OJT

Source: Christopher Chai

Pylon inspection requirements

The engines on a Boeing 747 aircraft are mounted to pylons located on the underside of each wing. The pylons transmit thrust from the engine to the airframe and are designed to withstand flight loads from normal operation along with transient dynamic loading in the event of an engine failure. Hydraulic, fuel, electrical and air conditioning lines all pass through the pylon structure.

Boeing’s ongoing maintenance requirements included an inspection of the engine pylon every 48 months. A general visual inspection within the torque box for corrosion damage was required along with a more detailed inspection for cracking of the ribs. The procedures noted that during the inspection, particular attention should be applied to the cutaway where stringers passed through each rib.

Qantas reported to the ATSB that several years prior to this occurrence, it had independently increased the frequency of the zonal inspections from 48 to 24 months due to repeated instances of corrosion and cracking damage within the pylon region. That new inspection interval was formalised by Qantas within its maintenance documentation for its 747 fleet.^[2]

Figure 2: Boeing 747 engine pylon showing the location of the cracking (highlighted)

Source: Boeing, annotated by the ATSB

Manufacturer’s examination

Following removal from the aircraft, four cracked outboard strut ribs were sent to Boeing for metallurgical examination and fracture analysis. The examination confirmed that five cracks had developed in the strut ribs, all emanating from the rib cutaway region, ranging in length from 3—19 mm (Figure 3 to Figure 5). Boeing’s report^[3] contained the following detail:

• Optical examination of the fracture surfaces was conducted and no anomalies were found that might otherwise have contributed to the development of the cracking.
• Scanning electron microscopy (SEM) of the fracture surfaces did not find any aberrations or defects on the surfaces of the strut ribs that might have otherwise contributed to the development of the cracking.
• SEM analysis of the fracture surfaces confirmed that each strut rib had striations and crack-progression marks consistent with fatigue cracking. No evidence was found of ductile tearing or plastic deformation at the crack origins, which might otherwise suggest that the fatigue cracking had initiated due to excessive vibration or torque loads from a transient engine event.
• Metallographic examination of the aluminium alloy microstructure from each rib did not reveal any abnormalities that could have led to the fracture of the strut ribs.
• Chemical analysis of each strut rib found the composition met the materials specifications, as listed within the engineering drawings.
• Hardness and conductivity measurements from each strut rib confirmed that they were within the correct range, as listed within the engineering drawings.
• The ribs were painted and primed in accordance with manufacturing specifications.

Boeing also conducted a structural loads analysis on the pylon, which confirmed that despite the presence of the cracking, the No. 2 engine pylon retained sufficient residual strength for all certified loading conditions.

Figure 3: A strut rib in-situ within the No. 2 pylon with two cracks identified

Source: Boeing, annotated by the ATSB

Figure 4: The strut rib from Figure 3 after removal from the pylon

Source: Boeing, annotated by the ATSB

Figure 5: Close-up of the cracking in the cutaway region shown at Figure 4

Source: Boeing, annotated by the ATSB

Aircraft history

On 24 June 2016, a few months before the detection of the pylon cracks, VH-OJT sustained a failure of the low-pressure turbine from the No. 2 engine during the take-off roll. This resulted in a high-speed seizure of the engine resulting in significant vibration and torque loads on the pylon. Following removal of that engine, the pylon was inspected and no damage to the critical structural members was evident.

Maintenance records noted that when the cracks were discovered in the Hong Kong SAR, the aircraft had accumulated 79,928 hours and 8,313 landings. There were no other identified defects reported from previous inspections on the No. 2 strut support assembly for VH-OJT.

Safety analysis

Cracking of the engine pylon strut ribs

During routine maintenance on 5 October 2016, four strut ribs from the No. 2 engine pylon were found to have developed fatigue cracks. The parts met the manufacturer’s engineering drawing and material specifications. With no obvious defects detected, it is likely that the fatigue cracking developed from exposure to vibratory loading that occurred within the pylon during normal engine operation. A factor that may have contributed to the crack growth was the number of hours and flight cycles that the airframe had accumulated within its service life. The manufacturer’s analysis indicated that despite the crack damage to the ribs, the pylon retained sufficient residual strength for all load cases, which indicates that the aircraft remained safe to operate.

Approximately 6 months prior to the detection of the strut rib cracking, the No. 2 engine sustained a seizure that resulted in the transmission of significant vibratory and torque loads into the pylon structure. Metallurgical analysis of the fracture at the crack origins did not show any evidence of plastic deformation or tearing of the alloy that might otherwise suggest that the seizure was a contributing factor to the cracking.

Pylon inspection

When compared with the 48-month inspection interval recommended by the manufacturer, the 24‑month inspection interval set by Qantas provided greater opportunity for the timely detection of cracking and other defects within the pylon. The reduced inspection interval proved effective in this case, detecting cracks before they reached a level considered by the manufacturer to be a safety risk.

Findings

These findings should not be read as apportioning blame or liability to any particular organisation or individual.

• The fatigue cracking in four outboard strut ribs in the No. 2 engine pylon is likely to have developed from exposure to vibratory loading within the pylon during normal engine operation over the aircraft’s service life.
• The extent of fatigue cracking in the ribs did not affect the structural integrity of the pylon for all certified load cases, which meant that the aircraft’s safe operation was not affected.

Safety message

The Qantas Boeing 747 maintenance program for the engine pylon area is completed at a higher frequency and in more detail than is required by Boeing. This enhanced inspection regime was effective in detecting cracking of the strut rib assembly, before the cracks reached a level considered to be a safety risk.

This occurrence highlights the importance for vigilance during ongoing routine maintenance activities. It also highlights that, when relevant, operators should consider reviewing prescribed maintenance schedules in order to address issues that may develop as their aircraft age and accumulate time in service.

About this report Decisions regarding whether to conduct an investigation, and the scope of an investigation, are based on many factors, including the level of safety benefit likely to be obtained from an investigation. For this occurrence, a limited-scope, fact-gathering investigation was conducted in order to produce a short summary report, and allow for greater industry awareness of potential safety issues and possible safety actions. Publishing information  Released in accordance with section 25 of the Transport Safety Investigation Act 2003 Published by: Australian Transport Safety Bureau © Commonwealth of Australia 2018 Ownership of intellectual property rights in this publication Unless otherwise noted, copyright (and any other intellectual property rights, if any) in this report publication is owned by the Commonwealth of Australia. Creative Commons licence With the exception of the Coat of Arms, ATSB logo, and photos and graphics in which a third party holds copyright, this publication is licensed under a Creative Commons Attribution 3.0 Australia licence. Creative Commons Attribution 3.0 Australia Licence is a standard form licence agreement that allows you to copy, distribute, transmit and adapt this publication provided that you attribute the work. The ATSB’s preference is that you attribute this publication (and any material sourced from it) using the following wording: Source: Australian Transport Safety Bureau Copyright in material obtained from other agencies, private individuals or organisations, belongs to those agencies, individuals or organisations. Where you wish to use their material, you will need to contact them directly.

__________

1. The Boeing 747 has four engines numbered one through to four. The No. 1 engine is located at the outboard position on the left wing and the No. 4 engine is located at the outboard position on the right wing. Engines No. 2 and No. 3 are at the respective left and right inboard wing positions.
2. Qantas Engineering and Maintenance procedure, J-744-C2-54-010 (Zonal inspection of No.2 engine strut).
3. Boeing Research and Technology Lab Report SR 15601, dated 20 May 2017.

Safety summary

What happened

On the morning of 16 November 2016, a single-engine Cessna 208B aircraft, registered VH-LNH, operated by Aviair Pty Ltd, departed Solomon Airport, Western Australia on a charter flight to Karratha. On board the aircraft were two flight crew and 11 passengers. Approximately 8 km from the airport, while climbing through an altitude of approximately 4,600 ft, the aircraft sustained an engine failure. The flight crew heard a loud bang and observed smoke billowing from the exhaust.

The flight crew elected to conduct an emergency landing on a nearby dirt road associated with the Solomon mine precinct. The landing was accomplished without injury to the occupants and the aircraft sustained only minor damage.

What the ATSB found

The ATSB found that a compressor turbine blade from the Pratt & Whitney Canada (PWC) PT6A‑114A engine developed fatigue cracking and fractured after approximately 1.8 hours of operation, leading to an in-flight engine failure and forced landing of the aircraft. A repaired compressor turbine vane ring that was fitted to the engine was identified by PWC to contain variations in aerofoil geometry. These variations likely led to an increase in vibratory stresses and the associated development of fatigue cracking and fracture of the compressor turbine blade.

From this and other recent occurrences, the investigation identified that PT6A-114A engines fitted with compressor turbine vane rings that had been repaired in accordance with the United States Federal Aviation Administration (FAA) approved repair scheme STI 72-50-254 had an increased likelihood of CMSX-6 single-crystal compressor turbine blade fracture and subsequent failure of the engine.

The ATSB also found that the flight crew’s handling of the engine failure and subsequent emergency landing reduced the risk of damage to the aircraft and/or injury to the passengers or crew.

What's been done as a result

On 16 March 2017, the holder of the major repair specification for the compressor turbine vane ring, Southwest Turbine Inc. (STI), ceased conducting repairs on CT vane rings for fitment into PWC PT6A-114A engines.

On 29 March 2017, the engine manufacturer, PWC, released Service Instruction Letter (SIL) PT6A-252 to all operators of PT6A-114 and -114A engines. The SIL advised of the heightened risk of CMSX-6 single-crystal compressor turbine blade fatigue fracture when combined with a compressor turbine vane ring that had been repaired using processes that were not approved by PWC.

Proactive safety action from the aircraft operator, Aviair Pty Ltd, included adopting the recommendations contained in PWC SIL PT6A-252. Their Cessna 208B fleet was reviewed, and any PT6A-114A engines containing compressor turbine vane rings that had been repaired using ‘non-PWC approved processes’, were replaced with CT vane rings produced by PWC.

On 19 August 2019, Transport Canada released airworthiness directive (AD) CF-2019-30 linking the low-time fatigue-fracture of CMSX-6 single-crystal compressor turbine blades and the use of CT vane rings that had been repaired in accordance with repair specification STI 72‑50‑254. The AD required operators to check for, and remove, STI‑repaired CT vane rings from PT6A‑114 and PT6A-34 series engines within a defined period of 9 calendar months, or 250 hours of operation.

The Civil Aviation Safety Authority automatically adopted AD CF-2019-30. This required Australian operators to remove STI-repaired compressor turbine vanes rings fitted to PT6A-114A and PT6A‑34 series engines.

On 17 August 2020, the FAA released a Notice of Proposed Rulemaking advising that they were considering mandating AD CF-2019-030. Public submissions to that process closed on 1 October 2020.

The vibratory effect of repaired CT vane rings on CMSX-6 single-crystal CT blades within PWC PT6A-114A engines was identified as a safety issue by the ATSB. However, due to the significant safety action taken by the directly involved parties since the occurrence, including action from the engine and component manufacturers, as well as the aviation regulatory authorities, the ATSB considers that the risk of CMSX-6 single-crystal CT blade fractures in Australian‑operated PWC PT6A-114A engines has been adequately addressed.

