Select Committee on Chinook ZD 576 Report


CHINOOK ZD 576

PART 5: EVIDENCE BEFORE THIS COMMITTEE

Witnesses

88.  In addition to Mr Holbrook we heard the evidence of:

89.  We also received a considerable number of documents from the MoD and many others which were intended to assist or influence us in our deliberations. We are grateful to all these persons for their assistance; they are listed in Appendix 2.

90.  The Secretary of State for Defence, the Rt Hon Geoffrey Hoon MP, kindly offered to give evidence himself to us on the position of the MoD[27]. We decided not to trouble him, since he was not in office at the time of the crash and the Board of Inquiry, and since the question before us concerned a decision reached by others at that time rather than any aspect of current policy. We record our appreciation however for the full and helpful co-operation of staff at all levels in the Royal Air Force and the Air Secretariat of the Ministry of Defence. At a time when other serious matters demanded their attention, they responded promptly to our many queries and requests. In particular they arranged for us to fly to the Mull in a Chinook on 25 September, and for one of our number to inspect a partially dismantled Chinook at RAF Odiham on 22 October.

91.  What follows is not an exhaustive summary of the evidence, but a considered analysis of those elements on which we base our conclusions.

Written evidence of Mr Malcolm Perks

92.  In addition to the detachment of the balance spring on 10 May 1994 the aircraft had experienced some other problems in the weeks prior to the accident. On 21 April 1994 a torque mismatch occurred which lasted for about 1 second before returning to normal. No fault codes were indicated on the DECUs. On 17 May 1994 number one engine power caption came on and engine temperature reached 950 degrees. The engine was thereafter rejected due to overheating but on inspection nothing suspect was found. On 26 May 1994 number two engine failed caption came on spuriously and went out after 10 seconds.

93.  In 1989 a Chinook HC2 engine owned by the MoD was destroyed on test at Wilmington, Delaware, USA as a result of a runaway[28]. The MoD initiated claims against Textron-Lycoming the manufacturers of the engine and Boeing who were responsible for the test. Boeing settled but the case against Textron-Lycoming went to arbitration later and resulted in a substantial award in favour of the MoD. The claim against Boeing rested solely on negligence in carrying out the test whereas that against Textron-Lycoming proceeded on the basis that they had failed to exercise due care in the design, development and test of the FADEC system.

94.  Mr Malcolm Perks, who now lives in Canada and who had spent many years working in the field of FADEC, provided technical evidence to the arbitration on behalf of the MoD. He was invited to assist us and accepted the invitation by furnishing two memoranda (p 87 of HL Paper 25(ii)). Although it has been repeatedly maintained on behalf of the MoD that the problem at Wilmington was due entirely to negligent testing Mr Perks suggested to us that if that had been the case it would have been unlikely that, Boeing having settled, Textron-Lycoming would also have been found negligent in respect of testing alone without any liability for design or development. Mr Perks had not seen the reasons for the finding of the arbitration board and nor have we.

95.  The FADEC system incorporates fault codes in its memory so that maintenance staff can see what has happened and correct the fault. The fault condition implied by the E5 fault code was heavily implicated in the incident at Wilmington and the same fault code was found in the surviving DECU after the crash of ZD 576 (AAIB statement para 7.3.2.4). There has been much speculation that the coincidence of the E5 faults pointed to a runaway engine in ZD 576 as at Wilmington.

96.  Mr Perks in his memoranda explained to us that the Wilmington incident was not caused by an E5 fault alone but by its conjunction with another fault and that by 1994 because of the action of the system designers an E5 fault was being dismissed as a nuisance fault of no significance. Furthermore Mr Cable explained that the DECU had two portions to the memory of faults namely (i) retained faults since its last overhaul and (ii) faults since the last engine start. The E5 fault in ZD 576's DECU was found in the former historical portion and not in that of the last flight. In all these circumstances we are satisfied that an E5 fault had no relevance to the accident.

97.  The FADEC system used in Chinook Mk 2s has a built-in protection to prevent destructive overspeed resulting from a runaway engine. Mr Perks explained that, even if the runaway was contained within the engine's speed limits, it could cause major controllability issues due to rapid acceleration of the rotor system.

