Select Committee on Chinook ZD 576 Report



Introduction of Chinook Mark 2 into RAF service

42.  At the time of the accident ZD 576 had flown 66.5 hours since receiving the mid-life update (MLU) by the makers, Boeing, in the USA, which turned it from a Mark 1 into a Mark 2. The most significant part of this MLU was the installation of a Full Authority Digital Engine Control (FADEC) system. This consists of a number of components of which the two major ones are a digital engine control unit (DECU) and a hydro-mechanical assembly for each of the two engines. The purpose of FADEC is by control of the fuel supply to maintain approximately 100 per cent rotor speed in all conditions and to match engine torque between the two engines.

43.  FADEC had been fitted in a number of RAF Chinooks over the preceding years and had given rise to certain problems. In the summer of 1993 an independent defence IT contractor, EDS-SCICON, was instructed to review the FADEC software; after examining only 18 per cent of the code they found 486 anomalies and stopped the review. In October 1993 the Aeroplane and Armament Experimental Establishment (A&AEE) at Boscombe Down advised the MoD that because of the unverifiable nature of the FADEC software it could not recommend Controller Aircraft Release for the Chinook Mk 2. Both EDS-SCICON and A&AEE recommended that the FADEC software be rewritten but this was not done, and in November 1993 Chinook Mk 2s were released into operational service, subject to certain operational restrictions on the load which they could carry and the height at which they could fly thereby avoiding icing - restrictions which had not applied to the Chinook Mk 1. In addition intermittent engine failure captions were being regularly experienced by aircrew of Chinook Mk 2s and there were instances of uncommanded run up and run down of the engines and undemanded flight control movements (UFCMs).[15]

44.  On 1 June 1994 flying tests of the Chinook Mk 2 were suspended at Boscombe Down. Squadron Leader David Morgan, at that time Flight Commander of the Chinook Operational Conversion Unit, thought that this suspension was "connected with icing trials" and not due to FADEC difficulties (Q 484; cf MoD letter 2 Nov 2001, p 65 of HL Paper 25(ii)). However, in comments by the MoD[16] on a paper submitted to them by Lord Chalfont, it is stated:

    "Boscombe Down's decision not to authorise further trials flying in June 1994 was made against a background of several engine control system malfunctions that had occurred on the ground during start up checks, which had not at that point been explained to Boscombe Down's satisfaction by the aircraft or engine Design Authority. The necessary clarification was completed and accepted by Boscombe Down on 24 October 1994. Test flying was recommenced without any changes to the aircraft FADEC system, or any additional operating limitations. Operating flying continued within the weight restrictions applied."

45.  The MoD continued in relation to a memo of 3 June 1994 from Boscombe Down:

    "The 3 June 94 memo was an internal MoD working level document, the last sentence of which reads: "Notwithstanding the claims made in Textron's white paper, the problem remains that the product has been shown to be unverifiable and is therefore unsuitable for its purpose".

    But it is emphasised that this statement arose because Boscombe Down were unable to verify the software independently using their preferred analysis, which was neither mandated nor included in the development contract placed in 1985. Contractors had carried out their own validation and the fact that Boscombe Down could not verify the software using their preferred software should not be taken to imply that there was an inherent problem with its design."

46.  A statement on behalf of the MoD to the FAI[17] included the following:

    "As to the allegation that some pilots refused to fly the Chinook HC Mk 2 during CA Release trials at Boscombe Down, this is an over simplification of what actually happened and perhaps it would be helpful if some of the background was explained. On 7 March 1994 during one of the specified FADEC checks on the ground, the engine of an HC Mk 2 flamed out. Trials at Boscombe Down were halted while the failure was investigated. The failure was not due to a software fault and flying resumed on 20 April. However in the period up to 2 June 1994 there were a number of incidents involving airborne HC Mk 2 of which approximately 5 were due to FADEC malfunction whilst operating in normal mode. There had also been other incidents on the ground. The MoD(PE) Project Office sought explanations of the various incidents from the aircraft and engine manufacturers but in the absence of satisfactory explanations Boscombe Down suspended trials flying."

47.  It is clear from these quotations that at the time of the crash there were still unresolved problems in relation to the FADEC system of Chinook Mk 2s.

Circumstances of the accident

48.  The aircraft took off from Aldergrove at 17.42 hours to fly to Fort George, near Inverness. The aircrew were members of the RAF Special Forces Flight and were all very experienced. The pilots were Flight Lieutenants Tapper and Cook of whom the former was captain and probably non-handling pilot. In the cabin were Master Air Loadmaster (MALM) Forbes and Sergeant Hardie. The former was considered to be an exceptionally professional crewman and entirely competent to check the navigation of the aircraft; indeed he had been observed on the preceding evening checking maps for this purpose. A four-man Special Forces crew work together as a unit more than in conventional flying, and it was the practice for one of the crewmen to act as second navigator.[18] After obtaining weather reports Flight Lieutenant Tapper decided that the flight to Fort George would be a low-level sortie flying under visual flight rules (VFR) and intimated that he was not intending to fly under instrument flight rules (IFR). Indeed flying under IFR in the vicinity of the mountains to the east of Ballachulish and Fort William would not have been permitted by the icing restrictions.[19] He therefore selected a route which included flying to the vicinity of the Mull of Kintyre lighthouse, which was programmed into the SuperTANS navigation system (see para 53 below) as way point (WP) A, and from there to Corran on Loch Linnhe, programmed as way point B.