Safety message

This occurrence shows how subtle changes can have a detrimental effect on modern complex turbine engines. In this instance, geometry variations in a repaired compressor turbine vane ring likely led to rapid fatigue cracking and fracture of a compressor turbine blade, and subsequent engine failure.

The incident also reinforces the importance of communication and effective decision-making during an emergency. The crew’s handling of the forced landing minimised the potential for injury or aircraft damage.

VH-LNH Cessna 208B aircraft

Source: Jim Woodrow

The occurrence

What happened

On 16 November 2016, at about 0630 Western Standard Time,^[1] a single-engine Cessna 208B aircraft, registered VH-LNH, departed Karratha Airport, Western Australia (WA) on a flight to Solomon Airport, WA. On board the aircraft were the pilot-in-command, a safety pilot, and 11 passengers. The flight was operated by Aviair as a charter^[2] service, transferring workers to and from the Solomon Hub, an iron-ore mining and processing facility in the Hamersley Ranges of northern WA.

At about 0710, after a 40-minute flight, the aircraft arrived at Solomon Airport. Following disembarkation at the terminal, the flight crew prepared the aircraft for the return flight with 11 passengers. The start-up and pre-flight checks for the return flight to Karratha were conducted without any reported issues. At 0740, the aircraft was taxied to runway 27^[3] and about 3 minutes later, the flight departed for Karratha, initially conducting a left climbing turn overhead the Solomon Airport and then climbing on a north-westerly heading over the Hamersley Range (Figure 1).

A few minutes after taking off, while climbing through an altitude of approximately 4,600 ft above mean sea level (AMSL), the pilot-in-command reported hearing a ‘loud bang’ and ‘grinding noise’ from the engine followed by a decrease in aircraft performance. Blue and white smoke ‘billowed’ from the exhaust. Instruments within the cockpit showed rapidly reducing turbine speed and torque, indicating the aircraft’s engine had failed. At this point, the aircraft was about 8 km from the departure airport.

The pilot-in-command recalled lowering the aircraft’s nose to attain the recommended glide speed and reducing the engine power lever to idle as part of the initial checks. On commencing a right turn back toward the airport, the flight crew assessed that an emergency landing to the airport would not be possible. The aircraft would be gliding into a southerly headwind and given their altitude, it was uncertain whether they could safely avoid the high terrain between their current position and Solomon Airport.

Scanning the immediate area, the flight crew identified an unsealed dirt road as a potential emergency landing area. The road was oriented north-south and was near to an accommodation village associated with the mine construction. The pilot-in-command delegated the emergency calls to the safety pilot and concentrated on setting up for the forced landing. The safety pilot conducted the necessary actions to secure the engine further and communicate their situation. The fuel lever was set to cut-off and the propeller lever to feather.^[4] Suspecting that the aircraft was too low to communicate effectively with air traffic services, the safety pilot broadcast an emergency MAYDAY^[5] advising of their situation on the VHF common traffic advisory frequency for Solomon Airport. A departing Qantas aircraft detected the broadcast, along with Solomon Airport personnel who were monitoring the frequency. The safety pilot also activated the Spidertracks^[6] emergency notification system.

In preparation for the forced landing, the safety pilot provided an emergency passenger briefing.

The pilot-in-command manoeuvred the aircraft for an approach to the mine road from the south. Full flap was selected about 1,000 ft above the terrain. The flight crew observed three vehicles travelling along the mine road close to their intended landing area. Approximately 100 ft above the road surface, the pilot flashed the aircraft’s landing lights to warn the driver of an approaching mine vehicle that the aircraft was preparing for an emergency landing. On touchdown, heavy braking was applied, slowing the aircraft until it came to a controlled stop about 150 m short of a bend in the road.

All passengers remained seated until instructed to exit the aircraft. The safety pilot moved the passengers away from the aircraft, while the pilot-in-command inspected for fire and other damage. Emergency responders from the airport arrived on scene shortly after. With the exception of some flat spots on the main gear tyres due to the emergency braking, the aircraft was undamaged. There were no injuries.

Figure 1: Sequence showing the aircraft departure from Solomon Airport, the pilot‑reported location of the engine failure, and other events during the forced landing

Spidertracks aircraft location markers (identified green in the above image) are transmitted every 2 minutes under normal operation. When the SOS function is activated transmission signals are broadcast at 10-15 second intervals.

Image source: Google, Maxar Technologies, annotated by the ATSB

__________

1. Western Standard Time (WST): Coordinated Universal Time (UTC) + 8 hours.
2. Charter operations involve the carriage of passengers and/or cargo on non-scheduled flights by the aircraft operator, for trade or commerce.
3. Runway number: the number represents the magnetic heading of the runway.
4. Propeller blades can automatically turn (or be intentionally turned) to a feathering angle, following engine failure or malfunction, to minimise drag and prevent further damage.
5. MAYDAY: an internationally recognised radio call announcing a distress condition where an aircraft or its occupants are being threatened by serious and/or imminent danger and the flight crew require immediate assistance.
6. Spidertracks is a satellite-based tracking system that includes an SOS function for use in an emergency.

Context

Recorded information

Engine data

The aircraft was equipped with an engine condition trend monitoring system. A range of parameters were automatically recorded by the system, including:

• pressure altitude
• airspeed
• engine torque
• turbine temperature
• gas generator speed
• fuel flow
• propeller shaft speed.

The system was designed to assess engine performance by allowing an analyst to monitor recorded data over longer periods to identify emerging trends. It also recorded any turbine temperature or torque exceedances that can contribute to longer-term degradation in engine performance. The system recorded flight data at approximately 0.5-second time intervals.

The data from the flight from Karratha to Solomon Airport, conducted earlier on the day of the occurrence showed no major deviations or ‘events’, generated from an inter-turbine temperature exceedance, gas generator overspeed or fuel flow anomaly. Data from the monitoring system also indicated normal operation of the engine up to the point of failure.

The recorded data from the occurrence flight is plotted in Figure 2. Following departure, the aircraft was climbing at an indicated airspeed of 106 kt when at 4,615 ft AMSL and time 0747:47, the engine failed. This was represented in the data by a rapid decline in the measured parameters for torque, propeller shaft speed, fuel flow and gas generator speed. The pilot’s report of sudden loud noises from the engine were consistent with the abrupt variation in the recorded parameters.

Eight seconds after the initial reduction in engine parameters, torque and fuel flow had reduced to zero, correlating with the pilot’s recollection of the events immediately following the failure. At 0750:17, the aircraft had descended to 4,200 ft AMSL and the recorded propeller speed had reduced to zero. After the failure, the inter-turbine temperature rapidly increased from 715 C, peaking at 887 C, before steadily declining as the failed engine cooled. Twenty-nine seconds after the engine failed, the gas generator speed had reduced from 99.5 per cent to zero. Changes to the indicated airspeed and altitude identified that touchdown on the mine road occurred at 0751:09 at an airspeed of approximately 80 kt.

Figure 2: Presentation of the recorded data from the occurrence flight noting that the engine failure occurred at 0747:57 and the touchdown occurred at 07:51:09

Inter-turbine temperature, gas generator speed, indicated airspeed, pressure altitude, propeller shaft speed, torque and fuel flow are plotted. The data is presented 4 minutes after take-off until the emergency landing.

Image source: ATSB

Spidertracks

The aircraft was equipped with a Spidertracks tracking and communication system. The system was satellite-based and sent a position update for the aircraft every 2 minutes. This allowed Aviair to monitor the progress of the aircraft in near real‑time. An ‘SOS’ button on the keypad of the system could be activated by the flight crew. In the event the ‘SOS’ function was activated, the fidelity of the flight track was improved by transmitting a data point at 10‑15 second intervals.

Data from the Spidertracks system showed that the aircraft was tracking to the northwest when the engine failure occurred. Two ‘SOS’ pings were transmitted by the system late in the sequence, shortly before touchdown on the mine road, however, these transmissions did not continue at the expected 10‑15 second intervals for the remainder of the flight and landing. Spidertracks advised the ATSB that the unit on board the aircraft sustained a power shutdown shortly after the SOS signals were emitted. No further information was available to explain the signal loss during the emergency broadcast.

Emergency locator beacon

The flight crew did not activate the emergency locator transmitter (ELT) during the emergency. The ELT was a radio beacon that, once activated, could assist search and rescue operations locate the aircraft and its occupants.

The Australian search and rescue authorities provide the following advice^[7] for pilots in relation to the use of an ELT during an emergency:

When in the air, a beacon can be activated if you feel you are in grave and imminent danger. Facing a forced landing would be a reason to feel in grave and imminent danger. In a forced landing situation, when you are unsure of your position, when you are lost or have deviated from the planned or notified flight route, you should activate your distress beacon.

During this emergency, the flight crew of VH-LNH were in two-way radio communication with Solomon Airport authorities along with other commercial aircraft. The Spidertracks SOS alert had also been activated. These measures provided the flight crew with a level of assurance that an appropriate response was being provided. Had the emergency occurred later in the flight, in a remote area without access to a suitable landing area, in-flight activation of the ELT would likely have assisted any search and rescue operation identify the aircraft position and its occupants.

Flight data recorders

The aircraft was not fitted with a flight data recorder or cockpit voice recorder. VH-LNH had a maximum take-off weight under 5,700 kg, utilised a non-pressurized hull, and had a single turbine engine. Fitment of flight recorders was not required for such an aircraft.^[8]

Meteorological information

Weather observations were recorded every minute by the Solomon Airport automatic weather station (AWS). Between 0720 and 0751, the surface winds were from the southwest, averaging 11 kt, and gusting to 14 kt. No cloud or significant weather was observed within the vicinity of the airport.

The pilot-in-command reported that weather conditions were consistent with the AWS data, and that the flight was conducted in visual meteorological conditions.

Personnel information

Both the pilot-in-command and the safety pilot were appropriately qualified to conduct the flight. The pilot-in-command had a total aeronautical experience of 541 hours, with 222 hours flying Cessna 208B aircraft. The safety pilot had a total aeronautical experience of 1,972 hours, with 312 hours flying Cessna 208B aircraft. The pilot-in-command reported that due to recent company changes, this was his first flight to Solomon Airport. Aviair normally conducted single-pilot charter operations, however, for this flight a safety pilot experienced with the Solomon route and its surrounds was on board to help familiarise the pilot-in-command with the route.

Aircraft information

The Cessna 208B (Caravan), registration VH-LNH, airframe serial number 208B0590, was a high‑wing, all‑metal, non-pressurised aircraft with a fixed tricycle landing gear. It was manufactured in the United States in 1996 and was configured to carry a maximum of 13 people (two pilots and 11 passengers). The aircraft was first placed on the Australian civil register in December 2006. A nose-mounted 675 shaft horsepower Pratt & Whitney Canada (PWC) PT6A‑114A free-turbine, turboprop engine^[9] was fitted with a McCauley three-blade constant‑speed, full-feathering propeller.

Engine failure procedure

The initial memory actions for the Cessna 208B engine failure procedure are to establish 95 knots indicated airspeed (KIAS), power lever to IDLE, propeller to FEATHER and fuel condition lever to CUTOFF (Figure 3). The pilot-in-command’s recollection of the initial actions in response to the engine failure were consistent with the flight manual procedures.