Evidence of Mr Cable, AAIB

98.  As already mentioned, Mr Cable in evidence to us stressed that throughout the investigation the evidence was "remarkably thin" (QQ 956, 968, 1013). While the evidence available to him pointed strongly to the engines operating normally, i.e. without distress, at the point of initial impact, he conceded that this did not necessarily mean that this was in accordance with pilot commands (QQ 181-4). He further explained that the possibility of an intermittent fault prior to impact could not be dismissed (Q 182).

99.  He further explained that the detachment of the pallet inserts and the components carried by them could possibly cause a restriction or jam. "It would be very difficult - impossible - to dismiss the possibility that there had been a restriction and evidence had not been found" (Q 196). This explanation is readily understandable given the crowded equipment in the broom cupboard. A balance spring is some 6 inches long by 1½ inches in diameter and its mounting bracket about 1½ inches long.

100.  The only positive evidence of a fault possibly contributing to the accident was a radar altimeter system fault (AAIB statement para 7.2.17 and conclusions 48-9 and 52). However, in the light of all the evidence before us, we do not consider that this fault is likely to have been relevant.

101.  Mr Cable summed the situation up thus:

102.  As already referred to in paragraph 58, the AAIB investigation disclosed a considerable quantity of very small metallic particles in residual hydraulic fluid in parts of the boost actuator for both the pitch integrated lower control actuator (ILCA) and the thrust lower control actuators (LCAs) together with the presence of four fine metal slivers up to 0.2 inches long on one of the servo valve screens of the yaw ILCA boost actuator. This contamination was thought to have been present prior to the accident (AAIB statement para 7.4.4).

103.  In evidence Mr Cable expressed the opinion that a failure of both lower control actuator systems due to hydraulic contamination would be unlikely to be a major problem as it would merely reduce the boost on the pilot's control to the upper boost actuators which drive the rotor blades[29]. A jam of an upper boost actuator would be a very different matter (QQ 204-10).

104.  The US Army, however, who operate very large numbers of Chinooks, take a different view. In a report of June 1997[30] on an incident when a Chinook turned upside down at about 1100 feet and righted itself at about 250 and where no exact cause could be established, hydraulic contamination was considered to be a possible cause. The recommendations section of the report[31] referred to "uncommanded oscillations, flight control movements, and flight attitude changes" possibly related to the performance of the upper boost actuators and metal contamination in part thereof.

105.  The recommendations continued,

    "An additional critical area is the integrated lower control actuators (ILCA). The metal contamination and moisture found in the pitch, roll and yaw ILCAs are considered critical to FLIGHT SAFETY. The amount of contamination found in the pitch and roll ILCA were considered sufficient to cause a disturbance in the normal operation of these components at any time. One solution may be to establish a drain point for each system 2 ILCA, since the corrosion and moisture contamination appears to be primarily found in system 2. CCAD shop personnel reported that some ILCAs arrive with secondary valves jammed due to internal corrosion. This means the unit is operating on the primary control valve with no back-up or secondary valve available. If the primary valve jams, in this situation, the capability to direct hydraulic fluid flow ceases.

    The upper boost actuators and ILCAs deserve immediate and positive action, since these two areas are CRITICAL TO FLIGHT SAFETY, PERSONNEL SAFETY, AND EQUIPMENT SAFETY." [32]

Evidence of Squadron Leader David Morgan

106.  Squadron Leader Morgan (QQ 473-86) referred to the over-speed check in the FADEC system which prevented an engine running up and thereby damaging or losing the rotor system. Boscombe Down had insisted that this mechanism be checked before each flight because of "a number of unresolved issues" with FADEC. During such checks engines overheated and sometimes ran up (cf Burke Q 677). He had no personal experience of a run up nor of any flight-critical malfunctions but was aware of spurious engine failure captions in the control instruments which were capable of providing a distraction of up to 10 seconds depending upon circumstances. He explained that an engine run up which increased rotor RPM would increase vibration levels and render more difficult the reading of instruments (QQ 546-7).

