AAIB Bulletin: 1/2013

minute the MGB EMLUB caption illuminated on the CWP indicating that the emergency lubrication system had failed. The crew carried out the ‘Emergency Landing – Power ON’ checklist and successfully ditched the helicopter in the sea, close to a ship. There were no reported injuries. Health and Usage Monitoring System (HUMS) HUMS trend indicator MOD-45 is used to monitor the meshing frequency of the bevel gear and indicator MOD-70 the meshing frequency of the oil pump wheels. Both indicators have thresholds which are used to generate alerts when two out of five consecutive data points exceed the thresholds. This monitoring is carried out at a ground station post flight. Figures 1 and 2 compare the MOD-45 and MOD-70 indicators for G-CHCN and G-REDW. The indicator values are plotted with respect to flying hours relative to the time at which the MGB oil pressure was lost; the period covered by each figure is 30 flying hours. Also plotted are the threshold values of these indicators unique to each helicopter1 and applicable at the time of each accident. At the time of the first accident in May 2012, the MOD-45 and MOD-70 indicators only included amber thresholds; these were ‘learned’ thresholds each with a maximum value of 0.6. After the accident to G-REDW, Eurocopter published EC225 Service Bulletin No 45-001, in July 2012 that included the introduction of a red threshold and lowered the fleet-wide maximum threshold values for both indicators. For MOD-45 the amber alert was reduced to 0.3 and a red alert of 0.4 was introduced. For MOD-70 the amber alert was reduced to 0.4 and a red alert of 0.5 was introduced. After the accident to G-CHCN, Eurocopter published an Emergency Alert Service Bulletin (ASB), on 21 November 2012, which removed the maximum amber alert threshold for MOD 45 and lowered the red alert threshold to 0.2. No change was made to indicator MOD- 70 thresholds. These maximum thresholds are greater than G-CHCN’s and G-REDW’s ‘learned’ thresholds. Both helicopters were operating within the published HUMS monitoring procedures valid at the time of their accidents. Footnote 1 These are ‘learned’ thresholds that are a function of the mean of the indicator values recorded to date. These will, therefore, vary from helicopter to helicopter. Eurocopter also publish ‘maximum’ thresholds that are applicable fleetwide which can, if required, be set sufficiently low to predominate existing ‘learned’ thresholds. 5 ©Crown copyright 2013 AAIB Bulletin: S7/2012 G-CHCN EW/C2012/10/03 Figure 1 shows that the MOD-45 indicator for G-CHCN, exceeded its ‘learned’ amber threshold (0.10) 4.75 flying hours and its ‘learned’ red threshold (0.12) 3.63 flying hours prior to the loss of oil pressure. For G-REDW the MOD-45 indicator exceeded its ‘learned’ amber threshold (0.19) 4.62 flying hours before the loss of the MGB oil pressure. Figure 2 shows that for the MOD-70 indicator, the first instance that it exceeded the ‘learned’ amber Figure 2 Comparison MOD-70 trend indications between G-CHCN and G-REDW Figure 1 Comparison MOD-45 trend indications between G-CHCN and G-REDW 6 ©Crown copyright 2013 AAIB Bulletin: S7/2012 G-CHCN EW/C2012/10/03 threshold (0.14) for G-REDW was 2.95 flying hours before the loss of MGB oil pressure. However, for G-CHCN only the last recorded value of this indicator, which was captured 1.17 flying hours before the loss of the MGB oil pressure, exceeded both its amber (0.14) and red (0.16) ‘learned’ thresholds. Aircraft information A description of the development of the Eurocopter EC225 LP helicopter and a systems description of the Main Gear Box (MGB) and the emergency lubrication system was provided in AAIB Special Bulletin S3/2012. In comparison with the AS332 L2, the EC225 LP helicopter has a five-bladed spheriflex composite main rotor and uprated Turbomeca Makila 2A1 engines that deliver approximately 15% more torque to the main rotor system. The helicopter manufacturer advised that the EC225 LP fleet has flown approximately 300,000 hours. In comparison, the AS332 variants have flown approximately 4.3 million hours. MGB bevel gear vertical shaft The bevel gear vertical shaft consists of a main bevel gear wheel and a vertical shaft that are joined together by an electron beam weld. To ensure the integrity of the weld, the disrupted material at the end of the weld is removed by drilling and reaming a diameter (Ø) 4 mm hole. The inner and outer surface of the weld region is then machined to remove the cap and root of the weld. A plug is fitted into the hole to prevent leakage of lubrication oil. The bevel gear vertical shaft is supported in the gearbox by two upper bearings (roller and ball) mounted adjacent to each other above the bevel gear wheel, and a lower roller bearing mounted at the bottom of the vertical shaft above the oil pump drive wheels. Following the failure of the bevel gear vertical shaft, the bevel gear wheel will only be supported by the two upper bearings. On bevel gear vertical shafts originally designed for the AS332 variants, both parts of the shaft are manufactured from 16NCD13 steel alloy. The gear teeth are surface hardened, by a process called carburising, prior to the bevel gear wheel being welded to the vertical shaft. The manufacturer’s design does not require the vertical shaft, or the part of the bevel gear wheel that is welded to the vertical shaft, to be surface hardened. The parent material and surface hardening process were changed for the EC225 LP to accommodate the increased loads and the elevated temperatures in the MGB during the operation of the emergency lubrication system. This was achieved by changing the parent material to 32CDV13 steel alloy and applying a different surface hardening process, called nitriding, to the teeth on the bevel gear wheel. The vertical shaft, which is also manufactured from this steel alloy, is not subject to the nitriding process. The 32CDV13 steel alloy shaft can also be fitted to the AS332 variants. On G-REDW and G-CHCN, the cracks initiated and grew to failure in areas of the vertical shaft that had not been, nor were required to be, surface hardened. The bevel gear vertical shaft has a life of 20,000 flying hours with a requirement for overhaul every 2,000 flying hours for the EC225 LP and a life of 50,000 flying hours and overhaul every 3,000 flying hours for the AS332 L2. According to the manufacturer, no shaft manufactured from 32CDV13 steel alloy has flown sufficient hours to reach its second overhaul, at 4,000 flying hours. 7 ©Crown copyright 2013 AAIB Bulletin: S7/2012 G-CHCN EW/C2012/10/03 Engineering investigation Overview of bevel gear vertical shafts fitted to G-REDW and G-CHCN The bevel gear vertical shaft (serial number M385) fitted to G-REDW was manufactured in March 2011 and had operated for 167 flying hours2, which equates to approximately 20 million shaft cycles since new3. The shaft (serial number M122) fitted to G-CHCN was manufactured in March 2008 and had operated for 3,845 flying hours which equates to approximately 553 million shaft cycles. Shaft M122 had remained with the same MGB since new and had operated for approximately 1,800 flying hours since the MGB had been overhauled. Its second overhaul was due in approximately 200 flying hours. Footnote 2 These flying hours are recorded by the flight crew in the helicopter’s technical log and are taken as the time between the wheels off and wheels on the ground. 3 A shaft cycle is defined as one rotation of the bevel gear vertical shaft, which rotates nine times faster than the main rotor. Failure of the bevel gear vertical shaft on G-REDW and G-CHCN The failure of the bevel gear vertical shaft on both G-REDW and G-CHCN occurred as a result of high cycle fatigue cracking in the area of the weld and is thought to be as a result of the shaft bending (flexing) as it rotates. On G-REDW, the first crack to develop was identified as Crack ‘A’, which initiated at a corrosion pit approximately 0.06 mm deep located in the inner countersink of the Ø 4 mm hole (Figure 3). The crack then propagated along the fusion line between the area of the weld which had been previously melted, and the heat affected zone in the parent material in the vertical shaft. A second crack, Crack ‘B’, initiated after Crack ‘A’ at a small Beachmark Crack ‘B’ Crack ‘A’ Initiation point Corrosion Beachmarks G-REDW G-CHCN 4 mm hole Inner surface Figure 3 Location of crack initiation on G-REDW and G-CHCN Figure 3 Location of crack initiation on G-REDW and G-CHCN 8 ©Crown copyright 2013 AAIB Bulletin: S7/2012 G-CHCN EW/C2012/10/03 scratch in the internal surface of the hole. Numerous corrosion pits that could only initially be detected by a scanning electron microscope were also present around the circumference of the inner countersink. Crack ‘B’ ran into a third crack identified as Crack ‘C’ (Figure 4). On G-CHCN the fatigue crack initiated at an area of corrosion on the inner surface of the shaft (Figure 3), approximately 47° around the circumference from the Ø 4 mm hole (Figure 4). It has not yet been determined where the crack initiated in relation to the heat affected zone. Away from the initiation point, the crack propagated in both directions in the parent material in the vertical shaft. Small areas of corrosion were also visible in the machining marks around the inner flange that had been machined during manufacture to remove the root of the weld. The initial examination has not identified corrosion on any other part of the bevel gear vertical shaft. The fracture surfaces of the shafts on G-REDW and G-CHCN both displayed characteristic fatigue beachmarks, which can be formed when an event such as an engine start or significant change in torque has taken place (Figure 4). Beachmarks can be difficult to identify and can be interpreted in a number of ways. For G-REDW the first beachmark was identified at 4 mm from the initiation point. Crack growth estimates are complicated by variations in engine torque and the changing stiffness of the shaft as the crack grows to failure. A number of different models are being considered, one of which suggests that the time for the crack to grow from the first beachmark to failure could be Crack ‘A’ Crack ‘B’ Crack ‘C’ 4 mm hole (Initiation point) (Initiation point) Last beachmarks Last strong beachmarks Cracks without beachmarks Cracks with beachmarks G-REDW 4 mm hole G-CHCN Figure 4 Location of beachmarks on G-REDW and G-CHCN 9 ©Crown copyright 2013 AAIB Bulletin: S7/2012 G-CHCN EW/C2012/10/03 as low as 20 engine4 hours. From the recorded data this corresponds to approximately 15 flying5 hours. Another model has shown that the crack growth time could be as high as 31 flying hours. Detailed analysis of the beachmarks continues. At present it is not possible to determine how long it took for the crack to initiate and grow to 4 mm. The growth of the crack from the last identified beachmarks could have occurred during the accident flight. For G-CHCN, work continues to identify all the beachmarks on the fracture surface. The provisional examination and analysis indicates that the first beachmark was identified at 2 mm either side of the initiation point (giving a crack length of 4 mm). Early analysis shows that there is a possible close correlation in the crack propagation time in the shafts fitted to G-CHCN and G-REDW. Condition of MGB on G-REDW and G-CHCN Examination of the MGBs fitted to G-REDW and G-CHCN identified the presence of glycol throughout and no visual evidence of heat distress or significant damage to any other components in the MGB. On G-REDW, light, unsymmetrical, marks were found on the bearing cages fitted to the upper roller and ball bearings that are believed to have occurred after the shaft failed. Part of the outer race on the lower roller bearing had broken away as a consequence of the shaft failure. Footnote 4 Engine hours are based on the first engine start to the last engine shut down, which closely corresponds to the times that the bevel gear vertical shaft is rotating. 