Figure 3: Cessna 208B initial items for the engine failure procedure

Image source: Cessna - Textron Aviation

Aircraft glide performance

The aircraft was at an altitude of approximately 4,600 ft AMSL and Solomon Airport was about 8 km (4.3 NM) to the south‑east when the engine failed. The airport is at an elevation of 2,000 ft AMSL, indicating the aircraft was approximately 2,600 ft above the airport elevation when the engine failed. Taking into account an approximate 600 ft height loss due to the initial right turn towards the airport, the flight crew probably had an allowance of about 2,000 ft above the airport elevation when assessing the available landing options.

VH-LNH was configured with a cargo pod, which resulted in a glide range of 4 NM at 2,000 ft above the airport elevation. The best glide speed at the maximum weight is 95 KIAS, and 80 KIAS for a power-off landing speed with full flap. The airspeeds recorded by the engine-recording module were consistent with these figures.

A graph of the Cessna 208B glide performance from the pilot operating manual is depicted at Figure 4. Calculations indicated that, based on the published data, it might have been possible for the aircraft to reach the airport in a glide on a direct track in nil wind conditions. However, given the south‑westerly wind would have resulted in a headwind component, it was unlikely the aircraft would have reached the airport. The crew also considered that an attempt to reach the airport presented an unnecessary risk of a forced landing on unprepared ground in rugged terrain, and instead, they opted to conduct the forced landing on the mine road.

Figure 4: Cessna 208B maximum glide range with the glide distance highlighted

Image source: Cessna - Textron Aviation, annotated by ATSB

Engine information

A typical PWC PT6A engine cross-section is shown in Figure 5. The design incorporates two independent turbines: the compressor turbine that is coupled to the compressor, and the power turbine that is coupled to the propeller reduction gearbox. Inlet air enters the engine via a plenum chamber and is compressed by the axial/centrifugal compressor. Fuel is injected into the compressed air stream and ignited within the combustion chamber.

The pressurised outflow of expanding hot gases from the combustion chamber is deflected onto the blades of the compressor turbine by the preceding compressor turbine vane ring. The power turbine is rotated by the remaining energy in the gas flow exiting the compressor turbine which, in turn, drives the propeller shaft via a two-stage planetary reduction gearbox. Inter-turbine temperature is monitored by an integral probe installed between the compressor turbine and power turbine.

Figure 5: Basic layout of a PT6A engine

Image source: PWC, modified and annotated by ATSB

Engine disassembly

Following the forced landing, the aircraft was relocated from the mine road to Solomon Airport. An initial on-wing inspection through the exhaust duct of the engine revealed considerable damage to the power turbine (Figure 6). Several power turbine blades were missing, having fractured at places varying from the blade root to the tip. Aviair personnel subsequently removed the engine and transported it to an approved facility for controlled disassembly and examination.

Investigators from the ATSB supervised the examination. Representatives from PWC, the Civil Aviation Safety Authority (CASA), Aviair and other interested parties attended the examination. The examination found the most significant damage had occurred to the power turbine and compressor turbine sections of the engine.

Figure 6: Composite image displaying the observed damage to the power turbine (left) and compressor turbine (right) displaying the missing aerofoil section.

Image source: ATSB

Compressor

Inspection of the compressor section identified that it was relatively undamaged, and the rotor assembly was free to rotate. Slight tip rub to the compressor blades had occurred. There was no evidence of foreign object damage.

Power turbine

The power turbine had sustained the most severe damage, with several blades having fractured due to overstress at varying heights between the blade root and the tip due to contact with upstream^[10] debris (Figure 6). The upstream face of the power turbine inlet guide vane had sustained significant hard-body impact damage, most likely from metallic debris released from the compressor turbine. The power turbine was seized and not able to rotate freely. The propeller shaft was also unable to be rotated.

Compressor turbine

Examination of the compressor turbine (CT) identified that a single CT blade had fractured close to the blade root (Figure 7 and 8). Multiple blades had sustained bending from impact damage with the liberated aerofoil section. One of the adjacent blades displayed a clear imprint from contact with the liberated aerofoil section (Figure 9). The fracture surface of the failed blade was largely flat and located at the root, close to the platform. Characteristic ‘beach mark’ features were present that extended to approximately 50 per cent of the overall fracture surface, indicating that a fatigue crack had progressed through that portion of the blade. The remainder of the fracture was jagged and angular and representative of a sudden overstress fracture.

Manufacturing identifiers on the fractured blade base confirmed it to be a PWC component:

• 3079351-01 (part number)
• HWB5PL (serial number)
• 702055 (vendor supplier)

Aside from the missing aerofoil section from one of the CT blades, the engine disassembly did not reveal any other anomalies with respect to the CT disc and its installation.

Figure 7: The fractured CT blade

Fractured blade manufacturing identifiers; part number 3079351-01, serial number HWB5PL, vendor supplier 702055.

Image source: ATSB

Figure 8: Close view of the the CT blade fracture surface

Image source: ATSB

Figure 9: Composite view of a CT blade adjacent to the one that had fractured showing the collision imprint from the liberated aerofoil section

An exemplar blade (right image) from the disc has been used to illustrate the collision of the liberated aerofoil section from the fractured blade.

Image source: ATSB

Compressor turbine vane ring

The CT vane ring fitted to the engine was also examined (Figure 10). The vanes had sustained abrasion and minor impact damage. The shroud segments had sustained two major gouges from the engine failure. The following manufacturing part numbers were identified:

• P/N 3029051 (part number)
• S/N 931E (serial number)
• WO 0452R16 CL 6.25
• STI 72-50-254 (authorised repair number)
• STI 02-0356

Figure 10: Compressor turbine vane ring assembly, view of the downstream face with gouge damage (labelled)

CT vane ring manufacturing identifiers: (part number) 3029051, (serial number) 931E, (work order) WO 0452R16, (flow class) CL 6.25, (major repair specification) STI 72-50-254, (repair authority) STI 02-0356.

Image source: ATSB

Oil system

The magnetic chip detectors for the reduction and accessory gearboxes, as well as the oil filter, were removed from the engine and inspected. A large amount of metallic debris had accumulated on the magnetic poles of each detector. Metallic debris was also present in the oil and oil filter.

Significant engine maintenance

The engine, serial number PCE-17404, had accumulated 9,210.1 hours and 12,375 cycles since new, and 1,001.3 hours and 1,057 cycles since the last overhaul. This was within PWC’s normal recommendation of 3,600 hours between overhauls.

Airworthiness Directive (AD) CF-2013-21R1 Compressor Turbine Blade Failures was complied with on 15 November 2016. The AD required the existing compressor turbine (CT) blades^[11] manufactured from IN100 material to be replaced with PWC blades manufactured from CMSX-6, a single-crystal nickel-based superalloy.^[12]

Records confirmed that the replacement CT blade set had a Transport Canada authorised release certificate from PWC. The part number of those blades was 3079351-01, indicating they were a single-crystal blade-type intended for use in this application and were the blades specified in AD CF-2013-21R1.

Due to the identification of converging cracks on the inner portion of the CT vane ring during maintenance inspection, the CT vane ring assembly was replaced. A release certificate authorised by the United States Federal Aviation Administration (FAA) accompanied the replacement CT vane ring. The certificate showed that the replacement vane ring, part number 3029051, serial number 931E, was repaired by Southwest Turbine Inc. in accordance with major repair specification STI 72-50-254.

The CT disc assembly and CT vane ring were provided by a Brisbane-based facility authorised to repair and overhaul PT6 engines. The parts were transferred to Karratha, where they were installed into the engine by an approved maintenance organisation. Following reassembly of the engine, records indicated that all required parameters of torque, inter-turbine temperature, fuel flow, gas generator and propeller speeds were met during power assurance testing.

During ground operation following the reassembly, a ‘whistling’ noise was identified when the gas generator speed was less than 60 per cent. A borescope inspection of the engine CT section along with a check of the bleed valve was performed, with no faults found. The engine operated for an additional 1.8 hours and two engine start-stop cycles before the CT blade fracture occurred. The recent maintenance activity is summarised in Table 1.

Table 1: Recent significant maintenance items for engine serial number PCE-17404

Component examination

Following this occurrence, the CT disc, blades and vane ring from the engine were sent to the Pratt & Whitney Canada (PWC) facilities in Montreal for detailed examination. An Accredited Representative^[13] from the Transportation Safety Board of Canada assisted the ATSB by providing oversight of the examination.

Compressor turbine blade from the occurrence engine

The PWC metallurgical investigation confirmed that the engine failure was initiated by the fracture of a single compressor turbine (CT) blade. Fracture and release of the CT blade aerofoil section into the gas path led to many of the surrounding CT blades sustaining tip-bending and abrasion from rubbing against the CT shroud segments. Major downstream damage to the power turbine resulted from the blade release. Other CT blades exhibited secondary damage in the form of random nicks, cracks, bending and indentations from repeated impact against the liberated debris.

The fractured blade exhibited a very flat and smooth fracture surface at the leading edge, with features characteristic of fatigue crack propagation. The fracture plane was located at the platform-to-blade radial transition. Crack arrest features, or ‘beach marks’, were also clearly defined on the fracture surface, confirming the progressive growth of the fatigue crack through the blade section (Figure 11).

Figure 11: High-magnification optical image of the fractured blade showing that cracking initiated at the leading edge on the suction face, progressing through approximately 50 per cent of the cross-section prior to overstress fracture and release of the aerofoil

Image source: ATSB

Scanning electron microscopy

PWC’s examination of the fracture surface using a scanning electron microscope (SEM) further confirmed that the fatigue crack initiated from the blade-coating’s inter-diffusion layer at the leading-edge surface (Figure 12). River lines on the fracture surface indicated that the fatigue crack had initiated from several sites along the leading edge of the blade. No evidence of anomalies such as coating defects or impact damage were noted at the fatigue crack initiation site. The crack initiation site was not associated with the blade casting parting line.

The SEM examination did not identify any defective features within the nickel-base alloy or coating material that might otherwise have led to the cracking.

Figure 12: High-magnification SEM image of the crack origin at the leading edge showing the direction of propagation ‘river lines’ and other features associated with the blade coating

Image source: Pratt & Whitney Canada, annotated by ATSB

PWC reported that a semi-quantitative analysis of the fractured blade base material was performed using energy-dispersive spectroscopy, which confirmed that the blade material met the chemical requirements for CMSX-6, the alloy-type specified for use as the base blade material. The material analysis and part number identification 3079351-01 were coincident with the blade being manufactured from a CMSX-6 alloy, with a platinum-aluminide outer coating, as specified.

Metallurgical examination

PWC’s destructive cross-sectioning of the CT blade through the fatigue initiation zone did not reveal any metallurgical anomalies in this region. The microstructure of the fractured blade exhibited uniform cuboidal gamma / gamma-prime^[14] with no evidence of heat alteration. The microstructural examination identified some patches of secondary reaction zone^[15] underneath the blade coating. They were not, however, associated with the fatigue crack initiation site and PWC did not consider them to be contributory to the fatigue crack development. A cross-section through the aerofoil of an adjacent blade from the CT disc did not identify evidence of heat alteration to the blade microstructure. There was no report of micro-cracking of the blade coating in the region of the fatigue origin.

