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Research group of Prof. Peter B. Ladkin, Ph.D.
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Extracts from
UK AAIB Report 4/90
on the 8 January 1989 accident
of a British Midland B737-400
at Kegworth, Leicestershire, England

excerpts edited by P. B. Ladkin

P. B. Ladkin, 29 November 1996 Appendix of the Report

[The following excerpts from the Kegworth report have used some text digitised by Colin Sandall in

This project was supervised by Peter Mellor, to whom thanks for transmitting Colin's ASCII version to me.

This extract includes the entire SYNOPSIS and portions of various sections. Section 3: Conclusions and Section 4: Safety are included in their entirety.

  • Comments by the editor are included in square parens with initials, just like this one. PBL]

    Air Accidents Investigation Branch

    Aircraft Accident Report No: 4/90

    Registered OwnerKommanditbolaget 11, Malmo, Sweden
    Operator British Midland Airways Ltd
    Aircraft Type: Boeing 737
    Model:Series 400
    Place of accident:1/2 nm east of East Midlands Airport
    Latitude: 52° 49' 54" N
    Longitude: 001° 20' 54" W
    Date and time:8 January 1989 at 2025 hrs
    All times in this report are UTC


    The accident was notified to the Air Accidents Investigation branch during the evening of the 8 January 1989 and the investigation was initiated on-site at 0400 hours on the morning of the 9 January. The AAIB Investigating Team comprised Mr E J Trimble (Investigator in Charge), Mr J D Payling (Operations), Mr C G Pollard (Engineering, Powerplants), Mr S W Moss (Engineering, Systems), Mr R D G Carter (Engineering, Structures), Wing Commander D Anton, RAF Institute of Aviation Medecine (IAM) (Survivability). Mr P F Sheppard and Miss A Evans (Flight Recorders). In addition Mr R Green, Head of the Psychology Department of the RAF IAM, was co-opted to investigate the human factor aspects of this accident and Captain M Vivian of the Civil Aviation Authority (CAA) Flight Operations Department was co-opted to assist the final assessment of the operational aspects.

    G-OBME left Heathrow Airport for Belfast at 1952hrs with 8 crew and 118 Passengers (including 1 infant) on board. As the aircraft was climbing through 28,300 feet the outer panel of one blade in the fan of the No 1 (left) engine detached. This gave rise to a series of compressor stalls in the No 1 engine, which resulted in airframe shuddering, ingress of smoke and fumes to the flight deck and fluctuations of the No 1 engine parameters. Believing that the No 2 engine had suffered damage, the crew throttled thhat engine back and subsequently shut it down. The shuddering caused by the surging of the No 1 engine ceased as soon as the No 2 engine was throttled back, which persuaded the crew that they had dealt correctly with the emergency. They then shut down the No 2 engine. The No 1 engine operated apparently normally after the initial period of severe vibration and during the subsequent descent.

    The crew initiated a diversion to East Midlands Airport and received radar direction from air traffic control to position the aircraft for an instrument approach to land on runway 27. The approach continued normally, although with a high level of vibration from the No 1 engine, until an abrupt reduction of power, followed by a fire warning, occurred on this engine at a point 2.4 nm from the runway. Efforts to restart the No 2 engine were not successful.

    The aircraft initially struck a field adjacent to the eastern embankment of the M1 motorway and then suffered a severe impact on the sloping western embankment of the motorway.

    39 passengers dies in the accident and a further 8 passengers died later from their injuries. Of the other 79 occupants, 74 suffered serious injury.

    The cause of the accident was that the operatin crew shut down the No 2 engine after a fan blade had fractured in the No 1 engine. This engine subsequently suffered a major thrust loss due to secondary fan damage after power had been increased during the final approach to land.