5 The flying hours used in the HUMS were established from the operation of the air / ground switch. In the AAIB calculation the flying hours were established using recorded data from the radio altimeter. On G-CHCN, there was evidence of the rollers on the upper roller bearing having slipped along the outer race and there were light marks, similar to those seen on G-REDW, in the cage on the roller bearing. The lower roller bearing displayed no unusual marks. Ongoing investigation into the failure of the bevel gear vertical shaft The investigation has not identified the root causes of the failure of the bevel gear vertical shafts fitted to G-REDW and G-CHCN. It is possible that the failures occurred for different reasons. To date the investigation has carried out a detailed examination of the MGB and the bevel gear vertical shaft fitted to G-REDW. A component fatigue test has been carried out on a new bevel gear vertical shaft and the stresses in the component, determined using finite element modelling, have been verified against the stresses measured on a shaft run in the manufacturer’s dynamic test rig. A review of the manufacturing process of the bevel gear vertical shaft and the HUMS data from both accident aircraft has also been carried out. The investigation is currently seeking to confirm the material properties and the in-flight dynamic loads on the MGB and bevel gear vertical shaft. On-going work, some of which is anticipated to extend into 2013, includes: - Dimensional analysis, fractography and metallographic examination of the bevel gear vertical shaft and MGB fitted to G-CHCN. - Tests on parent and welded material samples (coupons) to confirm the material properties of the 32CDV13 steel alloy, used 10 ©Crown copyright 2013 AAIB Bulletin: S7/2012 G-CHCN EW/C2012/10/03 by the manufacturer in the design of the component, and the material’s susceptibility to cracking from small features. - A flight load and vibration analysis programme to confirm the predicted loads in the weld region, and to establish if there is an area in the flight envelope where the bevel gear vertical shaft might operate at one of its natural frequencies. - Examination of a sample of shafts removed from EC225 LP helicopters and an analysis of oil removed from other EC225 LP helicopters operating out of Aberdeen. Emergency Lubrication system investigation The first in-flight activation of the emergency lubrication system on the EC225 LP was during the G-REDW accident on 10 May 2012. The second in-flight activation was during the G-CHCN accident on 22 October 2012. On both occasions the EMLUB failure caption illuminated, resulting in the ditching of the helicopters. The initial examination of both helicopters revealed that the emergency lubrication system had operated. The lack of visual evidence of heat damage to the gearbox components indicates that the system had lubricated and cooled the MGB during the short period6 between the loss of oil pressure and the aircraft ditching. On 17 October 2012 the AAIB published Special Bulletin S5/2012, which contained a description of the emergency lubrication system and included the following Safety Recommendation: Footnote 6 Approximately 9 minutes for G-REDW and 7 minutes for G-CHCN. Safety Recommendation 2012-034 It is recommended that the European Aviation Safety Agency requires Eurocopter to review the design of the main gearbox emergency lubrication system on the EC225 LP Super Puma to ensure that the system will provide the crew with an accurate indication of its status when activated. Tests have been carried out on a ground test rig using a Turbomeca Makila 2A1 engine and all the parts of an EC225 bleed air system. The preliminary finding is that the bleed air pressure sensor was probably the source of the low-pressure signal that led to the MGB EMLUB caption illuminating on G-REDW. This pressure sensor has been tested and was found to operate within its specification. Tests on a complete emergency lubrication system, with and without an engine, and components from G-CHCN are ongoing. Further safety action taken by Regulatory Authorities and Eurocopter On 21 November 2012 Eurocopter issued revision 2 of EC225 ASB No.04A009 and revision 2 of AS332 ASB No.01.00.82. These were mandated by the EASA Emergency AD 2012-0250-E which superseded the previous EASA Emergency AD 2012-0225-E. These introduced operational changes and additional inspection requirements. The UK and Norwegian Civil Aviation Authorities have issued Safety Directives7 that prohibit flight in a hostile environment of AS332 and EC225 helicopters that are applicable to the EASA AD. Published 29 November 2012 Footnote 7 UK CAA SD-2012/005. 11 ©Crown copyright 2013 AAIB Field Investigation reports AAIB Bulletin: 1/2013 13 ©Crown copyright 2013 AAIB Bulletin: 1/2013 D-AEAP EW/C2012/04/04 ACCIDENT Aircraft Type and Registration: Airbus A300B4-622R(F), D-AEAP No & Type of Engines: 2 Pratt & Whitney PW4158 turbofan engines Year of Manufacture: 1994 (Serial no: 724) Date & Time (UTC): 14 April 2012 at 1753 hrs Location: East Midlands Airport, Derby Type of Flight: Commercial Air Transport (Cargo) Persons on Board: Crew - 3 Passengers - None Injuries: Crew - None Passengers - N/A Nature of Damage: Damage to nose landing gear and minor damage to a tug Commander’s Licence: Airline Transport Pilot’s Licence Commander’s Age: 53 years Commander’s Flying Experience: 5,878 hours (of which 111 were on type) Last 90 days - 79 hours Last 28 days - 20 hours Information Source: AAIB Field Investigation Synopsis After disconnection from the pushback tug, the aircraft taxied forwards into the tug before the ground crew had signalled that they and the tug were clear. There were no injuries to the two ground crew or the three aircrew but there was some damage to the aircraft and the tug. History of the flight The aircraft was on a scheduled cargo flight, departing at 1745 hrs from East Midlands Airport to Paris Charles De Gaulle Airport. It was being conducted as a training flight, with a captain-under-training in the left seat, acting as Pilot Flying (PF), and a training captain in the right seat, who was the commander of the aircraft. There was also a loadmaster onboard who was sitting in the cabin. The aircraft was parked on Stand 102 on the West Apron at East Midlands Airport, a dedicated cargo area (see Figure 1). At 1744 hrs, when the aircraft was ready to depart the commander made a radio call to ATC to request start and pushback clearance. The PF completed the Before Start scan and called for the Before Start checklist. When start clearance had been obtained, he contacted the ground crew headset operator on the flight interphone and advised him that the aircraft was ready to push. During the pushback, which commenced at 1747 hrs, the flight crew started No 2 engine. When the pushback was complete, with the aircraft facing west, the headset operator contacted the flight crew and the park brake was set. The headset operator then stood by while No 1 14 ©Crown copyright 2013 AAIB Bulletin: 1/2013 D-AEAP EW/C2012/04/04 engine was started. When both engines were running, PF told the headset operator to disconnect the tug and that he would look for his hand signals on the left side of the aircraft. The headset operator went to remove the pin from the tug end of the towbar but was unable to get the pin out of the towing eye, so he sought the assistance of the driver. Between them they withdrew the pin and disconnected the towbar from the tug. The headset operator then disconnected the towbar from the aircraft, turned his back on the aircraft and started to push the towbar to an area forward of the aircraft, to reconnect it to the rear of the tug. At the same time, the tug driver reversed the tug away from the aircraft, before driving forward to pick up the towbar. Meanwhile, the PF completed the After Start scan, with some intervention from the PNF, and called for the After Start checklist. The checklist was completed and the PF then asked the PNF to request taxi clearance from ATC. At 1753 hrs, the aircraft was cleared to taxi to the Golf 1 holding point for Runway 09, which was near the aircraft and in sight of the flight crew (see Figure 1). The PF switched on the taxi light and increased power to start taxiing. Figure 1 Cargo ramp area at East Midlands Airport 15 ©Crown copyright 2013 AAIB Bulletin: 1/2013 D-AEAP EW/C2012/04/04 As the headset operator was moving the towbar to reconnect it to the tug, he described “feeling” the aircraft above him. He pushed himself backwards off the towbar, pushing it clear of the nosewheels, and watched the aircraft pass in front of him, strike the tug and stop. He then called the flight crew on the interphone to inform them that he was still connected. The tug driver had begun to drive forward to pick up the towbar, when he realised the aircraft was moving. He attempted to drive clear but was unable to do so and the aircraft struck the rear left side of the tug, pushing it a short distance. As the tug and aircraft both stopped moving, the driver exited the vehicle and quickly moved to a safe area. The aircraft had taxied forward a total of 8 metres, before stopping. The aircraft engines were shut down a minute later, at 1754 hrs, following instructions from the headset operator. The flight crew advised ATC that they had a problem and that they could not move off the taxiway. ATC offered further assistance, which was declined. At 1801 hrs, the AFRS was notified of a ground incident and attended the scene. At 1818 hrs, ATC upgraded the incident to an aircraft accident. Damage to the aircraft and tug The aircraft’s nose landing gear drag strut was damaged and one of the drag strut attachment pins had sheared. There was also damage to the left steering actuator, on the nose landing gear leg, and damage to the nose landing gear doors. Both nosewheels were replaced. There was some damage to the bodywork on the tug which was subsequently repaired. Ground personnel and equipment The headset operator was experienced in a wide variety of airside operations but was relatively new to the job of headset operator. He had completed his training for the role a few days prior to the accident and this was his fourth shift as a headset operator. When the aircraft’s engines are started during pushback, the usual procedure is for the headset operator to walk alongside the aircraft while monitoring the starts. When the aircraft is in position on the taxiway the headset operator asks the flight crew to set the parking brake. He then waits until he receives a message from the flight crew that both engines are started and he is cleared to disconnect. He disconnects the towbar, first from the tug and then from the aircraft. Once disconnected, the tug reverses and then drives forward to the side of the aircraft, to allow the towbar to be re-attached to the rear of the tug by the headset operator. When the towbar is re-attached, the tug is driven clear of the aircraft to a position in full view