Evidence of Squadron Leader Robert Burke

107.  Squadron Leader Burke had extensive experience in flying helicopters including Chinooks Mks 1 and 2 and was described by his unit commander in April 1993 as having air-testing skills on the Puma and Chinook which were unique. He was able to provide us with useful information about the problems which he had experienced when testing Chinooks. At the outset of the investigation into the accident he was contacted by Mr Cable and had two or three telephone discussions with him in relation to control positions (QQ 658, 662). Thereafter he had nothing further to do with the Board of Inquiry.

108.  After Squadron Leader Burke gave evidence, Group Captain Pulford submitted a statement to us (p 68 of HL Paper 25(ii)) in which he sought to explain why Squadron Leader Burke had not been asked to give evidence to the investigating board. He stated that as the Chinook maintenance test pilot "his flying was conducted in accordance with limited and pre-determined flight test schedules and he therefore lacked the operational currency to provide relevant evidence to the inquiry". This reasoning seems to assume that problems which Squadron Leader Burke might have encountered on test would not or could not occur in operational flying - an assumption whose justification we feel to be in doubt.

109.  Squadron Leader Burke spoke to having experienced two engine run ups on the ground at the Boeing factory in Philadelphia while flying with an American Army test pilot (Q 655) and similar run ups when testing the overspeed limiter on the ground at Odiham (Q 680). He also spoke to problems with the multi-point connectors which went from the engines into the DECU. These were of bad design and liable to be displaced by vibration which then produced a power interruption. Although there was a back-up system this did not always work and on two or three occasions pilots had lost control of the engine condition lever. As a result squadrons introduced a procedure whereby crewmen every quarter of an hour checked that the connections had not been displaced in flight (QQ 677-9).

110.  At the time of the accident DECUs still presented recurring problems. They were removed from the aircraft when something had gone wrong and returned to the makers who on many occasions could find no fault (QQ 698-9).

111.  In relation to possible jams Squadron Leader Burke explained that, due to the complexity of the Chinook control system, a jam caused by a loose article such as the balance spring in the broom cupboard in one of the three axes, pitch, yaw or roll, could lead to quite random results in all three axes sometimes and certainly in two of them. He had personal experience while lifting off from the ground of a jam in one axis affecting the other two (Q 935). He also referred to the problems of DASH runaways in Chinooks of both marks causing temporary loss of control of aircraft (Q 929).

112.  Finally, Squadron Leader Burke commented on the rudder input of 77 per cent left yaw found in the wreck of ZD 576:

113.  Mr Cable told us that, though it was possible that this rudder input was applied before impact, it was also possible that it was due to the force of the impact itself (Q 999).

Evidence of Witness A

114.  Witness A, who was a member of the Special Forces Flight with considerable experience of flying Chinooks operationally, had, at the time of the accident, experienced intermittent engine fail captions on a reasonably regular basis. He had subsequently experienced torque mismatches on an intermittent basis (Q 784). Pilots were instructed that if the failed captions remained on for more than 12 seconds they were to be treated as though something was wrong with the engine but if they stayed on for less than that time they could be ignored. When a caption came on in flight one of the crew was directed to check engine instrumentation and the engine itself (Q 786).

115.  Witness A also had personal experience of UFCMs in Chinook Mk 1s (QQ 792-6). In one case over a period of days an aircraft bounced vertically every time it was turned right. Repeated unsuccessful attempts were made to find the cause and the problem eventually disappeared of its own accord. In another case in daylight the lights came on to maximum intensity, dimmed to minimum and the hydraulic gauges cycled between zero and maximum. The pilot reported that the aircraft was becoming difficult to control and Witness A ordered him to land at the first available opportunity. The subsequent engineering investigation found no fault.

116.  Witness A, like Air Commodore Crawford, expressed cogent reasons for thinking that the crew of ZD 576 could see the land mass when they changed the way point and that this change was entirely consistent with a continued VFR flight to Corran on the new course (QQ 797-8). When asked whether he could think of any reason why having changed way point the aircraft should have continued on its existing course he replied, "That is the crux of the matter. I cannot think of any reason why the crew would have elected to do that unless they were not doing it of their own volition" (Q 802). He found it very difficult to accept that the crew were unaware of the proximity of the Mull (Q 804).

117.  In answer to a question as to how much the unforeseen malfunctions occurring in the Chinook Mk 2 since its introduction were a matter of discussion among helicopter pilots, he answered,

118.  It is interesting to note that the father of Flight Lieutenant Cook told us that, a few days before the crash, his son had expressed to him the view that neither his crew nor ZD 576 was yet ready to go on operations in Northern Ireland (Q 446).