    The following factors contributed to the incorrect response of the flight crew:
    1. The combination of heavy engine vibration, noise, shuddering and an associated smell of fire were [sic PBL] outside their training and experience.
    2. They reacted to the initial engine problem prematurely and in a way that was contrary to their training.
    3. They did not assimilate the indications on the engine instrument display before they throttled back the No 2 engine.
    4. As the No 2 engine was throttled back, the noise and shuddering associated with the surging of the No 1 engine ceased, persuading them that they had correctly identified the defective engine.
    5. They were not informed of the flames which had emanated from the No 1 engine and which had been observed by many on board, including 3 cabin attendants in the aft cabin.

    31 safety recommendations were made during the course of the investigation.

    1. Factual Information

    [.......] EIS primary display

    The following engine parameters were displayed:-

    These parameters were presented in both analogue and digital form by the use of light emitting diodes (LEDs). The analogue presentation utilised 81 bars of LEDs, arranged radially around the outside of each display scale. The bars illuminated one at a time, in sequence, to simulate the movement of the end of a pointer sweeping around the outside of the display scale. Other design features concerning the movement of the LED 'pointer' were also incorporated, in order to mimic the behaviour of an electro-mechanical indicator.

    The digital presentation, which was common to both the EIS and earlier hybrid instruments, was situated in the centre of each indicator and also used LEDs. These simulated the rolling drum mechanism used on conventional electro-mechanical indicators by making the display digits appear to 'roll' past the viewing aperture, with half of each adjacent digit visible in the last 'window'. This preserved the rate and direction of motion cues available to the pilot. Red exceedance warning lights are positioned above each N1,N2 and EGT display and were designed to illuminate whilst the affected parameter remains above the 'read-line' limit. Exceedance information was stored in a non-volatile memory which could be interrogated by maintenance personnel.

    Both N1 displays also featured movable LED cursors to indicate 'target N1' which could be set manually be using two knobs located in the lower corners of the display bezel, or automatically by the flight management computer. When set manually, this information was repeated in digital form in two windows at the top of the display. A button at the bottom of the display bezel was used to change the reading of fuel flow rate to fuel used. After 10 seconds the display automatically revertd to 'fuel flow'.

    A three-character display at the top of the primary EIS annunciated the thrust mode as selected through the flight management computer.

    An 'abnormal start' algorithm was incorporated which would cause the EGT digits to flash if the unit detected conditions such as incipient 'hot', or 'hung', starts. EIS Secondary Display

    The second EIS displayed the following parameters for both engines, in analogue form only:-

    The system of scales and LED 'pointers' was similar to that used in the primary display but there were no digital repeaters and, since the secondary display are smaller [reference omitted PBl], there were 31 bars of LEDs to simulate the pointer. There were digital readouts of engine oil quantity, hydraulic quantity (% full) and total air temperature (TAT). In common with normal practice on engine instruments classified as secondary, there were no exceedence lights on the secondary EIS. Features common to both EIS displays

    Both primary and secondary displays were fitted with numerous sensors which varied the LED brightness according to the amount of ambient light falling on the display. Thus the displays were designed to remain legible under all conditions, including situations where the ambient light fell differentially across the displays. For night operation the scales were edge-lit and their brightness was controlled by the crew, through the normal panel lighting control.

    Both displays featured built-in test equipment (BITE), activation of which would cause the unit to run through a test programme. Use of BITE was restricted to ground maintenance only.

    Engine parameters were received by the EIS direct from the sensors on the engines with the exception of vibration, where signals from the sensors were processed by the airborne vibration monitor (AVM) before being passed to the EIS (see paragraph 1.6.4). The EIS fed its output of these parameters (except vibration and engine oil temperature) to the flight control computers, the flight data acquisition unit (FDAU) and the stall warning computer, as required.

    The EIS was connected to the aircraft wiring through four connectors located on the back of each unit. The input and output wiring associated with each engine was fed through a discrete connector, which had a baulk system known as 'clocking'. This physically prevented inadvertent cross-connection.

    The EIS reacted to input system failures in different ways, depending on the type of input and the nature of the failure. In all cases where a definite system failure was detected, the EIS would delete the affected parameters from the display.