of the flight crew. The headset operator then disconnects his headset from the aircraft and walks to the side of the aircraft, to a position from which he can signal to the flight crew that the ground equipment and ground crew are clear. Recorded information The aircraft’s flight data recorder (FDR) and cockpit voice recorder (CVR) were removed from the aircraft and downloaded at the AAIB. A closed-circuit television (CCTV) camera recording of the pushback was also available. This captured most of the pushback, all the tug movements, the moment of the collision and the subsequent ramp activity. The FDR recordings of the engine EPR and thrust lever angle parameters were only sampled every four seconds. Also, there was no requirement to record parameters 16 ©Crown copyright 2013 AAIB Bulletin: 1/2013 D-AEAP EW/C2012/04/04 associated with the use of aircraft’s brakes. However, the recording of the longitudinal acceleration enabled the calculation of the speed of the aircraft at the point of collision (2.4 m/s = 4.7 kt) and the distance travelled (8 m) (see Figure 2). Tug information The tug was 6 metres long, 2.5 metres wide and 1.8 metres high. It was fitted with an orange high visibility light which was on at the time of the collision. Its mass was 30 tonnes. Aircraft information The flight interphone system is selected on the radio panel and when selected all parties can hear the communications. Thus the headset operator could hear all the communications made between the pilots as well as those addressed to him. Visibility from flight deck The forward and downward visible segment from the flight deck to the ground is restricted by the aircraft Figure 2 FDR derived groundspeed and distance 17 ©Crown copyright 2013 AAIB Bulletin: 1/2013 D-AEAP EW/C2012/04/04 structure. With the pilot’s seat position adjusted for flight, a 1.8 m high object (the height of the tug) becomes visible when approximately 14 m in front of the pilot. Organisational and management information The operations manual contained, as the final item of the After Start procedure and before the After Start checklist, a check for a hand signal from the ground crew on one or other side of the aircraft. The After Start checklist did not contain an item to check that the hand signal had been received. There are a variety of Standard Operating Procedures (SOP) that have been developed by operators in order to prevent aircraft moving before ground equipment and personnel are clear. For example, some operators do not allow the After Start checklist to be read until both pilots have seen the hand signal, some include ‘ground crew clear’ as an item on the after start checklist and some do not allow a call to be made to ATC for taxi clearance until the ground crew’s hand signal has been seen. Other information The field of view from the flight deck of a large aircraft is very restricted and some aircraft are equipped with under body cameras to increase the flight crew’s situational awareness. Often, pushback tugs are fitted with a rotating light to increase their visibility. However, this is mainly effective at night. Some tugs are fitted with a flag on an extended aerial, above the cab, to increase their conspicuity to pilots in daylight conditions. Discussion The incident occurred in daylight and clear weather conditions on a dry ramp surface. The pushback proceeded without incident until the headset operator was unable to withdraw the towbar pin from the tug. With the assistance of the driver, it was then removed but the process of disconnecting the tug took longer than usual. The PF advised the headset operator that he would expect to see him on the left side of the aircraft, for hand signals. Once both the engines had been started, the PF carried out the After Start scan, during which he was corrected in some actions by the PNF. This took a little extra time and may have caused a distraction, possibly resulting in the omission of the action of waving off the headset operator. After completing the scan, the PF called for the After Start checklist and, immediately afterwards, for clearance to taxi, omitting the check for a hand signal from the ground crew at the side of the aircraft. Analysis of the CCTV, recorded flight data and the field of view Figure 3 Field of view from pilot’s seat 18 ©Crown copyright 2013 AAIB Bulletin: 1/2013 D-AEAP EW/C2012/04/04 from the flight deck indicated that it is likely that the tug and ground crew were not visible to the pilots when the aircraft started to taxi. Ground personnel carrying out tasks close to an aircraft, together with any equipment, may well be out of sight of the flight deck. The safe completion of the pushback procedure relies on the flight crew being certain that all is clear before starting to taxi. This is achieved by the means of SOPs for flight and ground crews. On this occasion, the final safety element relied on the flight crew observing a hand signal before starting to taxi. While there may be equipment available that increases the area observable from the flight deck, the operation is dependent upon an operator’s SOPs being robust. Safety action The operator has reviewed its procedures since the event. A revised After Start checklist has been introduced which includes ‘HAND SIGNAL .......... RECEIVED’ as the final check. 19 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-EZFV EW/C2012/02/03 ACCIDENT Aircraft Type and Registration: Airbus A319-111, G-EZFV No & Type of Engines: 2 CFM56-5B5/3 turbofan engines Year of Manufacture: 2010 (Serial no: 4327) Date & Time (UTC): 14 February 2012 at 1359 hrs Location: London Luton Airport Type of Flight: Commercial Air Transport (Passenger) Persons on Board: Crew - 6 Passengers - 142 Injuries: Crew - None Passengers - None Nature of Damage: All landing gear legs exceeded their maximum certified load Commander’s Licence: Airline Transport Pilot’s Licence Commander’s Age: 45 years Commander’s Flying Experience: 10,700 hours (of which 500 were on type) Last 90 days - 128 hours Last 28 days - 61 hours Captain U/T’s Licence: Airline Transport Pilot’s Licence Captain U/T’s Age: 37 years Captain U/T’s Flying Experience: 3,998 hours (of which 672 were on type) Last 90 days - 170 hours Last 28 days - 19 hours Information Source: AAIB Field Investigation Synopsis The flight crew carried out a manually flown ILS approach to Runway 26 at London Luton Airport. Shortly before touchdown, both pilots sensed the aircraft was sinking and a go-around was initiated. The aircraft made firm contact with the runway before starting to climb. The normal acceleration recorded at touchdown was 2.99g, which is classified as a Severe Hard Landing. The subsequent landing was uneventful. All three landing gear legs exceeded their maximum certified loads and were replaced; there was no other damage to the aircraft. History of the flight The aircraft was on a scheduled flight to London Luton Airport, from Faro, Portugal. The pilot flying (PF) was a captain-under-training (Capt U/T), occupying the left seat; the right seat was occupied by a training captain, who was the commander of the aircraft. 20 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-EZFV EW/C2012/02/03 The aircraft was inbound to Luton from the south on the LOREL 4C standard arrival procedure. This procedure requires the aircraft to cross the extended runway centreline, before positioning for the Runway 26 ILS/DME approach from the north. When traffic allows, ATC will vector the aircraft towards the final approach course before the arrival procedure is completed, thereby shortening the track mileage to the landing. The flight crew were familiar with the airport procedures and were prepared for this to happen. The aircraft was given an early radar vector towards the final approach track and the PF increased the rate of descent to close the correct descent profile from above. The aircraft was then allocated a heading of 220°M, cleared to intercept the localiser and, once established, to descend on the glidepath. The PF realised that the aircraft would be high and configured the aircraft with flap 2 and the landing gear down, to capture the 3º glideslope from above. He armed the localiser mode and then attempted to arm the approach mode but inadvertently selected the EXPED1 pushbutton. The expedite climb mode engaged but, to prevent a climb or any mode confusion and to regain the correct profile, the PF disconnected the autopilot and the autothrust. The aircraft passed through the localiser and ATC issued a revised heading to enable the aircraft to intercept from the south. The PF decided to continue flying the approach manually and the aircraft was established on the localiser at 5.5 nm. It was configured for landing, with full flap, at 5 nm. Landing clearance was issued at 1355 hrs, with a reported surface wind of 320°/15 kt. A subsequent wind check of 320°/16 kt was broadcast at 1356 hrs, three minutes before touchdown. The wind conditions were gusty and gave rise to some turbulence on the approach. Footnote 1 EXPED - Expedite mode is used in climb or descent to reach the desired altitude with the maximum vertical gradient. Stabilised approach criteria were met at 1,000 ft and 500 ft radio altitude (RA). The VAPP (final approach) speed was 129 kt and at 50 ft RA the approach remained stable. Just below 50 ft there was a small nose-up pitch input followed by two nose-down inputs and, below 50 ft, the flight data indicated an increasing rate of descent from about 600 fpm to about 850 fpm. The data also showed that, below 100 ft RA, there were some left and right roll control inputs. Below 30 ft, over the runway, both pilots sensed that the aircraft was sinking rapidly and both initiated a TOGA 102 go-around. The PF momentarily retarded the thrust levers to idle before advancing them to the TOGA (Takeoff and Go-around) position. At the same time, he made a full forward sidestick input, within one second, which was then rapidly reversed to full aft sidestick. As the PF made the forward sidestick input, the commander initiated an aft sidestick input which reached the full aft position within one second. He followed through the PF, pushing the thrust levers fully forward and announced “I HAVE CONTROL”. The aircraft made firm contact with the runway, on all three landing gear legs simultaneously, before lifting off and starting to climb. During this phase the PF relinquished control and reverted to the PNF role. The commander remained as the PF, completed the go-around and subsequently carried out an uneventful landing on the same runway. There were no reported injuries. Flight crew information The Capt U/T had completed nine sectors of command training without notable incident and the training reports prior to the event had all been positive. His command Footnote 2 TOGA 10 Baulked landing procedure. 