49.  The aircraft was seen at Carnlough coasting out low over the Antrim coast towards the Mull of Kintyre. Shortly before 18.00 hours a yachtsman, Mr Mark Holbrook, who was some two miles south west of the lighthouse, saw the aircraft for some five seconds about 300 yards away. He estimated that its height was 200-400 feet and speed 60-80 knots, but he had never seen a Chinook before and was at the time involved in the difficult operation of manoeuvring his vessel among a considerable number of trawlers which were fishing.

50.  Mr Holbrook was the last person to see the helicopter. It was however heard immediately before and as it crashed by a number of witnesses at, and in the vicinity of, the lighthouse. The initial point of impact was 810 feet above mean sea level and about 500 metres east of the lighthouse, but the bulk of the aircraft remained airborne for a further 187 metres horizontally north and 90 feet vertically before coming to rest in pieces. Fire broke out immediately. All those on board sustained injuries from which they must have died almost instantaneously. The points of impact were shrouded in local cloud with visibility reduced to a few metres, which prevented those witnesses who had heard the aircraft from seeing it.

Examination of wreckage by the Air Accidents Investigation Branch

51.  On 3 June 1994 Mr Tony Cable, a Senior Inspector of the AAIB, flew to the crash site and remained closely involved thereafter in the investigation until he signed on 5 January 1995 the statement to the Board which he had prepared. This detailed report, which ran to some 66 pages and is published in HL Paper 25(i), resulted not only from Mr Cable's own work at the site and at AAIB Farnborough but also from opinions expressed by the relevant manufacturers to whom salvaged components or information had been provided.

52.  The aircraft was very severely damaged: 80 per cent of the fuselage structure was appreciably damaged by ground fire, of which around 20 per cent was destroyed (AAIB statement para 7.1). It had neither cockpit voice recorder nor accident data recorder. This presented a formidable task to Mr Cable who expressed the view to us that "the evidence was remarkably thin" (QQ 956, 968, 1013). From the final position of the aircraft, which had by then extensively broken up, he estimated that at the initial impact the flight path was 20 degrees up relative to the horizontal, the pitch angle was 30 degrees nose up with a 5-10 degrees left roll and the ground speed was approximately 147 knots, as disclosed by detailed examination of the cockpit ground speed and drift indicator (AAIB statement para 6). The left rudder pedal appeared to have been applied 77 per cent (AAIB statement para 7.4.3, 7.4.9).

53.  The aircraft was fitted with a Racal Avionics "SuperTANS" Tactical Area Navigation System providing navigation information from two independent sources. The system enables a number of way points to be fed into it before a flight. When flying from the point of departure to the first way point the screen[20] shows bearing, distance and "time to go" from the aircraft's current position to the way point. When the pilot alters the system from the first way point to the second, the distance and bearing of the former are replaced on the screen by those of the latter and so on as way points are progressively changed. Racal confirmed that the system was performing perfectly at the time of loss of power and extracted from its memory the information that way point B had been selected when way point A was 0.81 nautical miles distant, bearing 018°. The distance from the way point change to the point of impact was 0.95 nautical miles. The system gave no information as to height or time at the way point change but had recorded that at approximately 15-18 seconds before power down the aircraft was at a height of 468 feet ± 50 feet (Board report para 49). The manufacturers have told us that "18 seconds is likely to be a better estimate".[21] The TANS had also recorded that the height above sea level at impact was 665 ft[22], whereas in fact it was 810 ft. The investigating board noted this discrepancy (para 49); they considered it probably due to "the mechanics of the crash and the developing fireball", but we know of no evidence to support this. The TANS is not intended to act as a Flight Recorder or what is colloquially known as a Black Box, and the information referred to above was achieved by a somewhat complicated and ingenious method of extraction employed by Racal.

54.  The AAIB considered the engines and controls and because of the reported FADEC service difficulties investigated the DECUs in detail. DECU no. 2 remained partially functionable with deficiencies consistent with impact damage, and with no faults or exceedances traced in its memory of the last flight. DECU no. 1 had suffered gross fire damage with part of its casing melted away and severe damage to the interior components whereby its memories of exceedance and fault listing had been destroyed.

55.  A detailed examination of the flight control mechanical linkages was carried out and the AAIB stated, "No evidence to suggest a control jam was found, although such a possibility could not be excluded, given the level of system damage" (para 7.4.2).