    1.11.1: Flight Data recorder

    The aircraft [a Boeing 737-400 PBL] was fitted with a Sundstrand Universal Flight Data Recorder (UFDR) with a recording duration of 25 hours on a magnetic (kapton) tape, and a Teledyne flight data acquisition Unit (FDAU). A total of 63 parameters and 90 discrete events were recorded. In addition, the FDAU was equipped with a computer type 3.5" [originally `3-1/2"' PBL] `floppy' disk, which recorded `snapshots' of routine information and data associated with specific exceedances. The FDR was located in the rear passenger cabin above the cabin roof, in line with the rear passenger exits.

    The UFDR takes flight data into one of the two internal memory stores, each holding about one second of data. When one memory store is full, the data flow is switched to the other store. While the data is being fed to this other store, the tape is re-wound and the previous second of data is checked. A gap is left on the tape and the data in the first store is then written to the tape, and the first memory store emptied. This whole 'checkstroke' operation takes much less than one second to complete so that once the other store is full, data is switched back to the first store, and the second store is written to tape using the `checkstroke' operation again to check its data. The procedure is then repeated.

    Thus the UFDR tape is not running continuously. The tape first accelerates from stationary to 6 inches per second to read the previous data block, leaves an inter-record gap and then writes the new data block. The tape then slows and rewinds ready to begin the next `checkstroke' operation. A total of 0.48 inches of tape is used to record one block of data and inter-record gap.

    Data is formatted by the FDAU into one second subframes, each subframe begins with a synchronisation code, and is followed by the other parameters in a 64 word set format. The start of a block of data stored in the internal memory may not coincide with the start of a subframe, so when a block is recorded onto tape it is preceded by 'pre-amble' data bits and followed by 'post-amble' data bits. These bits of data are recognised during replay and removed, producing a continuous datastream. The start of a frame is identified from the synchronisation code.

    When power is lost from the recorder, the data held in the volatile memory which has not been recorded on the tape is lost. As can be seen from the way in which data is temporarily stored on this UFDR and then recorded, this can mean that up to 1.2 seconds of data may be lost just before impact. Analysis of the raw signal on the UFDR tape from ME showed that the recorder had completed writing the contents of one memory store to tape, and this stopped at word 30 subframe 2. It was not possible, in this case, to know exactly how much data had been lost, and obviously as this was the last information prior to impact, such data could have been important to the investigation.

    [Three paragraphs, decribing respectively how the UFDR gets its information from the EIS, how the pointer display gets its signals, and how the UFDR gets its vibration signals, omitted. PBL]

    [Section 1.11.2, FDR Data Analysis, and Section 1.11.3, Cockpit voice recorder (CVR), a physical description of the CVR, omitted PBL]


    1.11.4 CVR transcript significant events

    From the CVR it was apparent that the first indication of any problem with the aircraft was as it approached its cleared flight level, when for a brief period, sounds of `vibration' or `rattling' could be heard on the flight deck. There was an exclamation and the first officer commented that they had 'GOT A FIRE', the autopilot disconnect audio warning was then heard, and the first officer stated 'ITS A FIRE COMING THROUGH'. The commander then asked 'WHIXH ONE IS IT?', to which the first officer replied, 'ITS THE LE..ITS THE RIGHT ONE'. The commander then said 'OKAY, THROTTLE IT BACK'.

    London ATC was then called by the first officer, advising them of an emergency, after which the commander asked for the engine to be shut down. The first officer began to read the checklist for 'Engine Failure and Shutdown' but was interrupted by ATC calls and by the commander's own calls to the operating company during which the decision was made to divert to East Midlands. Approximately 2 minutes after the initial 'vibration' the final command was given to shut down the engine. The first officer then recommenced the checklist and 2 minutes 7 seconds after the initial engine problem he moved the start lever of the No 2 engine to 'OFF'. He then started the APU. Throughout this period no fire audio warning was heard.