21 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-EZFV EW/C2012/02/03 training had included practice in TOGA 10 manoeuvres in the simulator but he had never carried out a TOGA 10 manoeuvre in the aircraft. The commander had previous experience of line training on another aircraft type, but was relatively inexperienced in this capacity on the Airbus 320 series aircraft. He reported that he had practised TOGA 10 manoeuvres in the simulator but this was his first experience of one in the aircraft. Meteorological information The ATIS information issued at 1320 hrs, and copied by the crew, was surface wind from 300°M at 12 kt, CAVOK, temperature 7°C, dewpoint 3°C and pressure 1024 hPa. There was no significant change between 1320 hrs and 1400 hrs. Wind data was obtained from the Runway 26 touchdown zone sensor. Readings are taken every 10 minutes and include the average direction and speed reported, along with min/max variations. The readings for 1400 hrs were average wind direction from 315°M, varying between 297°M and 342°M, and average wind speed 13 kt, varying between 9 kt and 17 kt. The crew were aware that some turbulence can be expected on the final approach to Runway 26 when the wind is from the north-west. Baulked landing procedure The operator provides the following Baulked Landing Procedure, entitled TOGA 10, in its operations manual. Either pilot may carry out this manoeuvre. ‘In the event of a rejected landing from flare initiation until thrust reverser selection • call “TOGA TEN” • select TOGA • pitch to 10° (this may mean holding the attitude or de-rotating to achieve or maintain 10° pitch.) • do not retract the flaps until a positive rate of climb is established. • When positive ROC confirmed, call “GO AROUND FLAPS” and apply normal go around procedure.’ Through its Flight Data Monitoring programme, the operator has carried out several analyses of TOGA 10 manoeuvres conducted by its crews. The most recent analysis, which was completed following a programme of TOGA 10 training for crews in the simulator, examined 67 events. The report concluded that the manoeuvre was generally well flown and there were no significant dual inputs recorded. Aircraft information Autothrust The manufacturer provides the following information concerning the use of autothrust: ‘The A/THR is, in particular, best suited to tracking a moving target speed, when flying in managed speed mode. Statistically, the A/THR provides the best protection against airspeed excursions and its use is, therefore, recommended even in turbulent conditions, unless thrust variations become excessive. 22 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-EZFV EW/C2012/02/03 A/THR response to airspeed variations is the result of a design compromise between performance and comfort, and it is optimized when the AP is engaged. Therefore, in turbulent conditions and when flying manually, the pilot may sometimes find it to be too slow or lagging.’ Sidesticks There is a sidestick for each pilot, located outboard of the seating position. There are two switches on the sidestick, one of which is the autopilot disconnect and sidestick takeover pushbutton. The sidestick controls move independently, so one pilot may not be aware of a control input being made by the other. The manufacturer advises: ‘When the Pilot Flying (PF) makes an input on the sidestick, an order (an electrical signal) is sent to the fly-by-wire computer. If the Pilot Not Flying (PNF) also acts on the stick, then both signals/ orders are added.’ And: ‘If the PNF (or Instructor) needs to take over, the PNF must press the sidestick takeover pushbutton, and announce: “I have control”.’ Further: ‘In the event of simultaneous input on both sidesticks the two green SIDE STICK PRIORITY lights on the glareshield come on and “DUAL INPUT” voice message is activated. A pilot can deactivate the other stick and take full control by pressing and keeping pressed his priority takeover pushbutton.’ The operator provides the following guidance for flight crew in their operations manual: ‘If a take-over becomes necessary during flight, the PNF must call clearly “I have control”, and press the sidestick priority pushbutton, keeping it pressed until the transfer of control is clearly established. During critical phases of flight the PNF should be in a position to takeover, this may be achieved by resting the hand on the console or indeed on the stick itself but it is imperative that no input is made on the sidestick.’ The use of the takeover pushbutton has been shown from previous incidents not to be instinctive.3 Training in taking over control, including the use of the takeover pushbutton, is provided by the operator. Flare Mode When the aircraft descends through 50 ft RA it enters Flare Mode. The manufacturer’s description is: ‘The system memorizes the attitude at 50 ft, and that attitude becomes the initial reference for pitch attitude control. As the aircraft descends through 30 ft, the system begins to reduce the pitch attitude, reducing it to 2° nose down over a period of 8 s. This means that it takes gentle nose-up action by the pilot to flare the aircraft.’ Engineering investigation The A319 is fitted with a system that senses when landing parameters have been exceeded and generates a LOAD<15> report, following which inspection of the aircraft for damage is required. Footnote 3 Ref ; AAIB Bulletin No: 11/2004 Airbus A320, C-GTDK, AAIB Bulletin No: 5/2001 Airbus A321, D-AIRE and Airbus A321-211, EI-CPE. 23 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-EZFV EW/C2012/02/03 A LOAD <15> report will automatically be generated during a landing if any of the following conditions are met: • The normal acceleration is greater than 2.6g at touchdown (+/-0.5 second). If the aircraft weight exceeds the maximum landing gross weight, the normal acceleration limit is reduced to 1.7g. • The rate of descent on the radio altimeter is greater than 9 ft/sec at touchdown (+/ 0.5 second). If the aircraft weight exceeds the maximum landing gross weight, the radio altimeter descent rate limit is reduced to 6 ft/sec. • During a bounced landing, the normal acceleration exceeds 2.6g. The normal acceleration parameter used within the LOAD <15> report computation is provided by an accelerometer mounted near to the aircraft’s centre of gravity; the same accelerometer is used by the flight data recorder (FDR) system. The accelerometer incorporates a filter that attenuates its output above a predefined frequency. Under certain conditions, such as during rapid changes in acceleration, the accelerometer output may not always reflect the maximum attained normal acceleration level. In addition, during various phases of flight, acceleration levels experienced by other areas of the airframe, such as the nose gear, may be different from those measured at the centre of gravity. A LOAD <15> report for the incident landing was automatically generated shortly after the hard landing, having recorded a normal acceleration of 2.99g and a rate of descent of 12.5 ft/sec. Aircraft examination The operator sent the FDR data to the manufacturer. Due to the high level of vertical acceleration and the fact that the aircraft had made a three-point landing, the manufacturer requested a comprehensive list of structural inspections which included several areas of the fuselage, the belly fairing, the pylons, the horizontal stabilizer and the wings. No damage was found in these areas. The manufacturer’s analysis confirmed that the following components had exceeded their design loads and needed to be replaced: • Nose Landing Gear shock absorber. • Left Main Gear sliding tube assembly (including shock absorber internals) • Right Main Gear Main fitting, including both pintle pins • Right Main Gear sliding tube assembly (including shock absorber internals) The operator replaced all three landing gear legs, which were returned to the landing gear manufacturer for overhaul. Recorded flight data The aircraft’s FDR and CVR were removed from the aircraft, downloaded and the recordings were analysed by the AAIB. Recordings on the CVR at the time of the hard landing had been overwritten by more recent recordings while the aircraft was on the ground, after the final landing. From the FDR data it was determined that at 8.3 nm DME range from Luton and a height of 3,500 ft agl, the Expedite Climb Mode was selected. Three seconds 24 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-EZFV EW/C2012/02/03 later, the autothrust was disconnected, followed by the autopilot. The Expedite Climb Mode was deselected after 10 seconds. The aircraft was established on the localizer at 5.5 nm DME and 2,000 ft agl. The flaps were fully extended by 5.0 nm DME. Figure 1 shows the salient recorded parameters approaching the touchdown. The data illustrated starts at 1358:38 hrs, with the aircraft at 200 ft agl, 130 kt indicated airspeed and descending at 625 ft/min. At this point the aircraft’s attitude was 3º nose-up and 2º left wing low. Left-seat sidestick control inputs were made which resulted in the aircraft rolling wings level as it descended though 100 ft agl. The wings remained level for just over one second before further inputs rolled the aircraft right (to 4.5º at 60 ft agl), then left (to 5.5º at 21 ft agl), then back to wings level at touchdown. As the aircraft descended through 60 ft agl, at 1358:46 hrs, a momentary 2º reduction in angle of attack was recorded while the pitch attitude remained steady at 3.5º nose-up. A maximum pitch attitude of 4.2º occurred at 32 ft agl, 2.5 seconds before touchdown. This was followed by some pitch-down control inputs (from the left seat) which resulted in a decreasing pitch attitude, the descent rate peaking at just under 900 ft/min, and the aircraft descended below the nominal glideslope. The left-seat pilot then applied an aft sidestick control input of 10º (out of a maximum of 16º) which was coincident with a small aft-stick input by the commander. Both thrust levers were then advanced to the TOGA position. However, there was also a simultaneous forward sidestick input of 15º from the left seat pilot, which was countered by an aft input of 8º by the commander (ie a net input of 7º forward). Both sidesticks were then moved to the fully aft position. Half a second later the aircraft touched down on all three landing gear simultaneously, during which a normal acceleration of 2.99 g was recorded. The aircraft then lifted off, completed a go-around and returned for an uneventful landing. The manufacturer’s analysis of the data provided the following information: ‘Inputs performed simultaneously by pilots were equivalent to a pitch down order at ~+7° of side sick deflection leading to a strong pitch decrease.’ Analysis In the early stages of the approach there was a period of increased workload for the PF, caused by the aircraft being above the nominal 3º descent profile. This was exacerbated when the PF inadvertently selected the Expedite Mode, instead of the Approach Mode, and missed intercepting the localiser. He recovered from this by reverting to manual flight and the aircraft was re-established on the expected profile by 5 nm. This does not appear to have directly affected the outcome of the approach other than that it led to the PF’s decision to use manual thrust for the remainder of the approach, which increased the workload in turbulent conditions. The manufacturer advises the use of autothrust in most circumstances but also notes that there are conditions in which autothrust may not be the best option. Thus, it is accepted that a pilot may need to use manual thrust. Once established on the ILS in the landing configuration, the remainder of the approach was stable and it was only at a late stage that it deviated from what would normally be expected. There was a momentary reduction in angle of attack at about 60 ft agl and an increasing rate of descent below around 50 ft agl, with an increasing 25 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-EZFV EW/C2012/02/03 Figure 1 Salient FDR parameters for incident touchdown at Luton 26 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-EZFV EW/C2012/02/03 pitch attitude. The PF then made two brief nose-down inputs and the aircraft’s pitch attitude decreased. One or all of these factors may have provided the cue to the pilots that the aircraft was sinking. As the pilots responded to the sink, a period of dual sidestick inputs was recorded. The dual input phase lasted for approximately four seconds before the commander established sole control and took over as PF. Initially these dual inputs were in opposing directions, before the Capt U/T’s input was reversed. The commander made a nose-up control input on the sidestick but did not use the takeover pushbutton to establish control, so the effect was limited to reducing the magnitude of the nose-down input made by the Capt U/T. If the commander had used the sidestick takeover pushbutton the severe hard landing may have been prevented. The brief nose-down inputs made by the Capt U/T occurred at a time when a nose-up control input would normally be expected and probably took the commander by surprise. The sidesticks move independently. So he would have had no knowledge of the inputs being made by the Capt U/T until the flightpath of the aircraft changed. In attempting to carry out the TOGA 10 manoeuvre, the Capt U/T appears to have made a sidestick input opposite to that expected and there was also a brief retardation of the thrust levers before they were pushed forward to the TOGA position. One possible explanation is that there was momentary confusion between the actions of his left and right hands. When the aircraft entered the Flare Mode at 50 ft the pitch attitude was 3.2° nose-up. The system would have ordered a nose-down pitch to reach 2° nose-up over a period of 8 seconds. However, this would have been a relatively gradual change and was not considered to be a significant factor in this event. Following this event, the operator provided additional simulator training for both pilots before returning them to line flying duties. The Capt U/T was returned to line flying as a co-pilot for a period. Conclusion Both pilots responded to an increased rate of descent approaching touchdown and each initiated a TOGA 10 go-around. Their initial sidestick inputs were in opposition and, without the use of the takeover sidestick pushbutton, the net effect was a pitch-down control input. If the commander had operated the sidestick takeover pushbutton, his nose-up pitch input would not have been counteracted by the nose-down input of the Capt U/T. In the event, his control input reduced the effect of the nose-down input made by the Capt U/T. 27 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-SUEZ EW/C2012/02/05 SERIOUS INCIDENT Aircraft Type and Registration: Agusta Bell 206B Jet Ranger II, G-SUEZ No & Type of Engines: 1 Allison 250-C20 turboshaft engine Year of Manufacture: 1970 Date & Time (UTC): 20 February 2012 at 1500 hrs Location: Approx 3.4 miles NW of Perth, Scotland Type of Flight: Aerial Work Persons on Board: Crew - 1 Passengers - 1 Injuries: Crew - None Passengers - None Nature of Damage: Engine compressor and case damaged Commander’s Licence: Commercial Pilot’s Licence Commander’s Age: 36 years Commander’s Flying Experience: 1,891 hours (of which 1,020 were on type) Last 90 days - 168 hours Last 28 days - 38 hours Information Source: Aircraft Accident Report Form submitted by the pilot and subsequent AAIB enquiries Synopsis After suffering an engine failure at 600 ft agl the helicopter completed a successful autorotation into a field. The engine failure was a result of a fracture in fatigue of a stage-two compressor blade. History of the flight The helicopter was carrying out a pipeline inspection approximately 2 nm north-west of Perth. At 600 ft agl a loud bang was heard by the crew and the helicopter yawed to the left. The main rotor rpm decreased and the engine was seen to “wind down”. The pilot completed a successful autorotation into a field after which he reported the incident to ATC. There were no injuries. Investigation Initial inspection revealed that the engine had suffered from a failure of the axial compressor and the compressor case had been breached. No damage was found to any other components or structure of the helicopter. The engine was removed and inspected at the operator’s maintenance organisation where numerous fragments of compressor blades and stator vanes were recovered. The engine was then dispatched to an approved overhaul facility where it was examined under the supervision of the AAIB and a representative of the engine manufacturer. Several ruptures were observed to the compressor case in the plane of the stage-two and three compressor discs. Removal of the compressor cases showed that 28 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-SUEZ EW/C2012/02/05 all the blades on stages two through to six of the axial compressor had separated from their associated discs. The stage-one compressor blades remained attached to the disc but had suffered from significant trailing edge damage. The stage-two and three compressor stator vanes had been heavily damaged and distorted. Most of the stator vanes from subsequent stages had separated from the compressor case. The centrifugal impeller showed signs of impact damage but was intact. Impact damage was observed on the compressor diffuser tubes and the turbine. The extent of the damage to the compressor assembly prevented the identification of the cause of the failure using optical examination so it was dispatched to the engine manufacturer for a detailed investigation. Metallurgical tests of the compressor components confirmed that no material abnormalities were present and there was no evidence of Foreign Object Damage (FOD). The fracture surface of a large number of the compressor vanes had been smeared during the incident which prevented their initial failure mode from being identified. Scanning Electron Microscope (SEM) examination of the remaining fracture surfaces showed some features indicating that they had failed due to tensile overload. The compressor blade fracture surfaces had also suffered from significant secondary damage, but most contained localised areas where the initial fracture surface was visible. SEM examination of these areas confirmed that these compressor blades had failed due to overload. However, a section of a fracture face on one stage-two compressor blade had evidence of crack progression in fatigue. Further examination suggested that the crack had propagated in High Cycle Fatigue (HCF) from the suction side of the blade. Secondary damage to the majority of the fracture surface prevented the initiation point of the crack from being identified. Further SEM examination revealed pits in the leading edge blade root area of several stage-two and three compressor blades, the largest of which was 0.0053” deep and 0.0055” wide. Maintenance requirements The manufacturer’s Maintenance and Operation manual (72-00-00 page 617) for the M250 engine contains a 300-flying hour inspection which includes a task that states: ‘Inspect the compressor case when operating in an erosive and/or corrosive environment.’ The United Kingdom is considered to be a corrosive environment. This task makes reference to Paragraph 1.D (9), 72-00-00, which states: ‘(9) Erosion and Corrosion Inspection If the aircraft is frequently subjected to sand or dust ingestion or operated in a corrosive environment (salt laden or other chemically laden atmosphere such as pesticides, herbicides, sulphur, industrial pollutants, etc), inspect compressor blades, vanes, and case plastic coating for erosion or corrosion damage. Engines operated in a corrosive environment should be subjected to daily water compressor rinses.’ There is no requirement to use additional optical magnification when completing the visual inspection. The compressor blade and vane inspection limits are specified in section 72-30-00 paragraph 5 of the Maintenance and Operation Manual. 29 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-SUEZ EW/C2012/02/05 Maintenance activity A review of the maintenance records for the helicopter confirmed that the compressor had been installed in the engine on 19 August 2011 and had operated 341 flying hours prior to the failure. The compressor had previously been installed on an engine fitted to a helicopter that had been operating outside the UK. Before its installation into G-SUEZ, the engine’s compressor had undergone a 300-hour inspection. A 300-hour inspection had been subsequently completed on the compressor in January 2012. The maintenance organisation’s 300-hour inspection for the Allison 250 engine contained the task shown in Table 1. Whilst the required maintenance task for the inspection of the compressor was detailed in the helicopter’s maintenance programme and provided a generic reference to the engine Maintenance and Operation Manual, it did not provide a reference to the specific section of the manual which detailed the full inspection requirements. In discussions with representatives of the maintenance organisation it became apparent that the use of a generic reference and the description of the inspection task in the maintenance programme was ambiguous. The maintenance organisation confirmed that G-SUEZ was subject to routine compressor washes as part of the normal daily maintenance requirements when operating from its maintenance facility. At the time of the incident, G-SUEZ had been operating away from its main maintenance base for two days and had not received a compressor wash during this period. It was not possible to determine how frequently the compressor had been washed prior to its installation in G-SUEZ. Analysis The damage observed to the compressor was consistent with a failure within the stage-two compressor rotor which resulted in significant downstream damage to the engine. The examination of the remains of one stage-two blade root indicated that the probable cause of the event was the fracture of a stage-two blade due to crack progression in fatigue. Whilst the origin of this crack could not be identified, there was no evidence of Foreign Object Damage to the stage-one compressor blades, or to the inlet guide vanes. It was not possible to eliminate the possibility of the presence of pitting which had been observed on other blades. If pitting was present this could have been a potential initiator of the fatigue crack. The manufacturer’s inspection programme for the engine type includes visual inspections of the compressor rotor DATA REFERENCE INSPECTION TASK DESCRIPTION Mech Insp Allison 250 M & O Inspect the compressor case halves. NOTE: only required if flown 300 hours since last compressor split – see six monthly inspection Table 1 30 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-SUEZ EW/C2012/02/05 for damage and pitting during the 300-hour compressor case inspection task. Whilst wording of the compressor inspection task in the maintenance organisation’s 300-hour inspection programme reflected the wording of the manufacturer’s manual, it did not provide the reference to the specific tasks associated with the inspection requirements contained in that manual. This lack of references were such that the requirements were ambiguous, and therefore could result in an incomplete visual inspection of the compressor rotor. Safety action As a result of this investigation, the maintenance organisation has revised its maintenance programme to include a specific task for the inspection of the compressor rotor during the 300 hour inspection. In addition, the inspection task now includes specific references to the sections of the manufacturer’s manual which lay out the inspection criteria and limitations. 