119.  We found Witness A an impressive witness, who plainly felt it his duty to assist us as he had assisted the Sheriff.

Evidence bearing on the Boeing simulation

120.  Since the investigating board and the Air Marshals placed considerable reliance on the Boeing simulation it may be convenient to refer to it again in more detail at this stage. Before doing so however it is necessary to examine the functions of the two major controls in the aircraft. The collective increases or decreases the pitch of all the blades of the rotors as it is raised or lowered thereby causing the aircraft to climb or descend. At the same time movement of this control by connection to the FADEC system increases or decreases engine power to maintain rotor speed at approximately 100 per cent. The cyclic stick alters the pitch of the rotor head which is then tilted in the direction in which the aircraft is intended to go, namely forward, sideways or backwards.

121.  Detailed examination by the AAIB of the flight control system disclosed that the DASH extensions found did not correspond to a high speed level flight condition whereas the LCTA extensions did, and it appeared possible that the settings could reflect a dynamic aircraft manoeuvre at the point of impact. Boeing were therefore asked to undertake a study to assess the consistency of the settings and to define the possible manoeuvre. The simulation was a mathematical exercise which, as Mr Cable stated, was "looking really for fairly gross manoeuvres over a pretty short period of time" (Q 957). It was not intended to produce an accurate reconstruction of events but rather to demonstrate what could have happened within certain parameters (Q 982).

122.  Mr Cable provided Boeing with his findings from the wreckage of the aircraft as to pitch attitude, flight path angle, actuator extensions and ground speed together with certain other information provided by the board (Q 950). Information from the SuperTANS disclosed that:

123.  The Boeing simulation considered a wide range of possible starting conditions, i.e. conditions pertaining immediately prior to a final manoeuvre. Having rejected possible conditions at an airspeed of 135 knots, they concluded that an airspeed of 150 knots (groundspeed 174 knots) with a ROC of 1000 feet per minute provided "a ready match" with the criteria and was therefore the most likely (AAIB statement para 8). From this simulation, using among other things the state of the actuator extensions and attitude of the aircraft as found by Mr Cable, Boeing deduced that 2.9 seconds after the final manoeuvre had been initiated[35] the airspeed was about 135 knots, the rotor speed 204 rpm or 91 per cent design speed and the groundspeed 158 knots.

124.  It will be noted that, apart from the evidence of Mr Holbrook, there was no other evidence of the speed of the aircraft prior to the moment of impact. In the absence of a time at which the way point change took place and a position at which the height of the aircraft was disclosed, there are no facts from which the speed of the aircraft prior to the initiation of any final manoeuvre could be calculated. It follows that Boeing's 150 knots airspeed is a postulated figure rather than one calculated from known facts. This postulated figure then becomes the basis for the further postulated figure for ROC. Furthermore, the simulation gives no indication of the length of time prior to the assumed final manoeuvre during which the aircraft had been proceeding at the postulated airspeed.

125.  The rotor speed of 91 per cent derived from the simulation is significantly different from that of 100.5 per cent found by the AAIB on the instrument panel (statement para 7.2.2). Maintenance of rotor speed at or about 100 per cent design speed is of critical importance for safe helicopter flying. If rotor speed falls much below 90 per cent there is a danger of the blades of each rotor coning up and meeting at the top due to the reduction in centrifugal force which at higher speeds keeps them apart (Q 918). FADEC is designed to keep rotor speed at normal design speed and, if rotor speed had fallen to 91 per cent, maximum if not emergency power from the engines would have been expected. The position of the DASH actuators was not consistent with the use of such power. Mr Cable doubted whether the 91 per cent figure was accurate (Q 971) but he also explained how difficult it was to know the time to which the 100.5 per cent reading on the instrument panel related given the fact that there were at least three different impacts before the aircraft came to rest (Q 967).