    The aircraft then started the descent to East Midlands Airport and the commander made his first announcement to the passengers during which he mentioned that they had had a problem with their right hand engine which had produced some smoke in the cabin. The flight crew were then fully occupied with the relevant checklists, calls to the operating company and ATC, who were routeing [sic] them into East Midlands, and reprogramming the flight management system (FMS) for an East Midlands diversion, with which they had some difficulty. During this period they also briefly discussed the symptoms that had occurred initially and the commander mentioned 'RAPID VIBRATIONS IN THE AEROPLANE - SMOKE'.

    The flight proceeded until the aircraft was on final approach with the landing checklist completed. Just after they had comfirmed with East Midlands ATC that the right engine had been shut down, there was a crackling noise on the CVR, possible due to electrical interference. This occurred 54 seconds before the first ground impact. Leading up to this event there were significant changes in the frequency content of the background noise on the CVR area microphone, which are discussed in paragraph 1.11.5. These changes would probably not have been audible to the crew.

    Immediately following this, a transmission was made to the tower indicating that the crew was having trouble with the second engine as well and the commander asked the first officer to 'TRY LIGHTING THE OTHER ONE UP - THERE'S NOTHING ELSE YOU CAN DO'.

    36 seconds before impact the (No 1 engine) fire bell sounded. The first officer asked the commander if he should shut this engine down. The commander replied in the negative. The CVR recording then indicated their intention to 'stretch the glide', but at 29 seconds before impact the ground warning proximity system (GPWS) 'glideslope' warning commenced and continued with increasing repetition rate, indicating that the aircraft was steadly diverging below the glidepath. The commander twice said "TRY OPENING THE OTHER ONE UP' and each time the first officer said 'SHE'S NOT GOING'. At 10 seconds before impact the commander made an ammouncement to the passengers to 'PREPARE FOR CRASH LANDING' (repeated). The stick shaker was them heard operating, followed by the sounds of impact. Relevant comments from the CVR transcript are shown in relation to the FDR information in the Appendix 4, figs 2, 5, & 7.

    [.......] Engine instrument system (EIS) and associated wiring

    The primary and secondary EIS units were still in position within the centre instrument panel but the secondary unit had suffered some distortion which had caused the glass face to crack. Prior to removal, the wiring and connectors to and from both units were checked for left/right sense and found to be satisfactory.

    Upon removal of the EIS display units it could be seen that, whilst the primary unit was virtually undamaged, the secondary unit had received an impact on the upper rear case due to contact with deformed structure behind the panel. Loose items, which were later found to be a small screw, nut and washers broken from the bezel, could be heard rattling within the case.

    Both EIS units were examined at the manufacturer's premises under AAIB supervision. Power was applied to the primary display and, as the unit appeared to function normally, it was subjected to a BITE check which revealed no faults. The exceedance memory was interrogated and it was established that no exceedances had been recorded. When the unit was subjected to a full acceptance test, it met all the requirements of this test satisfactorily.

    Despite the damage to the case of the secondary display, an internal inspection suggested that the unit should function. However, when power was applied a malfunction was apparent within the right-hand (No 2 engine) channel. The engine oil pressure and hydraulic pressure displays on this side at all times displayed multiple pointers whilst the digital oil quantity indication also displayed in an obviously faulty manner. The other parameters on this No 2 side (including the No 2 vibration display) read correctly, as did all the parameters on the No 1 left engine side. The fault was correctly diagnosed as a defective hybrid integrated circuit associated with the display drive. The board containing this part was removed and the defective component replaced. Thereafter the secondary EIS worked normally and fully met the acceptance test requirements.

    The defective microchip appeared to have broken and, since it was located close to the sheared screw mentioned above, it was suspected that it had received an impact at the same time as the screw. When the microchip was examined by the manufacturer in the USA, it was confirmed that it had broken due to mechanical shock.