56.  Important parts of the hydraulic flight control systems were housed in a small closet, colloquially known as the "broom cupboard", at the rear of the cockpit. There were two control pallets containing respectively 23 and 26 threaded inserts for component attachment. On 10 May 1994 the thrust balance spring attachment bracket on the aircraft's thrust/yaw control pallet had detached; this was due to the somewhat inadequate method of attaching the inserts to the pallet. This detachment had resulted in an undemanded flight control movement (UFCM) in the collective system.[23] An engineering report of the following day relating to this incident stated among other things,

    "Detachment of the bracket within the flying control closet during flight could present a serious flight safety hazard, with the danger of a detached bracket fouling adjacent flying controls".[24]

57.  After the accident the investigators found that both inserts for the thrust balance spring attachment bracket had detached as well as most of the other inserts to both pallets. The AAIB stated, "as an insert could apparently pull out of the pallet without appreciable distress to the components necessarily resulting, the possibility that insert(s) had detached prior to the accident could not be dismissed" (para 7.4.2). In the Flight Control Summary the AAIB reiterated that "the possibility of control system jam could not be positively dismissed" and further stated that "little evidence was available to eliminate the possibility of pre-impact detachment of any of the pallet components" (para 7.4.9).

58.  The investigators also found a considerable quantity of very small metallic particles and four fine metal slivers in the hydraulic system of the lower control actuators which form part of the flight control system. They concluded that this contamination had occurred pre-impact, but that it had not contributed to the accident (AAIB statement para 7.4.4).

59.  With a view to ascertaining the manoeuvre necessary to produce the initial impact conditions the airframe manufacturers, Boeing, at the request of the Board produced mathematical models simulating the aircraft's final behaviour from a postulated range of initial steady flight conditions. In order to do this they were provided by Mr Cable with what he considered to be the pitch attitude and flight path angle of the aircraft at initial impact together with the extensions found in two components of the flight control system, namely, the differential airspeed hold (DASH) actuators and the longitudinal cyclic trim actuators (LCTAs), and other pieces of information derived from the site (Q 950 and p 86 of HL Paper 25(ii)).

60.  The Boeing simulation tried a series of input conditions suggested by AAIB, but found that close simultaneous matching of the predicted conditions with the control criteria was possible in only a few cases. A ready match was found where initial conditions combined an airspeed of 150 knots[25] with a rate of climb (ROC) of 1000 ft/min (AAIB statement para 8). In order to achieve a maximum ROC, which far exceeds 1000 ft/min, airspeed requires to be reduced to 80 knots or below. Boeing then estimated that, some four seconds before the initial impact, input was applied to the controls to achieve a "cyclic flare", increasing the ROC and decreasing forward speed by bringing the nose up.

61.  At initial impact Boeing's simulation produced the following result (AAIB statement para 8):

Airspeed -135 kt
Normal Acceleration -2.2 g
Rotor Speed -204 rpm (91%)
DASH Extension -23%
LCTA Extension -Virtually Fully Extended
Aircraft Pitch Attitude -31° Nose Up
Aircraft Roll Attitude -5° Left
Aircraft Yaw Attitude -1° Left
Flight Path Angle -20° above the horizontal
Climb Rate -4670 ft/min
Horizontal Distance Travelled -822 ft (250 m)
Vertical Distance Travelled -128 ft
Groundspeed -158 kt

15   See Board of Inquiry report part 2 para 35(d). Back

16   31 July 2001. Unpublished. Back

17   See MoD letter 9 January 2002, p 86 of HL Paper 25(ii). Back

18   Sheriff's determination pp 102-5 of HL Paper 25(ii); Q 789. Back

19   Instrument flight must be conducted at or above a height ("safety altitude") 1000 ft above each of the high obstacles on the intended route (Q 280). Flt Lt Tapper had calculated the safety altitudes for the latter part of the sortie as 5900 ft Mull-Corran and 5800 ft Corran-Inverness. The CA Release precluded flight in an indicated temperature below +4° C. A weather aftercast shows that this temperature would have been reached at 5000 ft; and the Sheriff calculated that the crew may well have expected to reach it much lower, at 2500 ft (determination p 120 of HL Paper 25(ii)). Back

20   Illustrated in slide 23 on p 134 of HL Paper 25(i). Back

21   Note from Thales Avionics, 6 December 2001, unpublished for reasons of commercial confidentiality. Back

22   Report by Racal Avionics (now Thales Avionics), Annex Y to the Board of Inquiry report. A restricted document, released to us in confidence by kind permission of Thales Avionics. Back

23   According to the investigating board's report, para 35(b)(3). In evidence to us, however, Gp Capt Pulford said that the effect would be to change "the feel of the controls" (Q 38). Back

24   DHS follow-up report of serious occurrence or fault, signed by Gp Capt A M Verdon, HS31. Back

25   With a tailwind component of 24 knots, this would give a groundspeed of 174 knots. Back

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