PWC concluded that there were no material anomalies identified on the remnant CT blade that might otherwise have initiated the fatigue cracking. The ATSB subsequently reviewed the physical evidence in detail, examining the remnants of the fractured blade using an optical microscope and SEM (Figure 13). Although some patches of altered microstructure were identified underneath the blade coating, the patches were unlikely to have affected the crack initiation. After examining the CT blades and disc from the occurrence engine, and reviewing the PWC metallurgical report, the ATSB agreed with PWC’s findings in respect of the fractured CT blade.

Figure 13: Composite optical and SEM micrograph of the remnant fractured CT blade

The main SEM image is a cross-section through the crack origin. The primary microstructural features are described. Some alteration of the microstructure was observed with the formation of secondary reaction zone (SRZ) between the gamma / gamma -prime (ɣ / ɣ’) base microstructure and the blade coating.

Image source: ATSB

Compressor turbine vane ring from the occurrence engine

The CT vane ring fitted to the occurrence engine had the identifier STI 72-50-254. The identifying markings showed that the CT vane had been repaired by Southwest Turbine Inc. (STI), a supplier of components to the aviation gas turbine industry. The repair was performed under an approval granted by the FAA.

The CT vane ring from the occurrence engine was examined in the PWC Airfoil Laboratory and compared against the specifications for a PWC-manufactured vane ring^[16] for the PT6A-114A engine. STI did not participate in the inspection of the compressor turbine vane ring. PWC reported that notable differences existed between the STI-repaired vane ring and the specifications for the part. Several features on the STI-repaired vane ring were similar to a PWC part, but did not meet the specification for a PWC-manufactured CT vane ring, including:

• variations in the cooling air outlet slot dimensions and overall surface area
• significant differences in the vane sealing plate size and shape of the weldment
• variations in the leading and trailing edge aerofoil position with respect to the inner and outer shroud (Figure 14 to 16)
• deviation of the chord length of the CT vanes at two aerofoil planes
• a large variation in flow area between the individual vanes
• variations to the vane stagger angle, chord length and trailing edge thickness that did not meet the tolerances specified in the PWC design.

Appendix A contains comparison plots from PWC of the measured values for vane chord length, trailing edge thickness and vane stagger angles of the STI-repaired CT vane ring against the PWC design. Also contained within appendix A are plots of the vane stagger angle of other STI‑repaired compressor turbine vane rings from PT6A-114A engines that had sustained CT blade fracture.

Figure 14: Trailing edge positioning variations against the inner shroud from the STI‑repaired CT vane ring, highlighting a non-conformance

Image source: Pratt & Whitney Canada

Figure 15: Trailing edge positioning variations against the outer shroud from the STI‑repaired CT vane ring, highlighting a non-conformance

Image source: Pratt & Whitney Canada

Figure 16: Example of trailing edge positioning against the outer shroud from a PWC‑manufactured CT vane ring

Notable differences in vane positioning exist when compared with the STI-repaired item from the occurrence engine.

Image source: Pratt & Whitney Canada

PWC CT blade vibration measurement

PWC reported to the ATSB that during development of both the IN100 and CMSX-6 compressor turbine blades, the blade vibration characteristics were studied at various engine shaft speeds. A Mode-1 vibration, otherwise known as a fundamental frequency vibration, was identified as affecting the CMSX-6 blades. This increased the amplitude of aerofoil cyclic flexure, in turn, increasing stresses at the blade root.

PWC further reported that a resonant sub-harmonic of the Mode-1 vibration could excite the CT blades. Identified as the ‘Mode-1 7E’ (seventh excitation), PWC advised that it is driven by aerodynamic buffeting from the upstream combustor, which results in a pattern of non-uniform airflow around the CT disc. Different CT vane geometries and relative positioning can lead to buffeting of the CT blades from the upstream airflow and were reported to be a major contributor to the Mode‑1 7E excitation.

PWC advised the ATSB that, following the development and release from production of the single‑crystal CMSX-6 CT blades, a method was developed to evaluate the vibration experienced by the CT blades in PT6A-114A test engines. The method was a ‘non-contact strain measurement system’ (NSMS) and was intended to help manage high-cycle fatigue by identifying the vibratory conditions that might influence the fatigue life of the CT blades.

In basic terms, the NSMS testing utilised a laser-sensor arrangement to measure in-situ tip timing of each passing blade. The return signal from the time lag of each blade tip was analysed to assess the significance of aerofoil deflections, or the vibratory response of the blade, during engine operation.

PWC reported to the ATSB that their NSMS test program evaluated the effects of fuel nozzle variation, CT vane ring ‘clocking’ (or angular positioning), axial clearance, and new versus used engine hardware. In total, their NSMS testing involved 24 engine builds using different hardware and varying configurations. The testing indicated that the dynamic response (vibration) of the CT blades could be reduced by controlling variables such as:

• the inter-platform gap between respective blades
• CT blade trailing edge thickness
• the relative position of the vane (clocking)
• under-platform blade dampening.

PWC reported that the changes introduced for preferential CT vane clocking (angular positioning) and CT blade dampening (under-platform seals) was not intended to eliminate the Mode-1 7E excitation, but to reduce the magnitude of blade tip flexure at the operating speed. Incorporation of the seals was determined by PWC to achieve the greatest reduction in CT blade stresses by dampening blade tip flexure during operation.

The PWC NSMS test program also evaluated the vibratory response when an STI-repaired CT vane ring was used in place of their own product. The STI-repaired CT vane ring used in the testing program had been removed from a PT6A-114A engine that had sustained a CT blade fracture during post-overhaul testing.^[17] The CT vane ring had sustained impact and penetration damage to some of the vane aerofoil surfaces as a result of the blade failure.

PWC provided the ATSB with plots of the CT blade vibratory behaviour, comparing the two CT vane rings (STI-repaired and a PWC item). The PWC results indicated that blade deflections associated with the Mode-1 7E excitation could be increased by as much as 200 per cent when an STI-repaired CT vane ring was used in a PT6A-114A engine. PWC indicated to the ATSB that such deflections were due to airflow distortions leading to stresses that exceeded the design requirements for CMSX-6 CT blades.

It was noted by the ATSB that the deflection behaviour of the CT blades during the NSMS testing may not be totally representative because of the aerofoil damage previously sustained to the STI‑repaired item used during the NSMS test program.

When questioned by the ATSB about the damaged CT vane ring used during the NSMS test program, PWC indicated that the damage was unlikely to have contributed to an increase in CT blade flexure. The increases in CT blade flexure measured during testing was likely a function of the overall aerodynamic performance of the CT vane ring, rather than a product of local aerofoil damage.

In September 2019, PWC indicated to the ATSB that, although NSMS technology was a complementary tool, it was being implemented where practical and warranted during all future product testing and development activities. In addition, the NSMS test data analysis had been used to support the development of service bulletins released by PWC in reference to the use of single-crystal CMSX-6 compressor turbine blades within the PT6A-114A engine.

CT vane ring from occurrence engine

PWC’s geometric measurements taken of the STI-repaired CT vane ring removed from the occurrence engine revealed out-of-tolerance conditions across the aerofoils. Although the STI‑repaired CT vane from this occurrence was not subject to NSMS testing, PWC indicated that the aerofoil and geometry variations likely led to higher stresses being applied to the single-crystal CMSX-6 CT blades during operation of the PT6A-114A engine from VH-LNH. PWC further indicated that the observed geometry variations in the CT vane from the occurrence engine were similar to those on the repaired CT vane that was used during the NSMS testing. They concluded that those variations likely led to increased vibratory stresses, excitation of the CT blade and rapid failure due to high-cycle fatigue.

Compressor turbine vane ring repair

During the course of the investigation, the ATSB was provided with the major process specification, STI 72-50-254 for the repaired vane ring. The revision history of the document indicated that in December 2006, Southwest Turbine were authorised by the FAA to conduct the repair following the method set out in the specification. The STI specification contained the approved method for repair and alteration of a range of PWC part number CT vane rings. The PWC part number of the CT vane ring intended for the PWC PT6A-114A engine was listed within the document as an eligible candidate for the repair scheme. The specification described the methods and processes that:

…could restore the vane to at least equivalent to its original condition with respect to aerodynamic function, structural strength, resistance to vibration and deterioration, and other qualities that could affect airworthiness.

In broad terms, the repair workflow involved replacing a major portion of the existing PWC vane ring through machining and welding processes. Following the machining processes, only a small portion of the original structure remained (Figure 17). A brazing process was then used to weld the remnant PWC part to the Southwest Turbine replacement casting, which contained the vanes and supporting structure. The CT vane was then finish machined and coated for oxidation resistance. The flow area was adjusted, and the flow class recorded. The item then underwent non‑destructive penetrant inspection for cracks in the structure, dimensional inspection and part marking (inscribing manufacturing identifiers), followed by a final inspection before being issued an authorised release certificate.

Figure 17: Illustration of a PT6A CT vane ring with the blue highlighted section displaying the approximate portion that is replaced as part of the STI 72-50-254 repair

The remnant section of the original PWC part is identified as the grey ring located along the mounting flange. The ring of original material was kept while the remainder of the part discarded to be replaced as part of the STI 72-50-254 repair process.

Image source: Southwest Turbine Inc.

Regulatory review of major repair specification STI 72-50-254

Prompted by this occurrence and other PT6A-114A engine failures, in January 2017, representatives from Pratt & Whitney Canada (PWC) met with Transport Canada and the US Federal Aviation Administration (FAA) to review CMSX-6 single-crystal CT blade failures. The NSMS vibration data and analysis of the suspected contributory effects of non-PWC repaired CT vane rings on CT blade fatigue fractures was presented to the regulatory authorities at that meeting.

On 12 July 2017, representatives from the FAA visited Southwest Turbine Inc. (STI) to review the major repair specification, STI 72-50-254. During the visit, it was determined by the FAA that, from 2012 through to 2017, 73 CT vane rings intended for the PT6A-114A had been repaired in accordance with the STI 72-50-254 specification.

The FAA advised that the complete substantiation document, containing the technical analysis and engineering data utilised to develop the repair of the PT6A-114A compressor turbine vane ring, was not available at the time of their visit.

Subsequently, the FAA requested STI provide the data necessary to complete the repair. STI agreed to reproduce the substantiation document, in particular to address key design features that included:

• vane angle
• vane chord length
• trailing edge thickness
• vane throat area
• trailing edge position.

Data from Southwest Turbine Inc.

In August 2019 and September 2020, the ATSB received extensive communication from STI that provided the following details:

• STI advised that in March 2017 (prior to the FAA visit), they had voluntarily ceased conducting repairs in accordance with major repair specification STI 72-50-254 to the PWC CT vane rings (PN 3029051) for fitment to the PT6A-1114A. The halt in vane ring repair was reportedly due to the relatively small market, coupled with a perceived increased risk associated with installation of the mandated CMSX-6 single crystal blades.
• STI advised that following the FAA visit they commenced a project lasting in excess of 12 months to generate the technical data necessary to substantiate their repair.
• New and overhauled PWC CT vane rings were purchased by STI and optically scanned to produce digital models of the scanned part.
• The scanning process was then repeated for several STI-repaired CT vane rings to produce digital models of each vane ring type.
• In September 2017, a copy of the Southwest Turbine Inc. revised substantiation report, ‘Conformity Analysis Report’ CAR 72-50-254 containing the geometric data comparison for the repaired compressor turbine vane rings was provided to the FAA. STI again presented the data to FAA personnel in December 2019.
• STI indicated in their communication to the ATSB that the revised substantiation report demonstrated that:
  • ‘…..The data unequivocally demonstrated to the FAA… that dimensional and geometric features of the STI overhauled CT Vane Rings do not differ from the dimensional and geometric features of the PWC CT Vane Rings.’
• STI also reported that major repair specification, STI 72-50-254, is intended for the repair of multiple CT vane ring part numbers that are eligible for installation into engines in addition to the PT6A-114A engine type. Southwest Turbine Inc. reported that, in totality, they have repaired and returned into service over 1,155 CT vane rings in accordance with STI 72-50-254, with 73 of those eligible for installation into the PT6A-114A engine. Southwest Turbine has not performed the repair since March of 2017.