    It was concluded from these tests that both EIS units were functioning at the time of impact and correctly displaying the input information.


    1.17.3 Pilot opinion of the EIS

    During May and June 1989 an informal survey of pilot opinion of the EIS among Boeing 737 pilots in UK airlines was conducted by AAIB. Replies were received from 120 pilots, representing over 90% of all pilots in the UK with experience of the EIS. When the results of the survey were received, its was noted that the replies of the 53 pilots of BMA were more critical of the EIS than those of other airlines.

    The overall results, however, showed that a large majority of pilots considered that the EIS displayed engine parameters clearly and most considered that it also showed rate of change of parameter clearly. Fewer than 10% of pilots reported any difficulty in converting to the EIS from the earlier hybrid instruments and only a few reported minor difficulty in alternating between the EIS and hybrid displays. 64% of pilots stated that they preferred instruments with electro-mechanical pointers to the EIS (74% of BMA pilots, and 58% of non-BMA pilots).

    The most unfavourable assessments were received in response to the question 'Is the EIS effective in drawing attention to rapid changes in engine parameters'. 64% of BMA pilots and 39% of non-BMA pilots thought it was not. 43 pilots (31 from BMA and 12 others) wrote additional comments in answer to this question in which they criticised the LED pointers of the EIS as too small, lacking in visual impact and not as conspicuous as the pointers on electro-mechanical instruments. These pilots considered that their attention was more likely to be attracted by the movement of full-length pointers than by changing numeric displays and the shorter LED pointers. An analysis of the human factors aspects of the EIS design is included at Appendix 2.7.

    [.......] Incident to Dan Air Boeing 737-400, G-BNNL on 9 June 1989

    In this incident, the aircraft was climbing at -3C rated power through FL250 when the crew reported symptoms which were initially identical to those reported by the crew of ME [G-OBME PBL], including the almost immediate presence of smoke in the cockpit, but subsequently with a much higher continuing vibration level which persisted even whilst the engine windmilled after shut down. The crew identified the failed No 1 engine correctly, by reference to the engine vibration indicators, and completed a full shut-down drill.

    The FDR showed that the engine was unable to sustain power and ran down before the crew began the engine shut-down drill. Fuel flow began to reduce within 2 seconds of the onset of the vibration and reached zero within 9 seconds; the throttle was partially closed after 6 seconds and fully closed after 24.

    Although the No 1 engine had quickly shut itself down, it appeared to the commander that, except for the No 1 engine vibration indicator, all the engine instruments looked alike and it took time for him to be certain that he had correctly identified the affected engine. When asked to explain his observations in greater detail he said that he had glanced repeatedly at the engine instruments and they appeared to be well-matched. He saw no tumbling digits and made his assessment solely on the basis of the high vibration shown on the No 1 engine vibration indicator. However, he later qualified this statement by saying that, having identified the No 1 engine as the source of the vibration, he could not recollect monitoring the No 1 engine instruments again. He added that, in conditions of high vibration, the small size of the EIS instruments and their manner of presenting information made them difficult to read. The aircraft landed without further incident. [Remainder of section describes physical damage to the engine. PBL]


    1.17.19 Airborne closed circuit television monitoring

    With the advent of miniature television cameras and recording equipment it has become apparent that flight crews could be presented with external views of their aircraft during flight. A logical extension of this concept would be the monitoring of previously 'blind' areas such as cargo holds and closed wheel bays in addition to the recording of flight-deck activity and instrumentation for accident investigation purposes. AAIB became involved when the Robert Gorden Institute of Technology (RGIT) proposed a system based on the Panasonic miniature camera and invited AAIB comments. A meeting with the CAA confirmed their interest and it was proposed that proof of concept trial would be mounted using the Panasonic camera on the RAE [footnote: Royal Aerospace Establishment at Farnborough] BAC 1-11. Subsequent to this decision, other agencies declared an interest in the project and the CAA found it necessary to go to competitiev tendering for the proof of concept trial. During this investigation, AAIB have remained in close contact with RAE, RGIT, and CAA and have hosted several presentations by equipment manufacturers. A recommendation relating to the provision of flight deck crews with external views of their aircraft was made in the AAIB Report on the accident to Boeing 737 G-BGJL at Manchester International Airport on the 22 August 1985 (Report 8/88).