126.  The groundspeed of 158 knots at impact derived from the Boeing simulation exceeded by 11 knots that of 147 found in the cockpit ground speed indicator (AAIB statement para 6). Moreover, the postulated ROC of 1000 feet per minute at 150 knots airspeed is unattainable. Squadron Leader Burke doubted whether it was achievable with ZD 576's load (Q 920). Witness A explained that while flying he had tried to see whether Boeing's chosen ROC was obtainable at 150 knots and had found that it was impossible in similar conditions (QQ 813-23). He had achieved no more than 400 feet per minute at 150 knots.

127.  Sir John Day had arranged for someone to fly the Chinook simulator at 150 knots; they achieved a ROC of 650 feet per minute (Q 1075). He accepted that a ROC of 1000 feet per minute and a speed of 150 knots were not compatible (Q 1075). However, he put it to us that he had always said that the ROC was about 1000 feet per minute, not precisely that. This comment was made in the context of his own calculations, based on the range of heights shown by the TANS some 15-18 seconds before impact, which showed that between those times and the start of the final flare the range of possible ROC was between 650 and 1350 feet per minute. The simulator which produced a ROC of 650 feet per minute at 150 knots also produced one of 1150 feet per minute at 135 knots, which was a speed for which Boeing found it "very difficult or impossible" to match the predicted conditions with the initial impact data (AAIB statement para 8, QQ 1074-80).

128.  The Boeing simulation postulations of a ROC of 1000 feet per minute and a speed of 150 knots were essential to the conclusion that a final flare was initiated some 4 seconds before impact. Now that those postulations have been shown to be unattainable, the circumstances and indeed existence of any such flare must be very doubtful. That there was such a flare was crucial to the Air Marshals' conclusion that the crew must have been in control of the aircraft for the last 4 seconds before impact (e.g. QQ 280, 1088). Sir John's calculations (above) give no support to such a conclusion, since they are independent of and in no sense a substitute for Boeing's postulations.

129.  Furthermore Mr Cable explained that the Boeing simulation did not model FADEC. "It had to be a representation of a simple engine governor for each engine, which would have really quite different characteristics, I think, in small areas anyway, from the FADEC" (Q 957). The simulation presupposed that the aircraft was at all times under control and flying a straight course although there was no evidence that this was necessarily the case.

130.  Mr Perks, who had worked on such simulations with a MoD team in the late 1970s and early 1980s, explained that for a given transient manoeuvre all the key modelled parameters had to be matched within reason to actual historical records. Two of the most important parameters were rotor speed and torque from the engines, in relation to which no historical records were available. He remarked on the disparity between the rotor speed required for the simulation and that indicated on the rotor speed gauge; and also on the fact that, whereas the simulation manoeuvre required engine power to be at absolute maximum, the indications found by the AAIB were that the engines were at an intermediate power setting (Q 183). Furthermore, none of the witnesses on the Mull who had heard the aircraft had noticed any noise suggesting a violent manoeuvre, and there were no data to suggest that the engines had exceeded normal values.

131.  Mr Perks proceeded,

    "On the Chinook Mk 2 aircraft the engine control systems have aircraft rotor speed as a primary input, with collective pitch as a supplementary input. If the rotor speed is too high, the engine is driven to idle power. If too low, the engine is driven to maximum output power. If collective pitch is changed, the engines will also be affected. Any form of extreme manoeuvring would have forced the engine control systems to respond immediately. The engine controls should, therefore, have been anywhere other than at normal settings. Normal settings implies the engine controls were not seeing major changes in their inputs, and that is not consistent with the violent manoeuvring postulated by Boeing. Whatever the pilots were doing, collective pitch was not being affected, and neither was rotor speed, given the evidence in the wreckage."

Thereafter he expressed the belief that the simulation should be "discounted" as evidence.

132.  Where does this leave the simulation? We conclude that it would be quite inappropriate to treat the results of the simulation as proven fact.

Evidence of the Air Marshals

133.  Sir John Day began his evidence before us with an interesting and helpful presentation involving 34 slides which are reproduced in HL Paper 25(i). He explained that the VFR rules applicable to the flight allowed the aircraft to be flown as low as 50 feet above ground with a minimum cloud base of 250 feet and minimum visibility of 1 kilometre[36]. IFR required the aircraft to be at least 1000 feet above the highest obstruction en route (Q 280). Sir John reckoned that the way point change was made some 20 seconds before impact; and he accepted as facts from the simulation that the groundspeed of the aircraft at the time of height disclosure was 160-175 knots with the aircraft in a cruise climb[37], and that the crew started to flare the aircraft some 4 seconds before impact.