31 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-CWIK EW/C2012/05/02 ACCIDENT Aircraft Type and Registration: Pegasus Quik, G-CWIK No & Type of Engines: 1 Rotax 912 ULS piston engine Year of Manufacture: 2004 (Serial no: 8018) Date & Time (UTC): 12 May 2012 at 1013 hrs Location: 100 ft below summit of Ben More, Stirlingshire, Scotland Type of Flight: Private Persons on Board: Crew - 1 Passengers - 1 Injuries: Crew - 1 (Fatal) Passengers - 1 (Fatal) Nature of Damage: Aircraft destroyed Commander’s Licence: Private Pilot’s Licence (Microlights) Commander’s Age: 63 years Commander’s Flying Experience: 826 hours (of which 1 was on type) Last 90 days - 12 hours Last 28 days - 6 hours Information Source: AAIB Field Investigation Synopsis The aircraft was being flown by an experienced microlight pilot accompanied by the owner, who was a passenger, occupying the rear seat. They were transiting from Perth to Glenforsa, on the Isle of Mull, at about 6,000 ft, above scattered cloud. Approximately 2 nm east of Ben More mountain, in Stirlingshire, the aircraft descended in good visibility, remaining clear of the cloud. The descent and flight up to one second before impact was recorded on a video camera attached to the aircraft. The aircraft levelled off below the cloud base and approximately 100 ft above the summit of the mountain. It continued towards the mountain and encountered severe turbulence in the lee of the summit. This appeared to cause the pilot to lose control of the aircraft, which impacted the south side of the summit, fatally injuring both occupants. History of the flight A group of friends had agreed to fly from Perth Airport to Glenforsa, an airfield on the Isle of Mull, using four weight-shift microlight aircraft. The owner of G-CWIK had purchased the aircraft in October 2011 and was taking flying lessons in it. On the day of the accident the pilot and the owner (his passenger) arrived between 0700 and 0730 hrs and prepared their aircraft. G-CWIK had been refuelled the day before the accident and at about 0800 hrs the group met to discuss the flight. They would not be flying in formation, or as an organised stream, 32 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-CWIK EW/C2012/05/02 that the aircraft descended slightly, to about the same height as the summit and heading directly towards it, as if to pass over the top. The flight path appeared stable until about 300 m before the summit, when the aircraft began rolling from side to side, with some pitching motion. The engine speed increased significantly and the aircraft banked rapidly left and right and then pitched rapidly nose-down before impacting the mountain side. A witness on top of Ben More saw the last moments of the aircraft’s flight but did not see or hear the impact. He described the wind at the summit as “very strong” and that when he removed an item of clothing from his rucksack it was nearly “ripped” out of his hand by the wind. He did not realise that the aircraft had crashed; the noise of the wind had probably masked the sound of the impact. Shortly after this he met two other hill walkers and they came across the wreckage some time later. They reported the accident to the police, who mobilised the Search and Rescue response. Both occupants had been fatally injured. Aircraft description The Pegasus Quik is a tandem two-seat weight-shift microlight, powered by a Rotax 912 ULS piston engine driving a Warp Drive three-bladed propeller (Figure 1). Figure 1 Accident aircraft G-CWIK but would make their way independently, meeting at Glenforsa for lunch. According to the ATC movements log, the group of aircraft departed to the west, between 0917 and 0927 hrs with G-CWIK departing at 0920 hrs. The pilots described the weather at Perth when they departed as having good visibility, with scattered clouds at about 4,000 ft. One of the pilots later reported that he initially climbed to 4,000 ft, where he estimated, from his GPS groundspeed and his indicated airspeed, a headwind component of about 15-20 mph with moderate levels of turbulence. Due to the turbulence, he climbed to between 6,000 and 7,000 ft, where the flying conditions were smoother. G-CWIK was last seen by one of the other aircraft at about 6,000 ft to the northeast of Ben More, where it was seen to descend. A GoPro Hero video camera, attached to G-CWIK and facing forward in the direction of flight, was later used by the investigation to reconstruct the later stages of flight. The final camera recording commenced 3 minutes and 21 seconds before impact and showed the aircraft descending above a small patch of stratus cloud, with the snow-capped summit of Ben More (elevation 3,850 ft amsl) clearly visible through a gap. The local terrain was visible in sunshine with the slow-moving shadows of the scattered cloud. There was no smoke or other visual means to indicate the direction and strength of the wind, and no snow ‘spindrift’1 was being blown from the summit. The aircraft manoeuvred to the right and left avoiding entering cloud and then passed clear of the edge of a cloud, heading towards the top of the mountain. The aircraft levelled off and engine speed increased just below the cloud base, which was about 300 ft higher than the summit. The recording shows Footnote 1 Spindrift is the movement of the surface snow particles due to the effect of the wind. 33 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-CWIK EW/C2012/05/02 G-CWIK was fitted with the optional electric pitch trim system and the GoPro Hero video camera was mounted on the forward strut. The Permit to Fly had been renewed on 3 May 2012 and the airframe and engine had accumulated 438 hours. The wing, which had been replaced in July 2010 following an accident, had accumulated 62 hours. Accident site and wreckage examination The aircraft had struck Ben More mountain on its south-eastern side 100 ft below the summit (Figure 2). The accident site was consistent with the aircraft having hit a small rock in a steep nose-down attitude with some left bank. The nosewheel and parts of the nose structure were embedded in the ground by the rock and the main aircraft wreckage was lying inverted 8 m away from the rock in the direction of 225°(M). The fuel tank had split and was empty but there was a distinct smell of fuel at the accident site. All three propeller blades had failed near the root. The wreckage was recovered from the mountain by helicopter on 16 May 2012 and then transported to the AAIB’s facility in Farnborough for more detailed examination. All the failures within the airframe and wing structure could be explained as a result of impact forces. The pylon had failed aft due to buckling loads which permitted the propeller to strike the wing. The aft end of the keel and the aft end of the fin tube had been deformed as a result of propeller strikes indicating significant energy in the propeller. All failures within the rigging were due to overload resulting from impact forces or propeller strikes. The electric motor for the pitch trim system was found set to ‘six turns’. According to the aircraft manufacturer this trim setting, with two occupants, would result in an approximate trimmed airspeed of 60 to 65 mph. Figure 2 Accident site location, 100 ft below summit of Ben More (image extracted from video camera fitted to G-CWIK) 34 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-CWIK EW/C2012/05/02 The lap straps from both seats had failed in overload. However, the harnesses on UK microlights are only required to restrain occupants in the case of 9.0g forward loading and 4.5g upward loading – the impact loads in this accident would have been considerably higher. Recorded data Devices from the aircraft A number of electronic devices were recovered from the accident site, including a GPS eTrex Legend C and a GPS-enabled iPad. However, the only relevant recordings that were recoverable were from the memory card of the GoPro Hero video camera. Two video files were recovered from the video camera, both taken in the air during the accident flight. The first covered a period of one minute and four seconds while approximately 25 km east-south-east of the accident site. The second video file had not been completed properly, indicative of a loss of power, and required forensic techniques to make it playable. This video was three minutes and 21 seconds long and ended with the aircraft in a steep nose-down attitude, visually estimated to be 10 to 20 ft above the ground, within 20 m of where the main wreckage was found. The video images provided good evidence of the weather conditions and flight path, shown in Figures 3 and 4 and described in the ‘History of the flight’ section of this report. The Figure 3 images at ‘6 seconds’ and ‘5 seconds’ indicate a roll rate of about 55º/sec and subsequent images showed a nose-down pitch. Analysis of the recorded audio showed clear engine-related signatures. The engine speed varied for the bulk of the recording and towards the end increased in increments until reaching the maximum continuous speed of 5,500 rpm, 40 seconds before the end of the recording. This was maintained for 9 seconds before increasing to the redline speed of 5,800 rpm. 10 seconds before the end of the recording the engine speed increased to approximately 6,090 rpm. 3.6 seconds from the end of the recording, the audio signatures stopped, returned and then disappeared once more, coincident with moments of more extreme attitude apparent from the video images. Radar Radar return recordings from Kincardine and Lowther Hill radar heads were provided by the national provider of air traffic services, NATS. The aircraft was not using an ATC transponder so could only be tracked using primary radar. Microlight aircraft do not present a strong primary radar target and intervening terrain between the aircraft and the radar heads caused further problems in reconstructing the flight, resulting in parts of the flight path not being detected by radar and the other parts being subject to large errors. The last recorded radar return relating to the accident aircraft was 2.2 km east-north-east of the accident site. The radar data included sporadic coverage of the other microlight aircraft in the area, showing them generally flying several kilometres apart, following different paths. This concurred with GPS tracks recovered from other microlight aircraft involved in the journey. A secondary radar track from a helicopter in the area at the time was also reviewed. The helicopter flew from the south-east, between Ben More and the adjacent peak, below the height of the peaks, and then to the north-west (Figure 4). Photographs and video taken from this helicopter at about the time of the accident were reviewed but did not capture the accident aircraft. 