History of compressor turbine blade and related design changes

Inconel 100 compressor turbine blades

The PT6A-114A engine was originally developed to operate with compressor turbine blades manufactured from Inconel 100 (IN100) alloy. PWC reported to the ATSB their awareness of 80 events of IN100 CT blade failure between the years 1990 and 2010, including 27 resulting from creep and 26 from overstress. In those years, PWC published service bulletins and maintenance manual advice to provide information on over-temperature operating conditions that could lead to creep of the CT blades. High-cycle fatigue was not a large contributor to the fractures experienced by the IN100 CT blades, however, they were prone to creep fracture when the engine was operated outside the recommended power settings.

CMSX-6 single crystal compressor turbine blade development

In recognition of the temperature sensitivity of the IN100 CT blades, PWC developed a replacement blade (part number 307291-01) that was introduced into service in July 2008 via service bulletin (SB) 1669.^[18] The new blades were manufactured from CMSX-6, which is a low‑density superalloy using a vacuum investment casting process creating a single-crystal microstructure. The blade design also offered a platinum-modified aluminide coating system that was deposited to the external surfaces for barrier oxidation and thermal protection of the underlying CMSX-6 substrate. The main advantage of using a single-crystal alloy over a conventionally cast and directionally solidified alloy, such as the previous generation of CT blades manufactured from IN100, is the microstructural refinement that provides greatly enhanced high‑temperature creep-rupture properties.

Since the release of SB1669, PWC have progressively modified the CT blade design ‘to increase the vibratory stress margin’. The design changes and reasoning were published in several service bulletins as follows:

• In August 2013, SB 1727^[19] was released to advise of a new CT blade part number, 3072791‑02, that had reduced blade platform dimensions to avoid blade-to-blade contact during engine operation.
• In September 2014, SB 1749^[20] was released to advise of further dimensional changes to the CT blade design. The blade platform width had again been reduced to create greater clearances between blades, limiting the potential for blade-to-blade contact during service. Part number 3079351-01 CT blades were introduced via that service bulletin.
• In July 2015, SB1768^[21] was released to improve the vibratory stress margin of the CT blades. Design changes were made to allow for an optimised angular alignment and clocking of the PWC CT vane ring within the engine.
• In December 2015, SB 1769^[22] was released to incorporate the fitment of a seals (dampeners) underneath the CT blade platforms to dampen the effects of high-frequency blade tip vibrations. PWC reported that NSMS testing of the blades fitted with blade dampeners achieved the most significant reduction in blade flexure during operation. On 18 November 2016, PWC distributed Revision 1 to SB 1749, which recommended incorporation of under platform blade dampeners (SB 1769). PWC identified that this configuration provided the most significant reduction in blade flexure during engine operation.

Related compressor turbine blade failure occurrences

PWC reported to the ATSB that since the introduction of the single-crystal CMSX-6 blades,^[23] there have been 12 PT6A-114A engine failures due to CT blade fracture. Unlike the creep‑related failures involving the IN100 blades, each occurrence event of the single-crystal blades has been attributed to the initiation of high-cycle fatigue cracking. The cracking leads to overstress fracture of the blade aerofoil, and downstream damage to the power turbine and potentially to other sections of the engine. PT6A-114A engines that have sustained fracture of the single-crystal CT blades are listed in Table 2.

One of the CMSX-6 compressor turbine blade failure occurrences was during engine testing in which a blade fractured after just 0.25 hours of engine operation while the engine was fitted with a non-PWC CT vane ring. The remaining occurrences have predominantly resulted in forced landings with minor airframe damage and/or injuries to the occupants. There have also been two accidents leading to significant damage and serious and/or fatal outcomes for those on board. Both accidents were investigated by agencies within the state of occurrence as summarised below.

Tropicair – Papua New Guinea Accident Investigation Commission (PNG AIC) report 13‑1008

On 25 November 2013, a Cessna 208B Caravan, registration P2-SAH, was on a charter flight from Kamusi to Purarui River. There was one pilot and nine passengers on board. Shortly after being established in the cruise, the engine failed. The pilot elected to conduct an emergency landing at a remote airstrip adjacent Kibeni village. On touchdown, the aircraft bounced and overran the landing area, struck a palm tree, and impacted a river just beyond the village boundary. Three passengers sustained fatal injuries.

The PNG AIC investigation report contained a summary of technical findings from PWC concluding that the engine power loss was caused by the fracture of one CT blade in fatigue, which resulted in secondary damage to the remainder of the CT blades and downstream components. The fatigue originated from multiple origins on the pressure side of the blade’s trailing edge. The report indicated that the underlying factors for the fatigue initiation could not be determined with certainty.

PWC reported to the ATSB that the engine from the Tropicair aircraft had been fitted with a CT vane ring manufactured by PWC.

Aereo Servicios Empresarialis S.A – Mexican Directorate General of Civil Aviation (DGCA) Report ACCDTAFA011/2016MMDO

On 1 April 2016, a Cessna 208B Caravan, registration XA-ULU, was substantially damaged during an off-airport forced landing near Tayoltita, Mexico. The aircraft had departed from Tayoltita and was enroute to Durango. On board were nine passengers and one pilot. The engine failed while transiting over mountainous terrain, requiring the conduct of an emergency landing. The aircraft was landed along a rocky riverbed, however it struck a tree and large boulders leading to substantial damage and break-up of the fuselage. Two passengers and the pilot sustained fatal injuries.

PWC conducted the subsequent engine teardown and examination, the results of which were summarised in the Mexican DGCA report. The examination determined that a single compressor turbine blade had fatigue-fractured around the leading edge of the blade root. No metallurgical abnormalities were observed. In their analysis of the accident, the DGCA stated that the engine had been fitted with a CT vane ring, which had a repair not approved by PWC.

Table 2: PT6A-114A engine failures involving CT blade fatigue and fracture

Data source: PWC, current to May 2020

Of all the PT6A-114A engines that have sustained CMSX-6 compressor turbine blade fractures, PWC reported to the ATSB that none of those engines had under platform blade seals (dampeners) incorporated, in accordance with PWC SB 1769.

PT6A-114A engine in-flight shutdown rates

Certification

The Canadian Aviation Regulations^[24] (CARs) that describe the design, testing and construction of turbine engines are based on Part 33 of the United States Federal Aviation Regulations.^[25]

Engine manufacturers, such as PWC, must demonstrate they meet all aspects of the prescribed airworthiness standards (CARs). A Type Certificate covering all variants of the PT6A engine-type is issued by the national aviation administrator^[26] when the airworthiness standards are demonstrated, allowing for other countries to grant type acceptance.

Safety analysis

Part 533 - subsection 75 of the CARs prescribes that a safety analysis must be performed to demonstrate the likely consequences of all failures that can reasonably be expected to occur. The analysis of hazardous, major and minor engine effects is required. An engine failure resulting in an in-flight shutdown (IFSD) within the standard was defined as:

(1) An engine failure in which the only consequence is partial or complete loss of thrust or power (and associated engine services) from the engine will be regarded as a minor engine effect.

Although the CARs list an engine failure as a Minor Effect, PWC advised the ATSB that in consideration of the possible occurrences of loss of useable thrust coupled with an unavailability of an adequate landing site, their procedures classify an IFSD as a Significant Effect. Such classification has the potential to result in a substantial reduction in the safety margin, or the potential to have a more significant impact on flight. PWC further advised that for IFSD events associated with the design of a specific engine component, a probability of occurrence exceeding 1 failure per 1,000,000 flight hours would generally drive additional mitigation.

Failure rates

PWC reported to the ATSB that they use an industry standard method for calculating the mean time between IFSDs for their engines. A basic IFSD is a shutdown attributed to a malfunction directly related to the engine or an engine component. These occurrences are used in the failure rate calculations. They also include unsubstantiated events where investigations are still in progress to identify the primary part that has led to the failure. A non-basic IFSD is a failure caused by a component failure not directly related to the engine. These events are not included in the failure rate calculations and include occurrences such as fuel exhaustion, or bird ingestion.

The PT6A-114 engine was certified for use in the Cessna 208 Caravan when the aircraft entered service in June 1984. By May 2020, PWC reported that there had been 2,628 engines produced, including the -114(A) variant, which had accumulated 20,159,000 hours total flying hours.

Since the introduction of the single-crystal CMSX-6 blades, PN’s 3072791-01, 3072791-02 and 3079351-01, PWC reported that 1,958 engines had been fitted with the newer single‑crystal blade type. The accumulated -114A fleet time operating the single-crystal blades was approximately 6,742,000 hours.

As previously described in Table 2, six engines fitted with genuine, non-modified (or repaired) PWC CT vane rings had sustained single-crystal CT blade failure. For those engines, the calculated failure rate was approximately 0.89 failures per 1,000,000 flight hours. ATSB calculations indicated that the number of failures of -114A engines fitted with single-crystal blades and PWC CT vane rings represented 0.31 per cent of the overall ‘PWC-fleet’ population.

The ATSB notes that, at the time of writing, there have been no instances of CMSX-6 compressor turbine blade fractures on PT6A-114A engines that have platform blade dampeners incorporated. This is consistent with PWC’s assessment that blade dampeners have the greatest effect toward limiting in-service flexure of the CT blades at the engine operating speed and their use more generally across numerous turbine engine types to reduce CT blade tip flexure and vibration.

It is also noted that the failure rate without blade damper fitment is below the threshold that generally drives mitigation by PWC and that their incorporation, although recommended, has not been mandated by regulation. This indicates that the CT blade dampers provide a reliability improvement rather than resolution of a safety issue associated with PT6A‑114A engines.

Of the 73 CT vane rings that had been produced under the FAA-approved major repair specification STI 72-50-254, and fitted to PT6A-114A engines, there have been six occurrences of single-crystal CT blade failure. The time to failure of the single-crystal CT blade from when the STI 72-50-254 CT vane rings were installed into the respective PT6A-114A engines ranged from 0.25 to 617 flight hours. Assuming the 73 vane rings produced by Southwest Turbine were installed and operated, the calculated failure rate was approximately 24 failures per 1,000,000 flight hours. The number of failures of PT6A-114A engines fitted with single-crystal blades, and which had STI 72-50-254 repaired CT vane rings fitted, represented approximately 8 per cent of the overall PT6A-114A ‘STI-fleet’ population.