    The aims of such an Airborne closed circuit television monitoring system would be as follows:-
    (i) To provide flight deck crew with an extermanl view of the major areas of their aircraft in all weathers, both day and night.
    (ii) To provide flight deck crews with a view of cargo bays to aid in decision making following an associated fire warning
    (ii) To provide a record of flight deck activity and flight instrumentation for the purposes of in-flight post-incident recall by flight crews and accident investigation. This aspect would be of particular use in 'glass cockpits' and would provide a valuable supplement to the future provision of full digital recording of flight instrument parameters on the FDR.
    (iv) To provide cabin crews with a view of areas of the cabin not currently manned by cabin staff. This system could also be used as a security system whilst the aircraft was on the ground.
    (v) To provide a recording of all such information, that would be protected to the same standards as those specified for FDRs and CVRs.


    2. Analysis

    [.......] The engine instrument system

    The failure to detect, or at least to identify correctly, disparities in the readings of the engine instruments is perhaps most important with regard to the vibration indicators. Unlike the transient fluctuations that would have appeared on the primary engine instruments, the reading on the No 1 engine vibration indicator rose to maximum and remained there for about 3 minutes. On the EIS, however, not only is the pointer of this vibration indicator much less conspicuous than a mechanical pointer (Appendix 2, figs 1 & 2) but, when at maximum deflection, it may be rendered even less conspicuous by the close proximity of the No 1 engine oil quantity digital display, which is the same colour as the pointer and is the dominant symbology in that region of the display (Appendix 2, fig 3). In view of the limited attention both pilots appear to have given to the vibration indicators, it is a matter for conjecture whether or not they would have failed to notice such a maximum reading on the mechanical pointer of a hybrid display, clearly separate from any other distracting indication, but there can be little doubt that it would have been easier to see.

    The informal survey of pilot opinion of the EIS (paragraph 1.17.3) showed that 64% preferred engine instruments will full length mechanical pointers. This finding was almost certainly influenced by lack of familiarity since the survey was conducted when the EIS was relatively new in service, and it is not surprising that at that time most pilots should have expressed a preference for the older type of display, with which they were familiar. The result of the survey was also influenced by the replies of the BMA pilots, which were more critical of the EIS than those of other airlines. Also, because of a natural resistance to change, the fitness of new equipment for its purpose may not be judged on pilot preference alone, although this must be an important factor. With these reservations, the least favorable interpretation of the results was that the EIS displayed engine parameters clearly, but its ability to attract attention to rapidly changing readings was less satisfactory. The latter aspect, however, was less important in the case of this accident because the crew were alerted to abnormal operation by other signs and had time, or should have taken time, to study the engine instrument readings.

    One finding of the pilot survey, that the LED pointers of the EIS are less conspicuous than the mechanical pointers on the hybrid displays, is however a cause for concern. In this respect, whilst the introduction of the EIS may represent progress in terms of improved reliability and maintainability, it may be a retrograde step in terms of presentation of information. Moreover, although it was type-certified as fit for its purpose by both the FAA and the CAA during October 1988, it appears to have been introduced without any thorough evaluation of its efficiency in imparting information to line pilots. Now that the system has been in use for some time and EIS-equipped flight simulators are available, the reduced conspicuity of the pointers may assume less importance and it may be too late for a new evaluation of the system to be worthwhile. Nevertheless, this change in presentation indicates how important it is for all new developments in aircraft indicating systems to be subjected to comprehensive evaluation of their effect on line pilot performance before being introduced to service. It is therefore recommended that the regulatory requirements concerning the certification of new instrument presentations should be amended to include a standardized method of assessing the effectiveness of such displays in transmitting the associated information to flight crew, under normal and abnormal parameter conditions. In addition, line pilots should be used in such evaluations (Made 30 March 1990).