35 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-CWIK EW/C2012/05/02 Combined data The final section of the radar track of the accident aircraft was consistent with the position of the aircraft established by analysis of the video. The correlation was used to derive the approximate timings for the video (Figure 4). The paths and timings of the helicopter and the microlight indicate that the microlight impact was Start of video Video at ~1 minute Video at ~2 minutes 40 seconds Video at ~3 minutes 6 seconds to final frame 5 seconds to final frame Figure 3 Snapshots extracted from the recovered video, showing the approach to Ben More A secondary radar track from a helicopter in the area at the time was also reviewed. The helicopter flew from the south-east, between Ben More and the adjacent peak, below the height of the peaks, and then to the north-west (Figure 4). Photographs and video taken from this helicopter at about the time of the accident were reviewed but did not capture the accident aircraft. Figure 3 Snapshots extracted from the recovered video, showing the approach to Ben More 36 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-CWIK EW/C2012/05/02 between 0.4 nm and 1.1 nm ahead of the helicopter. This established that the helicopter was not a factor in the accident. Meteorological information On the day of the accident a large high pressure system was established to the west of the UK, extending its influence over Scotland. Over Scotland, the surface wind observations valid at 1100 hrs UTC show westerly winds of 10-15 kt with a 2,000 ft gradient wind of 310° at 25-28 kt. At Glen Ogle, near the crash site, the surface wind between 1000 and 1200 UTC was westerly 16-19 kt with gusts of 24-26 kt. Figure 4 Flight paths and timings of the microlight and helicopter The movement of the cloud shadows near the summit of Ben More was recorded on the video and analysis indicated a wind of 306°T at 32 kt, at about 4,000 ft. The visibility was approximately 40 km with the generally scattered cloudbase between 3,500 and 5,000 ft. The sea level temperature was about 12°C. Medical and pathological information A post-mortem examination of both occupants revealed that they had died of severe multiple injuries, consistent with having been caused when the aircraft struck the ground. The crash forces were outside the range of human tolerance and therefore the impact was not 37 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-CWIK EW/C2012/05/02 survivable. There was no evidence of any pre-existing condition that may have contributed to the accident and toxicology showed no evidence of drugs or alcohol in either occupant. Mountain flying guidance There are a number of documents available on the internet covering mountain flying. An example is the Civil Aviation Authority of New Zealand ‘Good Aviation Practice (GAP), Mountain Flying’ publication (www.caa.govt.nz/safety_info/good_aviation_practice. htm). It contains valuable information and clearly describes the potential hazards associated with flying in mountainous terrain. The illustrations below are reproduced from this document. Wind strength increases as it passes over a mountain feature, due to the Venturi effect of the mountain. As a result, wind strength on the summit of a mountain will be significantly greater than the ambient wind speed away from the summit at the same height. This fact is illustrated in Figure 5, with wind speeds for illustrative purposes only. The strength of the ambient wind will govern the degree of turbulence created. A gentle wind will simply flow over the terrain following the contours but as the strength increases the wind will curl over and around features, forming up and down drafts as well as vortices, the severity of which will increase with the strength of the wind. This effect is shown in Figure 6. Figure 5 The Venturi effect of the mountain increasing the wind speed at the summit Figure 6 The creation of hazardous turbulence in the lee of high ground related to wind strength 38 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-CWIK EW/C2012/05/02 Analysis The wreckage examination did not reveal any evidence of a technical fault or pre-impact structural failure. The engine was not examined in detail as audio evidence from the video camera, and the damage to the propeller blades, indicated that the engine was producing power at impact. The video retrieved from the GoPro camera recorded the flight path as stable up to a point about 300 m from the summit of Ben More. At this point the aircraft started to roll rapidly from left to right and pitched nose-down. The increase in engine power up to the redline speed of 5,800 rpm and then, in the last 10 seconds before the end of the recording, to approximately 6,090 rpm, suggests the pilot was trying to arrest his rate of descent and climb out of the turbulence. The aircraft’s motion and final flight path is consistent with the effect of turbulent air in the lee of a summit, which creates downdrafts, rotors and vortices. The evidence of the hill walker on the summit of Ben More, regarding the direction and strength of the wind, indicated that the aircraft’s track was downwind of the summit with a wind speed of 30 to 35 kt. This is supported by the recorded video data showing the clouds indicating a wind of 306° at 32 kt near the summit. The pilot of G-CWIK would have known that the winds were westerly from his takeoff at Perth but it is not known how he was conducting his en route navigation and whether that would have given him an appreciation of the wind speed and direction at Ben More. Further, the video recording shows that there was no compelling visual evidence of the wind speed and direction at the summit, such as snow ‘spindrift’. It is likely that, in this case, a lack of awareness of the wind conditions, and of the likelihood and severity of turbulence downwind of high ground, were factors in this accident. In summary, the severity of the turbulence created by the wind, close to the summit of Ben More, was such that it exceeded the safe conditions for flight in the microlight aircraft. This resulted in a loss of control, which led to the impact close to the summit of the mountain. Safety Recommendation The UK CAA produces a series of Safety Sense Leaflets covering a wide range of aviation activities but this does not currently include a leaflet covering mountain flying. The following Safety Recommendation is made: Safety Recommendation 2012-037 It is recommended that the Civil Aviation Authority produce a Safety Sense Leaflet, or other guidance material, covering the activity of mountain flying for the UK general aviation community. 39 ©Crown copyright 2013 AAIB Bulletin: 1/2013 AAIB correspondence reports These are reports on accidents and incidents which were not subject to a Field Investigation. They are wholly, or largely, based on information provided by the aircraft commander in an Aircraft Accident Report Form (AARF) and in some cases additional information from other sources. The accuracy of the information provided cannot be assured. 41 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-AIYR EW/G2012/09/04 ACCIDENT Aircraft Type and Registration: DH89A Rapide Dragon, G-AIYR No & Type of Engines: 2 De Havilland Gipsy Queen 3 piston engines Year of Manufacture: 1943 (Serial no: 6676) Date & Time (UTC): 8 September 2012 at 1136 hrs Location: Duxford Aerodrome, Cambridgeshire Type of Flight: Commercial Air Transport (Passenger) Persons on Board: Crew - 1 Passengers - 7 Injuries: Crew - None Passengers - None Nature of Damage: Fire damage to the outer section of the right stub wing and inner section of the lower right mainplane Commander’s Licence: Airline Transport Pilot’s Licence Commander’s Age: 55 years Commander’s Flying Experience: 4,134 hours (of which 40 were on type) Last 90 days - 9 hours Last 28 days - 4 hours Information Source: Aircraft Accident Report Form submitted by the pilot and investigation report by the aircraft operator The aircraft was being prepared for flight, with the commander and seven passengers on board. The left engine was started first, followed, after priming, by the right engine. Immediately after the right engine started the pilot saw a flame, which appeared to originate from the outboard section of its engine cowling. Believing the right wing to have caught fire, he shut down both engines. The pilot then assisted with passenger evacuation, which had been initiated by the ground crew when the fire broke out. Ground crew also tackled the fire, with hand held appliances, and had extinguished the fire when the aerodrome fire service arrived on scene. An investigation conducted by the aircraft operator concluded that the probable cause of the fire was over-priming of the hot engine, leading to the ignition of overflowed fuel. 42 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-BKJW EW/G2012/09/23 ACCIDENT Aircraft Type and Registration: Piper PA-23-250 Aztec, G-BKJW No & Type of Engines: 2 Lycoming IO-540-C4B5 piston engines Year of Manufacture: 1971 (Serial no: 27-4716) Date & Time (UTC): 28 September 2012 at 1204 hrs Location: Southend Airport, Essex Type of Flight: Private Persons on Board: Crew - 1 Passengers - 4 Injuries: Crew - None Passengers - None Nature of Damage: Damage to propellers and nose section, engines shock- loaded Commander’s Licence: Private Pilot’s Licence Commander’s Age: 64 years Commander’s Flying Experience: 1,536 hours (of which 1,337 were on type) Last 90 days - 21 hours Last 28 days - 1 hour Information Source: Aircraft Accident Report Form submitted by the pilot The pilot was conducting a normal landing on Runway 24 at Southend Airport following a local flight. The weather was fine, with a surface wind from 240° at 17 kt. The landing gear was selected down and three green ‘down’ indicator lights obtained, together with visual confirmation that the nose gear was extended. The landing was normal until near the end of the landing roll when wheel braking was applied. The gear warning horn sounded and, two or three seconds afterwards as the aircraft was travelling at about 15 kt, the nose landing gear gently moved towards the retracted position. The aircraft’s nose lowered to the ground and both propellers contacted the runway. The aircraft came to a stop within about 20 m and the pilot secured the aircraft. The five occupants were uninjured; they vacated the aircraft through the cabin door as the airfield emergency services arrived. The aircraft was checked by a local maintenance and repair organisation, which found no faults with the undercarriage system. Several retraction/extension cycles were carried out, on each occasion achieving positive downlock and green indicator lights for all three undercarriage legs. The cause of the accident was thus unresolved. 43 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-SACY EW/G2012/09/14 ACCIDENT Aircraft Type and Registration: Aero AT-3 R100, G-SACY No & Type of Engines: 1 Rotax 912-S2 piston engine Year of Manufacture: 2007 (Serial no: AT3-029) Date & Time (UTC): 5 September 2012 at 1050 hrs Location: Fishburn Airfield, Co Durham Type of Flight: Private Persons on Board: Crew - 1 Passengers - 1 Injuries: Crew - None Passengers - None Nature of Damage: Damage to nose landing gear and propeller Commander’s Licence: National Private Pilot’s Licence Commander’s Age: 62 years Commander’s Flying Experience: 114 hours (of which 36 were on type) Last 90 days - 4 hours Last 28 days - 1 hour Information Source: Aircraft Accident Report Form submitted by the pilot After an uneventful flight from the aircraft’s base at Sherburn in Elmet, the pilot prepared for a landing on Runway 26 at Fishburn. The weather was fine, with a surface wind from 330° at 8 to 10 kt. The pilot discontinued his first approach because the aircraft was too high on final. He was content with the second approach, despite the crosswind presenting some difficulty, but as he was about to reduce power and flare for landing, the stall warning activated, which he recalled thinking had sounded unusually early. He thought he may have lowered the nose briefly in response, but the aircraft appeared to stall before striking the runway in a nose-low attitude. The nosewheel broke off and the aircraft came to rest on its main wheels and propeller spinner. The pilot and his passenger were uninjured and both vacated the aircraft to await assistance from airfield personnel. The pilot thought that the aircraft may have stalled as it flew into the lee of hangars and trees, impacting the runway before he had time to take recovery action. 