PT6A-114A turbine blade and vane ring regulatory action

Replacement of IN100 blades with single-crystal CMSX-6 blades

On 1 August 2013, Transport Canada announced the formal requirement to replace IN100 blades fitted to PT6A-114 and -114A engines with the newer, single-crystal CMSX-6 variant through the release of airworthiness directive (AD) CF-2013-21. The AD stated that:

There have been a number of reported incidents where Compressor Turbine (CT) blades failures have caused power loss on PT6-114 & PT6A-114A engines, resulting in in-flight shutdown (IFSD). Investigation by engine manufacturer Pratt & Whitney Canada (P&WC) has determined that when operated at high power and high temperature settings, the subject CT blades are prone to crack/fracture as a result of creep and/or sulfidation.

The AD provided a criteria and several options to maintain the ongoing airworthiness of the -114 / -114A fleet. These included:

• a reduced time between borescope inspection of the existing IN100 CT blades
• submitting a sample pair of IN100 CT blades for metallurgical evaluation
• replacing all existing IN100 blades with the newer single-crystal CT blades manufactured from CMSX-6 within a 36-month time period.

Removal of repaired STI 72-50-254 CT vane rings

On 19 August 2019, Transport Canada released AD CF-2019-30 linking the low-time fatigue‑fracture of CMSX-6 single crystal CT blades and the use of compressor turbine vane rings that had been repaired in accordance with major repair specification STI 72-50-254. The AD required:

1. Within 9 months or 250 hours airtime, whichever occurs first, from the effective date of this AD, determine if a CT vane, repaired in accordance with repair specification number STI 72-50-254, is installed on the affected engine and replace it with a serviceable non-STI repaired CT vane.

2. Within 9 months or 250 hours airtime, whichever occurs first, from the effective date of this AD, replace and discard any CMSX-6 CT blade that has been operating in service on an engine with an above-mentioned STI repaired CT vane installation.

3. As of the effective date of this AD, it is prohibited for anyone to allow the installation of an above-mentioned STI repaired CT vane on affected engines.

On 17 December 2019, Revision 1 to AD CF-2019-030 was released by Transport Canada. The revision contained some minor amendments to specifically restrict the installation of STI 72‑50‑254 repaired CT vane rings.

__________

7. Australian Maritime Safety Authority, 2019, www.amsa.gov.au/safety-navigation/search-and-rescue/aviation
8. Requirements relating to the fitment of flight recorders are detailed in Civil Aviation Order (CAO) 20.18, Section 6, Recording equipment.
9. A lower powered PT6A-114 variant, rated to 600 shaft horsepower, was installed in early-model Cessna 208 aircraft.
10. Because the internal mechanical architecture of the PT6 engine is of a reverse and folded configuration, and the installed orientation of the PT6 engine is reversed, directional and positional designators such as ‘forward’ and ‘rear’ are ambiguous. The more accurate designator of ‘upstream’ and ‘downstream’ with respect to the engine’s internal airflow is used when describing internal parts. Installed orientation of external engine components are referenced with respect to the airframe.
11. IN100 compressor turbine blades were PWC part number 3102401-01.
12. CMSX-6 compressor turbine blades were PWC part number 3079351-01.
13. International Civil Aviation Aircraft Organisation, Annex 13 Aircraft Accident and incident Investigations: A person designated by a State, on the basis of his or her qualifications, for the purpose of participating in an investigation conducted by another State.
14. Gamma / Gamma-prime (ɣ / ɣ’) is a microstructural constituent found in nickel-base heat-resistant alloys, which provides an enhanced creep resistance and strengthening effect when distributed as a fine precipitate.
15. A secondary reaction zone is a microstructural instability in nickel-base superalloys that may develop beneath the blade coating leading to reduced rupture strength and as a potential crack initiation site.
16. Pratt & Whitney Canada compressor turbine vane ring, part number 3029051
17. Turbo Power PT6A-114A engine failure, see Table 2 of this report.
18. Pratt & Whitney Canada, Service Bulletin 1669, ‘Turboprop Engine Compressor Turbine Blades – Inspection/Replacement of’, original issue 22 July 2008
19. Pratt & Whitney Canada, Service Bulletin 1727, ‘Turboprop Engine Compressor Turbine Blades – Replacement of’’, original issue 23 August 2013
20. Pratt & Whitney Canada, Service Bulletin 1749, ‘Turboprop Engine Compressor Turbine Blades – Replacement of’’, original issue 23 September 2014
21. Pratt & Whitney Canada, Service Bulletin 1768, ‘Turboprop Engine Compressor Turbine Vane Ring and No. 2 Bearing Housing Assembly Cover – Replacement / Modification of’, original issue 30 July 2015
22. Pratt & Whitney Canada, Service Bulletin 1769, ‘Turboprop Compressor Disk Balancing Assembly – Replacement / Modification of’, original issue 21 December 2015
23. Pratt & Whitney Canada, single crystal compressor turbine blade part number: 3072791-00, 3072791-01, 3072791-02 and 3079351-01
24. Canadian Aviation Regulations, Part V – Airworthiness Manual Chapter 533 – Aircraft Engines
25. United States Code of Federal Regulations, Title 14 Aeronautics and Space, Chapter I, Part 33 – Airworthiness Standards: Aircraft Engines
26. Transport Canada is the national aviation administrator from the state of manufacture.

Safety analysis

While on a charter flight within Western Australia from Solomon Airport to Karratha, a Cessna 208B aircraft, registered VH-LNH, sustained the failure of its single-turbine engine. The aircraft had departed just a few minutes prior and was climbing steadily when the engine failed. The pilots assessed that the safest option was to conduct a forced landing onto a nearby dirt road that was part of the Solomon mine precinct.

This analysis discusses the contributing factors that led to the engine failure, as well as the decision-making aspects of the flight crew during the conduct of the emergency landing.

Compressor turbine blade fracture

The failure of the Pratt & Whitney Canada (PWC) PT6A-114A engine was the result of fatigue cracking and fracture of a single compressor turbine (CT) blade. Release of the aerofoil section led to multiple consequent power turbine blade failures from impact debris.

The engine had undergone maintenance activity in the days prior to the occurrence, having had the CT vane ring and CT blades replaced. The blades were replaced to comply with airworthiness directive CF-2013-21R1. Following reinstallation of the engine, the aircraft was flown for approximately 1.8 hours prior to the failure. Information obtained from the engine teardown did not identify any maintenance-related factors that might have contributed to the engine failure.

The blade was manufactured from CMSX-6, a single-crystal nickel-base alloy. A coating to improve resistance to oxidation was applied to the blade surfaces comprising a thin layer of a platinum-modified aluminide material. Though the CMSX-6 alloy type and the aluminide coating system can, in combination, develop brittle sub‑surface microstructural phases and associated micro-cracks, no such features were identified that might have contributed to the fatigue cracking. There were no other material defects or damage identified that had the potential to initiate the fatigue cracking.

In the short time since installation, the fatigue crack had initiated at the leading edge of the CT blade, along the suction side of the aerofoil. It subsequently transitioned and propagated through the blade-chord toward the trailing edge, through approximately 50 per cent of the blade cross‑section, prior to overstress fracture. The high-cycle fatigue cracking seen in this occurrence was produced via bending and the associated high-frequency vibrations during engine operation.

Compressor turbine vibratory stresses

The compressor turbine vane ring from the occurrence engine had been repaired by Southwest Turbine Inc. The repair had been completed and the part released as airworthy in accordance with major repair specification STI 72-50-254. Initial approval to repair the PWC CT vane rings using STI 72-50-254 was granted to Southwest Turbine by the United States Federal Aviation Administration (FAA) in 2006.

At that time, the CT blades from PT6A-114 and -114A engines were manufactured from Inconel 100 (IN100) alloy. High cycle fatigue was not a large contributor to fractures experienced by the IN100 CT blades. Component life for the IN100 blades was generally limited by thermal effects, where exposure to higher operating temperatures led to oxidation, creep and overstress of the blade materials.

In 2008, PWC introduced single-crystal blades manufactured from CMSX-6 alloy. These blades were resistant to the temperature-related failure modes of the IN100 blades but were more susceptible to high-cycle fatigue cracking. In their early analysis of the single-crystal blade design, PWC identified a resonant modal frequency that could lead to increased flexure and cyclic vibration of the single-crystal blades. PWC’s subsequent engineering analysis of the vibration characteristics of the single-crystal blades, when fitted to the PT6A-114A engine, led to several design changes intended to increase the ‘vibratory stress margin’.

Changes to the single-crystal design included adjustment of the platform dimensions and the incorporation of under-platform blade dampeners. The under-platform dampeners, introduced by PWC through service bulletin 1769, were shown to have the greatest effect toward limiting the in‑service flexure of the CT blades and reducing the likelihood of a blade fatigue fracture from occurring. When the dampeners were incorporated with other hardware changes identified as a result of the PWC vibration testing, such as optimisation of the CT vane ring angular alignment, vibratory stresses within the CMSX-6 CT blades from blade flexure were minimised.

CT vane aerofoil geometry and their relative positions were also identified to have a significant effect on CT blade vibratory stresses. PWC identified through vibration testing that CT blade flexure could increase significantly when an engine was operated with a CT vane ring that contained aerofoil geometry variations outside of their design tolerance. The vibration testing was conducted using an STI‑repaired CT vane ring that had been removed from a PT6A-114A engine that had also sustained a CT blade failure. Although the CT vane ring had sustained damage from the blade release, PWC indicated that such local damage would not be detrimental toward the overall aerodynamic performance of the part.

The increase in CT blade dynamics was attributed by PWC to subtle, but important, overall dimensional variations in areas that that had been designed by PWC to control the incoming airflow onto the compressor turbine blade aerofoils. This included vane stagger angle, which can affect gas pressures from the inlet to the outlet of the vane, accelerating the mass flow, and leading to low‑amplitude, high-frequency excitations of the aerofoil surfaces.

The STI-repaired CT vane ring from the occurrence engine was not subjected to this vibration testing. However, it contained similar geometry variations to the tested component. Both PT6A‑114A engines had sustained a CT blade failure in a short period of operation. The identified out of tolerance geometry variations were also similar to other PT6A‑114A engines that had sustained CT blade fatigue-fracture in relatively rapid time when operated with STI-repaired CT vane rings. In the absence of any identified material defect associated with the failed occurrence CT blade itself, it can be concluded that variations in aerofoil geometry likely led to an increase in vibratory stresses, excitation and the associated development of fatigue cracking and fracture of the CT blade from the occurrence engine.

Failure rates

Major repair specification STI 72-50-254 by Southwest Turbine Inc. (STI) was a method approved by the FAA to restore the functionality of PWC CT vanes for fitment into numerous PWC PT6 engine models, one of which included the PT6A-114A. In part, the repair involved machining away a large portion of the vane structure and welding an equivalent casting in place. The replacement casting ensured structural fitment to the engine and also contained the vanes that determined the aerodynamic performance of the part.

Both STI and the FAA reported that, since the introduction of the repair, there were 73 STI‑repaired CT vanes released into service for installation into the PT6A-114A engine. Of that population, six PT6A-114A engines (of the 12 total reported failures) sustained a compressor turbine blade failure that led to an in-flight shutdown of the engine. This represented approximately 8 per cent of the repaired population. All CT blade failures occurred in relatively short time periods after fitment of the CT vanes, leading to an overall failure rate of approximately 24 failures per million flight hours. In contrast, the remaining six PT6A-114A engines that had sustained CT blade fatigue-failure were fitted with PWC‑manufactured CT vanes and none of them had been repaired/overhauled by PWC. This represented 0.31 per cent of the PT6A-114A fleet and a significantly lower failure rate of approximately 0.89 failures per million flight hours.