    The layout and methods of displaying information on engine instruments are considered in Appendix 2.7, which concludes that although the EIS provides accurate and reliable information to the crew, the overall layout of the displays and the detailed implications of small LED pointers, rather than larger mechanical ones, and of edge-lit rather than reflective symbology, required further consideration. Neither pilot noticed the maximum reading on the No 1 engine vibration indicator and at least one of them gained the impression from the engine instruments, or from some other cue, that the No 2 rather than the No 1 engine was failing. If they gained this impression from the engine instruments then there is a possibility that the methods of displaying information on these instruments may have contributed to the error.

    The error would probably not have been made if the vibration indicators had included a visual warning of which engine was affected by excessive vibration. It may be seen from the preamble to the FAA requirement for a vibration monitoring system (paragraph 1.17.4) that it was intended to provide the flight crew with a vibration warning, and later aircraft powered by certain high-bypass turbofan engines, such as Boeing 757 aircraft and the Airbus series, do include high engine vibration in their crew alerting systems. It is therefore recommended that the CAA should require that the engine instrument system on the Boeing 737-400 aircraft type, and other applicable public transport aircraft, be modified to include an attention-getting facility to draw attention to each vibration indicator when it indicates maximum vibration (Made 30 March 1990).


    2.5 Flight recorder design requirements

    The system of recording using temporary buffer storage as employed by the UFDR can mean that at impact, if the contents of the buffers have not been transferred onto the recording medium, then that information will be list. In the UFDR this can be up to 1.2 seconds of data. In this instance a knowledge of the impact parameters was important to the survivability investigation. The loss of the last moments of data meant that the impact parameters had to be estimated. The lost data in the buffer may have yielded more accurate information. If a recorder has to employ a temporary buffer storage, that storage medium should be made non-volatile (i.e. recoverable after power off) and contained within the armour protected enclosure.

    The European Organisation for Civil Aviation Electronics (EUROCAE) are at present formulating new standards [footnote: Minimum Operational Performance Requirement for Flight Data Recorder Systems. Ref:- ED55] for future generation flight data recorders; these standards will permit delays between parameter input and recording of up to 0.5 seconds. These standards may be adopted worldwide and do form the basis of the new CAA specifications for flight data recorders. It is therefore recommended that the manufacturers of existing flight data recorders which use buffering techniques should give consideration to making the buffers non-volatile and hence recoverable after loss of power, and EUROCAE and the CAA should reconsider the concept of allowing volatile memory buffering in flight data recorders (Made 30 March 1990 and also included in AAIB report 2/90 [Lockerbie. PBL]).

    Because of the length of time (64 seconds) between successive samples of the engine vibration, it was not possible to be precise about when vibration levels increased or decreased. Whilst this is not a parameter that the CAA specifications require to be recorded, it is recommended that, where engine vibration is an available parameter for flight data recording, the CAA should consider making a requirement for it to be recorded at a sampling rate of once per second (Made 30 March 1990)


    3. Conclusions

    (a) Findings

    The aircraft

    1. The aircraft had a valid certificate of airworthiness in the transport category (passenger) and had been maintained in accordance with an approved schedule.

    The flight deck crew

    2. The flight deck crew were properly licensed and rested to undertake the flight.

    3. The flight deck crew experienced moderate to severe engine induced vibration and shuddering, accompanied by smoke and/or smell of fire, as the aircraft climbed through FL283. This combination of symptoms was outside their training or experience and they responded urgently by disengaging the autothrottles and throttling-back the No 2 engine, which was running satsifactorily.