134.  In Sir John's view, by continuing a cruise climb towards the Mull after the way point change, when they could have selected a high rate of climb, turned away from the high ground and stayed out of cloud or slowed down and then flown along the coast, the pilots "grossly breached the rules of airmanship" (Q 280). Sir John was clear that the gross negligence occurred at the way point change or a little before but he did not know when. If as he believed the crew voluntarily climbed into cloud, "the moment they decided to climb into cloud and go on to instrument conditions was the moment of negligence, because at that point they needed to take decisive action to make safety altitude as quickly as possible or not to fly any further towards the Mull until they had established that safety altitude" (Q 314).

135.  During the course of his evidence Sir John on more than one occasion emphasised that his conclusions were based on fact and not on hypotheses. It is therefore appropriate to look at some of the matters which he treated as fact. (Page references are to HL Paper 25(i).)

    (a)  "We know that about 20 seconds before impact with the ground the crew made a way point change" (Q 280, p 118 col 1). This figure which derives from the Racal report on the SuperTANS is based on a power down speed of 150 knots and a straight course from the WP change to impact at that speed. It is therefore at best an estimate and not a fact since the only factual evidence of speed at or after the change is the indication from the ground speed and drift indicator of 147 knots at initial impact (AAIB report, paragraph 7).

    (b)  "We know for a fact … that some four seconds before impact the crew started to flare the aircraft" (Q 280, p 117 col 1; Q 1088). Not so. The Boeing simulation, using assumptions now shown to be incompatible, produced this result. On no view could it be described as fact and there is no evidence either way as to what caused the aircraft to impact the ground in the position described in the AAIB report.

    (c)  "They had chosen to fly straight over the Mull of Kintyre, and we know that because they had set up this 1000 feet a minute ROC" (Q 301). There is no evidence that they had chosen to overfly the Mull, and indeed the making of the way point change suggests the contrary. Furthermore the 1000 feet a minute ROC derives entirely from the Boeing simulation with all its deficiencies referred to above.

    (d)  "What is for sure is that they were in a 1000 a minute cruise climb in that last 20 seconds before the final four seconds of flare" (Q 304). This is far from being sure given the deficiencies in the simulation already referred to.

    (e)  "We know they did not pull emergency power" (Q 311). Sir John later agreed that the impact could have destroyed any evidence of emergency power being pulled (Q 1097).

136.  An example of Sir John's reliance on facts appears in the evidence given on his first appearance before us: "The judgment I have made about gross negligence is not based on what I think may have happened, it is based on what I know happened from the facts I have described to you" (Q 321). The majority of these "facts" were the matters referred to in the preceding paragraph.

137.  Sir John in his remarks had discounted the possibility of a control jam, saying that the crew flew "a serviceable aircraft" into the hill. On first appearing before us he was asked about the possibility of the crew having lost control of the aircraft due to a control jam; he explained that if this had happened the pilots would have pulled emergency power, which they had not done because the relevant captions had not been activated. He therefore discounted this as a possibility (QQ 310-11). Nevertheless he later said that he could not exclude the possibility of a control jam having played a part in the accident (Q 339); and his acceptance that evidence of emergency power having been pulled could have been destroyed (see para 135(e) above) necessarily weakens his argument against such a jam having taken place. He emphasised however that "the crew put themselves into a position where they were going to hit the mountain and if any subsequent technical failure happened they had forsaken all the margins of safety which are imposed upon our aircraft". Although he could not exclude the possibility that some technical event such as an engine failure caption distracted the pilots, he considered it incomprehensible "that a minor emergency would have so distracted them that they forgot they were about to hit a mountain" (Q 340). Likewise Sir William Wratten conceded that the possibility of a control jam or engine malfunction could not be disproved (Q 1068).