Comparing the failure rate from both populations showed that PT6A-114A engines fitted with the FAA-approved and repaired CT vane rings from STI had a significantly increased likelihood of CMSX‑6 compressor turbine blade fracture, compared to those engines fitted with PWC‑manufactured compressor turbine vane rings.

Flight crew response

The ATSB found that the flight crew response to the engine failure and subsequent emergency landing was handled in a competent and proficient manner that reduced the risk of damage to the aircraft and/or injury to passengers or crew.

Due to the aircraft’s position when the engine failed, and the southerly headwind component, it is unlikely the aircraft could have reached the airport. In consideration of the rugged terrain surrounding the Solomon Airport, the decision to land on the access road considerably reduced the risk of damage and injury. The delegation of the passenger briefing and emergency radio calls from the pilot-in-command to the safety pilot was an effective practical application of crew resource management.

Findings

From the evidence available, the following findings are made with respect to the engine failure and subsequent forced landing involving a Cessna Aircraft Company 208B, registered VH-LNH, that occurred 8 km northwest of Solomon Airport, Western Australia on 16 November 2016. These findings should not be read as apportioning blame or liability to any particular organisation or individual.

Safety issues, or system problems, are highlighted in bold to emphasise their importance. A safety issue is an event or condition that increases safety risk and (a) can reasonably be regarded as having the potential to adversely affect the safety of future operations, and (b) is a characteristic of an organisation or a system, rather than a characteristic of a specific individual, or characteristic of an operating environment at a specific point in time.

Contributing factors

• A compressor turbine blade from the engine developed fatigue cracking and fractured after approximately 1.8 hours of operation, leading to an in-flight engine failure and resultant forced landing of the aircraft.
• The compressor turbine vane ring that was repaired in accordance with STI 72-50-254 and fitted to the engine contained variations in aerofoil geometry that likely led to an increase in vibratory stresses and the associated development of fatigue cracking and fracture of the compressor turbine blade.
• Pratt & Whitney Canada (PWC) PT6A-114A engines fitted with compressor turbine vane rings that have been repaired in accordance with the United States Federal Aviation Administration-approved scheme STI 72-50-254 have a significantly increased likelihood of CMSX-6 compressor turbine blade fracture and subsequent failure of the engine compared to those engines fitted with PWC‑manufactured compressor turbine vane rings. [Safety Issue]

Other findings

• The ATSB found that the flight crew’s response to the engine failure and subsequent emergency landing on the mine road was handled in a competent and proficient manner that reduced the risk of damage to the aircraft and/or injury to passengers or crew.

Safety issues and actions

Compressor turbine blade failure

Safety issue number: AO-2016-155-SI-01

Safety issue description:  Pratt & Whitney Canada (PWC) PT6A-114A engines fitted with compressor turbine vane rings that have been repaired in accordance with the United States Federal Aviation Administration‑approved scheme STI 72-50-254 have a significantly increased likelihood of CMSX-6 compressor turbine blade fracture and subsequent failure of the engine compared to those engines fitted with PWC‑manufactured compressor turbine vane rings.

Sources and submissions

Sources of information

The sources of information during the investigation included:

• Aviair Pty Ltd
• The pilots from the occurrence aircraft
• Pratt & Whitney Canada
• Cessna - Textron Aviation
• United States Federal Aviation Administration
• Transport Canada
• Southwest Turbine Inc.

Submissions

Under Part 4, Division 2 (Investigation Reports), Section 26 of the Transport Safety Investigation Act 2003 (the Act), the ATSB may provide a draft report, on a confidential basis, to any person whom the ATSB considers appropriate. Section 26 (1) (a) of the Act allows a person receiving a draft report to make submissions to the ATSB about the draft report.

A draft of this report was provided to the operating flight crew of VH-LNH, Aviair Pty Ltd, Transport Canada, Transport Safety Board of Canada, Pratt & Whitney Canada, United States Federal Aviation Administration, Southwest Turbine Inc., United States National Transportation Safety Board, and the Civil Aviation Safety Authority.

Submissions were received from:

• Aviair Pty Ltd
• the Civil Aviation Safety Authority
• Pratt and Whitney Canada
• Southwest Turbine Inc.
• the United States National Transportation Safety Board
• the United States Federal Aviation Administration.

The submissions were reviewed and were considered appropriate, the text of the report was amended accordingly.

Appendix A: Compressor turbine vane ring measurements

Figure A1: The measured vane chord length and trailing edge thickness measurements of the STI-repaired CT vane ring from the occurrence engine (Aviair Cessna 208, VH-LNH) compared to the Pratt & Whitney Canada design tolerance.

Image source: Pratt & Whitney Canada

Figure A2: The respective vane stagger angles of the STI-repaired compressor turbine vane ring from the occurrence engine (Aviair Cessna 208, VH-LNH) compared to the Pratt & Whitney Canada design tolerance.

Image source: Pratt & Whitney Canada

Figure A3: The vane stagger angles of three STI-repaired compressor turbine vane rings fitted to three PT6A-114A engines that had sustained compressor turbine blade fracture, compared a CT vane ring with a Pratt & Whitney Canada CT vane ring that had also sustained a CT blade failure.

Image source: Pratt & Whitney Canada

Purpose of safety investigations The objective of a safety investigation is to enhance transport safety. This is done through: identifying safety issues and facilitating safety action to address those issues providing information about occurrences and their associated safety factors to facilitate learning within the transport industry. It is not a function of the ATSB to apportion blame or provide a means for determining liability. At the same time, an investigation report must include factual material of sufficient weight to support the analysis and findings. At all times the ATSB endeavours to balance the use of material that could imply adverse comment with the need to properly explain what happened, and why, in a fair and unbiased manner. The ATSB does not investigate for the purpose of taking administrative, regulatory or criminal action. Terminology An explanation of terminology used in ATSB investigation reports is available here. This includes terms such as occurrence, contributing factor, other factor that increased risk, and safety issue. Publishing information Released in accordance with section 25 of the Transport Safety Investigation Act 2003 Published by: Australian Transport Safety Bureau © Commonwealth of Australia 2020 Ownership of intellectual property rights in this publication Unless otherwise noted, copyright (and any other intellectual property rights, if any) in this report publication is owned by the Commonwealth of Australia. Creative Commons licence With the exception of the Coat of Arms, ATSB logo, and photos and graphics in which a third party holds copyright, this publication is licensed under a Creative Commons Attribution 3.0 Australia licence. Creative Commons Attribution 3.0 Australia Licence is a standard form licence agreement that allows you to copy, distribute, transmit and adapt this publication provided that you attribute the work. The ATSB’s preference is that you attribute this publication (and any material sourced from it) using the following wording: Source: Australian Transport Safety Bureau Copyright in material obtained from other agencies, private individuals or organisations, belongs to those agencies, individuals or organisations. Where you wish to use their material, you will need to contact them directly.

What happened

At about 0936 Central Daylight-saving Time (CDT) on 10 November 2016, a QantasLink Bombardier DHC‑8‑315 aircraft, registered VH-SBB (SBB), was on approach to Adelaide Airport South Australia. Two flight crew, two cabin crew and 43 passengers were on board the regular public transport flight.

At this time, the flight crew contacted air traffic control and informed the approach controller that they were at 9,000 ft. Soon after, the first officer, who was the pilot flying,^[1] noticed that the captain’s electronic attitude director indicator (EADI) screen had gone blank. The flight crew conducted the display failure checklist. As the captain turned the EADI screen off, they noticed the screen return to normal for about five seconds and then the entire screen pulsed on and off, before returning to a blank state.

The flight crew were then cleared to descend by air traffic control and at about 6,000 ft they noticed a faint electrical smell. They were not able to identify the source of the smell, but suspected that it originated from the failed EADI screen.

The cabin crew had prepared the cabin for landing and all passengers and the two cabin crew were seated. The flight crew contacted the cabin crew through the aircraft’s intercom and informed them that there was a smell in the cockpit. The cabin crew reported that they had not identified any unusual smells where they were seated at the rear of the aircraft.

The flight crew were cleared by air traffic control to conduct a visual approach for a landing on runway 12.

The flight crew determined that the smell was getting worse and conducted their phase one-memory checklist items for a fuselage fire or smoke. Both flight crew fitted their oxygen masks and smoke goggles and turned off the air-conditioning recirculation fans for the flight deck and the cabin.

One of the cabin crew went to the forward area of the cabin and could smell an electrical smell that did not go past row one and reported this to the flight crew. At this stage, the flight crew informed them that they were using their supplemental oxygen.

At about 0943, the flight crew made a PAN PAN^[2] call to air traffic control and advised that they had an instrument failure and electrical smell. As the aircraft was already, established on approach for runway 12 they were cleared to land by air traffic control. As the electrical smell was increasing in intensity, the flight crew advised the tower controller that they would require emergency services on standby.

The aircraft landed, taxied onto taxiway E and stopped just past the holding point clear of runway 12 (Figure 1). The first officer made an alert announcement through the aircraft’s public announcement (PA) system that informed the passengers to remain seated and await further instructions. The PA also signalled to the cabin crew that there was an abnormal situation that may require an emergency or precautionary evacuation. The flight crew contacted one of the cabin crew using the intercom and the cabin crew informed them that the smell in the cabin was getting stronger and that there were no other issues in the cabin. The flight crew contacted the ground controller at about 0950 to inform them that they would be conducting an evacuation at their position on the taxiway. They then conducted the precautionary disembarkation checklist. When the engine propellers had stopped, the captain made the precautionary disembarkation PA. The cabin crew at the front of the aircraft opened the main entry door and directed the passengers towards the airport safety officers that were located on the grassed area near the taxiway. The first officer disembarked the aircraft after the twelfth passenger and also directed the passengers to the grass area where the airport safety officers were located.

Figure 1: SBB parked on taxiway E after the crew and passengers had disembarked

Source: Airport operator

When all passengers had disembarked, the captain and cabin crew disembarked the aircraft with their emergency equipment. The captain briefed the airport fire fighters about the nature of the fumes. The fire fighters informed the captain that no toxic fumes or hot spots were detected although they were able to smell the strong electrical smell.

There were no injuries as a result of the occurrence and the aircraft was not damaged.

Captain’s comment

The captain reported practicing emergency procedures about four months prior to the occurrence when conducting simulator training. They commented how valuable that training was to be prepared for this type of occurrence. The captain indicated that the operator provided different tools for different situations. In this emergency, the captain reported using the GRADE model (gather information, review information, analyse alternatives, decide and evaluate the outcome of the action) and that it was helpful to evaluate the situation and decisions to ensure that there was a safe outcome.

The captain indicated that there were no issues with the communication between the flight crew and the cabin crew and that everyone worked together well.

Cabin crew comment

A member of the cabin crew reported that before the disembarkation, the passengers remained seated and calm. Most passengers followed the instructions to leave everything with only 2 or 3 people taking small bags from the aircraft. They advised that everything went smoothly, in accordance with their training and that they worked well as a team.