44 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-ATBI EW/G2012/09/25 ACCIDENT Aircraft Type and Registration: Beech A23 Musketeer II, G-ATBI No & Type of Engines: 1 Continental Motors IO-346-A piston engine Year of Manufacture: 1965 (Serial no: M-696) Date & Time (UTC): 5 September 2012 at 0830 hrs Location: Fishburn Airfield, Co Durham Type of Flight: Private Persons on Board: Crew - 1 Passengers - 1 Injuries: Crew - None Passengers - None Nature of Damage: Damage to right wingtip and aileron, stabilator and fuselage Commander’s Licence: Private Pilot’s Licence Commander’s Age: 62 years Commander’s Flying Experience: 1,150 hours (of which 500 were on type) Last 90 days - 10 hours Last 28 days - 2 hours Information Source: Aircraft Accident Report Form submitted by the pilot The pilot was taxiing the aircraft toward the threshold of grass Runway 26, prior to an into-wind takeoff. The runway had several distinct gradients and the ground sloped downwards as the aircraft backtracked toward the start of the runway. The grass was wet and the surface wind was from 330° at 12 kt. The aircraft gained excess speed as it taxied downhill and, as the pilot attempted to turn the aircraft to line up on the runway, it skidded to the right and into a hedge, which arrested its motion. The pilot considered that excess taxi speed and his lack of appreciation of how much the braking effect was reduced on the wet grass had caused the accident. He also observed that the grass surface may have been wetter towards the bottom of the slope, where the attempt to turn the aircraft was made. 45 ©Crown copyright 2013 AAIB Bulletin: 1/2013 N35SN EW/G2012/09/16 ACCIDENT Aircraft Type and Registration: Beechcraft 33 Debonair, N35SN No & Type of Engines: 1 Continental IO-470-J piston engine Year of Manufacture: 1960 Date & Time (UTC): 22 September 2012 at 0930 hrs Location: Perranporth Airfield, Cornwall Type of Flight: Private Persons on Board: Crew - 1 Passengers - None Injuries: Crew - None Passengers - N/A Nature of Damage: Damage to the nose landing gear, lower forward fuselage and propeller Commander’s Licence: Private Pilot’s Licence Commander’s Age: 49 years Commander’s Flying Experience: 463 hours (of which 203 were on type) Last 90 days - 9 hours Last 28 days - 5 hours Information Source: Aircraft Accident Report Form submitted by the pilot The aircraft was landing at Perranporth after a flight from Old Sarum. The weather was fine, with a 15 kt easterly wind. The pilot made a normal approach to Runway 09, with both a green light and a mechanical indicator confirming that the landing gear was locked down. As the pilot lowered the nose after touchdown, he heard and felt an impact. He immediately pulled back on the control column to keep the nose off the runway, while steering with rudder and brakes. Eventually, the nose dropped fully to the ground and the aircraft slid to a halt on the runway, about 45° offset from the centreline. The pilot, who was uninjured, secured the aircraft and vacated it. This required going forward from the cabin door and over the wing, as the rear step was too high off the ground. The pilot reported that the nose landing gear had suffered a mechanical failure. A knowledgeable witness in the control tower had reportedly observed the landing gear to be down prior to landing, and saw the nose gear collapse after a short ground roll. The reason for the failure of the nose landing gear has yet to be established. 46 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-BSZI EW/G2012/09/19 ACCIDENT Aircraft Type and Registration: Cessna 152, G-BSZI No & Type of Engines: 1 Lycoming O-235-N2C piston engine Year of Manufacture: 1984 (Serial no: 152-85856) Date & Time (UTC): 22 September 2012 at 1230 hrs Location: Carrickmore Airfield, Co Tyrone Type of Flight: Private Persons on Board: Crew - 1 Passengers - 1 Injuries: Crew - None Passengers - None Nature of Damage: Damage to propeller, engine mounting and bulkhead, left wing and nosewheel Commander’s Licence: Private Pilot’s Licence Commander’s Age: 50 years Commander’s Flying Experience: 161 hours (of which 69 were on type) Last 90 days - 13 hours Last 28 days - 8 hours Information Source: Aircraft Accident Report Form submitted by the pilot and further enquiries Synopsis The pilot decided to abort the landing soon after touchdown. He selected full power and wing flaps 0° for the climb away. However, the aircraft did not climb as expected. With obstacles in the projected flight path, the pilot elected to land immediately. The aircraft landed in a ploughed field and turned over; neither occupant sustained injury. History of the flight The aircraft was being flown from its base at City of Derry Airport to Carrickmore Airfield near Omagh in Northern Ireland. The runway at Carrickmore was hard-surfaced and 505 m (1,656 ft) long; it occupied an elevated position and was orientated 08/26. The weather at Carrickmore was fine, with a surface wind from 180° at 5 kt and a temperature of 13°C. The pilot flew a normal approach profile to Runway 08, configured with 20° flaps and flown at about 70 kt. He reported that he was satisfied with the approach parameters. The aircraft touched down about one third of the distance along the runway, and one wheel briefly left the narrow paved surface. The pilot corrected the deviation, but was generally dissatisfied with the landing. He therefore elected to abort the landing and fly a further approach. He applied full power and selected flaps 0°, rotating at 47 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-BSZI EW/G2012/09/19 a speed he believed to have been around 60 kt. It soon became apparent to the pilot that the aircraft was hardly climbing and would not clear a building and nearby power lines which lay on rising ground ahead and slightly to the left of the extended centreline (the ground fell away after the runway end before rising again). He turned the aircraft to the right but saw another building, also on elevated ground, ahead. He decided that the best course of action was to land the aircraft in the field immediately below, and warned his passenger. The field had been ploughed, and the aircraft’s nosewheel dug into the ground on landing, causing the aircraft to pitch nose-down and invert. Neither occupant was injured in the accident. The pilot recalled hearing the stall warning horn after he warned his passenger, but thought that he maintained control of the aircraft and it had not actually stalled. The pilot thought that the aircraft may have experienced carburettor icing, and that he may have overlooked the need to apply carburettor heat before landing, leading to reduced engine power following the aborted landing. However, apart from the poor climb performance, there were no other unusual indications, such as engine noises or rough running. Although the flap control was found in the 0° position, one wing flap was found at 10° and the pilot could not be sure that the flaps had fully retracted. The pilot thought the aircraft had reached a maximum height of no more than 100 ft. Aircraft performance Based on mass and balance figures provided by the pilot, the aircraft was only some 5 kg below its maximum allowable weight at the time of the accident, which would have placed it above the maximum weight at takeoff, 45 minutes earlier. The pilot had been aware that weight was an issue prior to flight, and fuel had been offloaded earlier in the day when his original plans had changed and it was decided that a passenger would accompany him on the flight. He estimated that the aircraft had departed City of Derry with about 11 or 12 US gallons (about 45% of maximum fuel). The Pilot’s Operating Handbook (POH) for the Cessna 152 listed takeoff and landing performance figures for ground roll and distances to and from 50 ft. However, these were based on the most favourable situation, which was achieved using ‘short field’ techniques. For landing, this entailed using 30° flap and an airspeed of 54 kt. The recommended initial climb configuration and airspeed for a baulked landing was 20° flap and 54 kt. The landing ground roll using POH short field techniques should have been 477 ft, with a total distance from 50 ft of 1,203 ft, before any safety factors were applied. Discussion The aircraft was operating at close to its maximum weight on a narrow and relatively short runway. The pilot was not familiar with the airfield and it was considerably different from his home airfield which was an international airport. It would appear that the aircraft gained very little altitude after becoming airborne, with separation from the terrain arising partly through the downwards slope of the ground after the runway end. A reduction in available engine power through carburettor icing could not be ruled out, although there was no rough running, low rpm, or other symptoms (other than an apparent lack of power). Another possibility is that the aircraft may have been unable to climb as a result of a combination of weight, configuration and airspeed. The decision to abort the landing was made quickly, and the actions taken by the pilot were those appropriate to a touch-and-go landing, with which he was very familiar, including selection 48 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-BSZI EW/G2012/09/19 of flaps 0°. The lack of headwind and narrow runway could conceivably have provided misleading visual cues that the aircraft was travelling at greater airspeed than was the case. Consequently, it may have become airborne at too low an airspeed, possibly with flaps still retracting, compromising its initial climb performance. 49 ©Crown copyright 2013 AAIB Bulletin: 1/2013 G-BDOE EW/G2012/08/20 ACCIDENT Aircraft Type and Registration: Cessna FR172J Reims Rocket, G-BDOE No & Type of Engines: 1 Continental Motors Corp IO-360-J piston engine Year of Manufacture: 1975 (Serial no: 559) Date & Time (UTC): 20 August 2012 at 1515 hrs Location: Farm strip near Truro, Cornwall Type of Flight: Private Persons on Board: Crew - 1 Passengers - 1 Injuries: Crew - None Passengers - None Nature of Damage: Damage to landing gear, propeller, wing and tail Commander’s Licence: Private Pilot’s Licence Commander’s Age: 65 years Commander’s Flying Experience: 81 hours (of which 81 were on type) Last 90 days - 7 hours Last 28 days - 3 hours Information Source: Aircraft Accident Report Form submitted by the pilot The aircraft was landing back at a private farm strip, orientated 03/21, after a 55-minute flight. The weather was fine, with a south-westerly wind of 10 kt. The approach, which was being made into wind, was slightly higher than usual but otherwise seemed normal. Full flap was selected and power was increased. A small valley lies across the approach to the airstrip and, typically, causes an area of sink, which was anticipated. However, as the pilot opened the throttle, the engine did not respond. The aircraft was unable to clear a hedge in the undershoot, so the pilot raised the nose of the aircraft just prior to impact. The underneath of the aircraft struck the hedge before it came to rest on the gra