138.  On 11 December 2001, some weeks after our last public hearing and when the first draft of our report had been almost completed, we received a document from the MoD entitled "Turning Performance of Chinook" (p 73 of HL Paper 25(ii)). This document stated the position of the MoD and was no doubt intended to support the views of the Air Marshals that the aircraft was already in a position of danger at the way point change. The document set out information about the radius of turn of which a Chinook is capable, in an attempt to show that it was inevitable that if the aircraft had turned at the point of way point change, banking 30 degrees, there would have been a crash. It is unfortunate that the Air Marshals did not provide us with this information on either of the two occasions on which they gave evidence. The document assumed a fairly high airspeed and did not take account of the possibility that, if the crew had on initiating a turn reduced speed, the radius of turn could have been greatly reduced. The formula for calculating the radius of turn contained in the document was helpful in demonstrating the dramatic effect of a reduction of speed in reducing that radius.

139.  On 7 January 2002 we received a further letter from the MoD offering "key information about the speed of ZD 576". This letter stated that at 1747 hrs the aircraft was observed on radar to be about 7 nm on the 27° radial from Belfast VOR. "The crash took place at 1759.30, 35 nm down track, so the aircraft must have maintained average cruising speed in excess of 150 knots groundspeed".

140.  If this information was considered to be of crucial importance, we fail to understand why it was not drawn to our attention several months previously. However, we do not consider that it is of such importance. First, while it produces an average speed over the 35 nm, it throws no light on the actual speed at the way point change. Second, even if the speed at the way point change were 150 knots or more, this would not reflect on the pilots' ability to reduce speed on making a turn to port. Third, it adds no support to the Boeing simulation, since 150 knots groundspeed equates in the circumstances to only 125 knots airspeed. Boeing found it impossible to match the predicted conditions with the initial impact data as found when the airspeed at the start of the final manoeuvre was 135 knots or below.

141.  Both Sir John and Sir William gave evidence for the second time after we had heard all the other witnesses. Sir John explained that way point A entered in the TANS was in fact some 280 metres south east of the lighthouse. He produced a map (reproduced in HL Paper 25(i), p 157) showing the track of the aircraft to the programmed way point with a mark thereon for the point at which the aircraft would have been one kilometre from the land. He considered that if the crew had intended to fly VFR along the west side of the Mull they should, at the high speed at which they were travelling, have altered course at that point; and that they were negligent in not having done so (Q 1042). Had they turned at the point of way point change which was only 600 metres from the cliffs they would have been in a dangerous position. Had they turned where they thought they were when they made the way point change, namely to the west of the actual track, the cliffs would have been about 1 kilometre ahead (see map) and the change would not have attracted criticism from Sir John and Sir William (Q 1039).

142.  Sir William also considered that if the crew were visual with the Mull they should have turned left at 1 kilometre away (Q 355). He stated that if there had been a control jam preventing an alteration of course immediately after the way point change the pilots would have been negligent because "they had brought upon themselves an emergency, a crisis of time" (Q 376).

143.  Sir William summed up his position by agreeing with the proposition that the aircraft should not have been in either of the way point change positions (namely that in which they actually were and that in which they thought they were) and that any supervening circumstances affecting the flight at or after reaching these positions were irrelevant (QQ 438-42). In fairness to Sir William it should be explained that these answers were simply agreements to propositions put to him, and should be read together with his answers to QQ 335 and 1039. Both Sir John and Sir William considered that the aircraft must have been under control when the way point was changed and when the final flare was initiated 4 seconds before impact.


27   See p 84 of HL Paper 25(ii). Back

28   An engine "runaway" is an engine run up which is not controlled. Back

29   Mr Cable compared the effect to losing the power steering in a car. Back

30   By Corpus Christi Army Depot Analytical Investigation Division, ref SIOCC-QP-AI USASC 97-305. Back

31   Reproduced on p 69 of HL Paper 25(ii). Back

32   Emphasis original. Back

33   Report by Racal Avionics. See footnote 22. Back

34   Investigating board report para 49. Back

35   Mr Cable told us, "The manoeuvre started at the one second point from an arbitrary zero" (Q 987). Hence the Air Marshals' description of the final manoeuvre as lasting 4 seconds (eg QQ 280, 304). Back

36   If flying slower than 140 kts indicated airspeed. If faster, 1.5 km. Illustrated on slides 4 and 5. Back

37   A cruise climb is a climb at high forward speed and low ROC. To increase ROC, forward speed must be reduced. Back


 
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