Aircraft operator comment

The aircraft operator reported that the electronic attitude director indicator (EADI) had failed. The fumes were caused by damage to a circuit card assembly due to a blown resistor on the video driver. There was no damage to any other aircraft parts or components.

Findings

These findings should not be read as apportioning blame or liability to any particular organisation or individual.

• The captain’s electronic attitude director indicator (EADI) failed, resulting in fumes in the aircraft.
• As a result of the fumes in the aircraft, the crew reduced the potential risk to the aircraft occupants by conducting a precautionary disembarkation onto a taxiway.

Safety message

Many factors come into play when pilots make decisions in the aviation environment. There are many different models and tools that pilots can use for effective decision making, such as GRADE, as was used by the captain in this event. The models involve a systematic approach to decision making, to consistently determine the best course of action in response to a given situation.

An understanding of the decision-making process provides a pilot with a foundation for developing aeronautical decision-making skills. Some situations require a pilot to respond immediately using established procedures, with little time for detailed analysis. These decisions are based upon training, experience, and recognition. Other situations require a more reflective response, where greater analysis is necessary.

Additional information is provided in the following publication:

US Federal Aviation Administration (FAA), Pilot’s handbook of aeronautical knowledge, Chapter 2: Aeronautical decision-making, available from the FAA website.

Aviation Short Investigations Bulletin - Issue 58

About this report Decisions regarding whether to conduct an investigation, and the scope of an investigation, are based on many factors, including the level of safety benefit likely to be obtained from an investigation. For this occurrence, a limited-scope, fact-gathering investigation was conducted in order to produce a short summary report, and allow for greater industry awareness of potential safety issues and possible safety actions. Publishing information  Released in accordance with section 25 of the Transport Safety Investigation Act 2003 Published by: Australian Transport Safety Bureau © Commonwealth of Australia 2017 Ownership of intellectual property rights in this publication Unless otherwise noted, copyright (and any other intellectual property rights, if any) in this report publication is owned by the Commonwealth of Australia. Creative Commons licence With the exception of the Coat of Arms, ATSB logo, and photos and graphics in which a third party holds copyright, this publication is licensed under a Creative Commons Attribution 3.0 Australia licence. Creative Commons Attribution 3.0 Australia Licence is a standard form licence agreement that allows you to copy, distribute, transmit and adapt this publication provided that you attribute the work. The ATSB’s preference is that you attribute this publication (and any material sourced from it) using the following wording: Source: Australian Transport Safety Bureau Copyright in material obtained from other agencies, private individuals or organisations, belongs to those agencies, individuals or organisations. Where you wish to use their material, you will need to contact them directly.

_________

1. Pilot Flying (PF) and Pilot Monitoring (PM): procedurally assigned roles with specifically assigned duties at specific stages of a flight. The PF does most of the flying, except in defined circumstances; such as planning for descent, approach and landing. The PM carries out support duties and monitors the PF’s actions and the aircraft’s flight path.
2. PAN PAN: an internationally recognised radio call announcing an urgency condition which concerns the safety of an aircraft or its occupants but where the flight crew does not require immediate assistance.

What happened

On 12 November 2016, a Robinson R44 helicopter, registered VH-YMJ, departed from a landing site on the Riversleigh Station property for the return leg of a private sight-seeing flight to Adels Grove aircraft landing area (ALA), Queensland. On board the helicopter were the pilot and three passengers.

The pilot used the helicopter to ferry two groups of people from Adels Grove ALA, to a dirt road landing site near a river in the morning for swimming, and then ferried the first group back to Adels Grove after lunch. At about 1420 Eastern Standard Time (EST), the pilot attempted the last planned departure of the day from the landing site for the return flight. The take-off direction followed the road, which was a south-south-east direction. The pilot reported that there were no power performance issues in the hover, but then during the initial climb, at about 100–130 ft above ground level (AGL) and between 15–20 kt airspeed, the helicopter started to experience a loss of performance. The helicopter started to descend and the pilot advised there was insufficient engine power to prevent the descent.

As the helicopter approached the ground, the pilot raised the collective lever^[1] to cushion the landing. The right skid of the helicopter landed first, on the side of the road and the left skid landed off the side of the road below the right skid, which resulted in the helicopter rolling onto the left side before coming to rest (Figure 1). One passenger received minor injuries and the helicopter was substantially damaged.

Figure 1: VH-YMJ accident site

Source: Operator

Terrain and weather conditions

Riversleigh Station is located in the north-west quarter of Queensland, about 200 km south-west of the southern corner of the Gulf of Carpentaria. The countryside along the river in the vicinity of the accident site was described by the pilot as hilly with plateaus and escarpments and tall gumtrees.

The pilot reported that the weather conditions started fine in the morning, but changed late in the morning with a hot wind, which was gusting in strength and varying in direction. Wind strength varied from 0–20 kt, and wind direction varied between south and south-east. The temperature was about 38 °C, the elevation of the landing site was about 430 ft above mean sea level (AMSL), and the QNH^[2] was about 1010 hPa. There were also rain showers in the vicinity.

Aircraft performance

The maximum take-off weight (MTOW) published for the Robinson R44 in the rotorcraft flight manual is 1,088 kg. The weights of the occupants and estimated fuel on board at the time of the accident indicate the all-up weight (AUW) of the helicopter was about 1,041 kg. With an estimated elevation of 430 ft, QNH of 1010 hPa and maximum height on take-off of 130 ft, the pressure altitude was about 650 ft when the pilot noticed there was insufficient power to continue the climb. The pilot reported that they started the take-off with about 10 kt of head wind and the helicopter had just passed through translational lift^[3] and was at about 15–20 kt before the descent started.

If the wind speed dropped during the initial climb, then the helicopter could have been below translational lift at 100-130 AGL when the descent started. In this case the helicopter would have been in the hover out of ground effect^[4] (HOGE) flight regime (Figure 2). At the reported temperature and AUW, this would place the helicopter at the limit of the hover altitude for the power available (point B, Figure 2).

In comparison, the hover in ground effect (HIGE) performance chart indicated the helicopter could maintain a hover at about 4,500 ft pressure altitude at the AUW and 38 °C.

Air density

The power produced by the engine and the lift produced by the helicopter rotors are influenced by air density. Low atmospheric pressure^[5] and hot and humid conditions decrease air density. A decrease in air density decreases power available from the engine, but increases the power required for rotor thrust, because a larger angle of attack^[6] is required from the rotor blades to produce the same lift.

Tail rotor and demand for power

The helicopter tail rotor is an anti-torque device, which is controlled by the tail rotor pedals to increase or decrease the angle of attack of the tail rotor blades. The engine provides the power for the tail rotor drive. Therefore, an increase in demand for anti-torque ‘bleeds off’ engine power. When the relative wind is directly in front of the nose of the helicopter, the helicopter airframe behaves like a weathervane, holding the nose of the helicopter into the wind and reducing the requirement for anti-torque. However, if the wind strikes the helicopter from the right side, this will increase the demand for left tail rotor pedal to maintain heading, which will bleed off engine power.

Figure 2: R44 HOGE performance

Source: Manufacturer, annotated by ATSB

Previous incidents

Previous ATSB reports of Robinson R44 helicopters descending with insufficient power in low airspeed and low air density (high density altitude) conditions include the following:

• Collision with terrain involving a Robinson R44, VH-HLB (AO-2014-154)
• Collision with terrain involving a Robinson R44, VH-UGC (AO-2013-203)
• Collision with terrain 10 km west of Gunpowder Mine, Qld, 21 February 2006, VH-HBS (200600979)

Safety analysis

The AUW of the aircraft was below the published MTOW and within the published limits for HIGE operations. At the time of take-off, there were no unusual noises or vibrations, the engine was delivering power to the rotors and the rotor speed did not decrease below limits. The pilot estimated they had a 10 kt headwind component, but that the wind was gusting in strength and variable in direction. Therefore, it is likely that the forced landing was the result of insufficient power for the prevailing environmental conditions at the helicopter’s AUW.

Two possible scenarios for insufficient power are a shift in wind direction to the right side of the helicopter or a decrease in head wind strength. In the first scenario, a shift in wind direction to the right would demand more left tail rotor pedal to maintain take-off heading and decrease the power available from the engine for main rotor thrust. In the second scenario, a decrease in wind strength just after translational lift would place the helicopter inside the HOGE flight regime and at the limit for the take-off AUW and temperature.

Findings

These findings should not be read as apportioning blame or liability to any particular organisation or individual.

• The forced landing was the result of power required to continue flight in excess of power available for the take-off AUW, temperature and wind conditions.
• The helicopter was within the published maximum take-off weight limit.

Safety message

This incident highlights the effect of high AUW, high air temperature and gusting wind conditions on the R44 helicopter’s performance. In particular, the combination of high AUW and high air temperature increase the power required and decrease the power available, which can lead to a significant difference between the HIGE and HOGE performance. In addition, the pilot reported that it is important to keep a close eye on changing wind conditions as they had never previously experienced a similar loss of performance.

Aviation Short Investigations Bulletin - Issue 57

About this report Decisions regarding whether to conduct an investigation, and the scope of an investigation, are based on many factors, including the level of safety benefit likely to be obtained from an investigation. For this occurrence, a limited-scope, fact-gathering investigation was conducted in order to produce a short summary report, and allow for greater industry awareness of potential safety issues and possible safety actions. Publishing information Released in accordance with section 25 of the Transport Safety Investigation Act 2003 Published by: Australian Transport Safety Bureau © Commonwealth of Australia 2017 Ownership of intellectual property rights in this publication Unless otherwise noted, copyright (and any other intellectual property rights, if any) in this report publication is owned by the Commonwealth of Australia. Creative Commons licence With the exception of the Coat of Arms, ATSB logo, and photos and graphics in which a third party holds copyright, this publication is licensed under a Creative Commons Attribution 3.0 Australia licence. Creative Commons Attribution 3.0 Australia Licence is a standard form licence agreement that allows you to copy, distribute, transmit and adapt this publication provided that you attribute the work. The ATSB’s preference is that you attribute this publication (and any material sourced from it) using the following wording: Source: Australian Transport Safety Bureau Copyright in material obtained from other agencies, private individuals or organisations, belongs to those agencies, individuals or organisations. Where you wish to use their material, you will need to contact them directly.

1. Collective: a primary helicopter flight control that simultaneously affects the pitch of all blades of a lifting rotor. Collective input is the main control for vertical velocity.
2. QNH: the altimeter barometric pressure subscale setting used to indicate the height above mean seal level.
3. Translational lift is additional lift gained from induced airflow through the main rotors gained from forward airspeed. It reduces the power required for rotor thrust and becomes noticeable as the helicopter experiences about 12–14 kt of airspeed.
4. Out of ground effect: helicopters require less power to hover when in ‘ground effect’ then when out of ‘ground effect’ due to the cushioning effect created by the main rotor downwash striking the ground. The height of ‘ground effect’ is usually defined as more than one main rotor diameter above the surface. This is 33 ft for the R44.
5. Standard atmospheric pressure is 1013 hPa.
6. Angle of attack is the angle between the rotor blade chord line and resultant relative wind. A larger angle of attack produces more drag and therefore requires more engine power.