    4. After the autothrottle was disengaged, and whilst the No 2 engine was running down, the No 1 engine recovered from the compressor stalls and began to settle at a slightly lower fan speed. This reduced the shuddering apparent on the flight deck, convincing the commander that they had correctly identified the No 2 engine as the source of the problem.

    5. The first officer reported the emergency to ATC, indicating that they had an engine fire and intended to shut and engine down, although there had been no fire warning from the engine fire detection system.

    6. Whilst the commander's decision to divert to East Midlands airport to land with the minimum of delay was correct, he thereby incurred a high cockpit workload which precluded any effective review of the emergency or the actions he had taken.

    7. The flight crew did not assimilate the readings on the engine instruments before they decided to throttle-back the No 2 engine. After throttling back the No 2 engine, they did not assimilate the maximum vibration indication apparent on the No 1 engine before they shut down the No 2 engine 2 minutes 7 seconds after the onset of vibration, and 5 nm south of EMA. The aircraft checklist gave separate drills for high vibration and for smoke, but contained no drill for a combination of both.

    8. The commander remained unaware of the blue sparks and flames which had issued from the No 1 engine during the period of heavy vibration and which had been observed by many passengers and the three aft cabin crew.

    9. During the descent, the No 1 engine continued to run apparently normally, although with higher than normal levels of vibration.

    10. Flight crew workload during the descent remained high as they informed their company at EMA of their problem and intentions, responded to ATC height and heading instructions, obtained weather information for EMA and the first officer attempted to reprogramme the flight management system to display the landing pattern for EMA. Some 7.5 [written 71/2 PBL] minutes after the initial problem, the commander attempted to review the initial engine symptoms, but this was cut short by further ATC heading and descent information and instructions to change to the EMA ATC radio frequency.

    11. Fifteen minutes ater the engine problem occurred and some 4 minutes 40 seconds before ground impact, the commander increased power on the No 1 engine as the aircraft descended towards 3000 ft amsl and closed with the centreline of the instrument landing system. At this point, the indicated vibration on the No 1 engine again rose to its maximum value of 5 units but did not attract the attention of either pilot.

    12. Fifty three seconds from ground impact, when the aircraft was 900 feet agl and 2.4 nm from the runway with landing gear down and 15° flaps selected, there was an abrupt decrease in power from the No 1 engine.

    13. The commander immediately called for the first officer to relight the No 2 engine. The attempted restart was not successful, probably because there was insufficient bleed air pressure from the No 1 engine, pressure air from the APU was not connected and the bleed air crossfeed valve was closed. Even if pressure air had been available it is unlikely that power could have been obtained from the No 2 engine before the aircraft hit the ground.

    14. The training of the pilots met CAA requirements. However, no flight simulator training had been given, or had been required, on the recognition of engine failure on the electronic engine instrument system or on decision-making techniques in the event of failures not covered by standard procedures.

    15. The change from hybrid electro-mechanical instruments to LED displays for engine indications has reduced conspicuity, particularly in respect of the engine vibration indicators. No additional vibration alerting system was fitted that could have highlighted to the pilots which of the two engines was vibrating excessively.

    The Cabin Crew

    16. All members of the cabin crew were properly trained to undertake the flight.

    17. Although the cabin crew immediately became aware of heavy vibration at the onset of the emergency and three aft cabin crew saw flames emanating from the No 1 engine, this information was not communicated to the pilots.

    18. During the descent, the cabin crew carried out their emergency drills, checking that all passengers had their lap belts fastened and stowing all loose carry-on luggage in the overhead bins.

    No 1 (Left) engine

    19. The No 1 engine suffered fatigue of one of its fan blades which caused detachment of the blade outer panel. This led to a series of compressor stalls, over a period of 22 seconds, until the engine autothrottle was disengaged.

    20. The severe mechanical imbalance which arose because of the outer panel separation led to blade tip rubbing, particularly on the fan and booster sections abradable seals, which caused smoke and the smell of burning to be passed into the air conditioning system.

    21. About 3 seconds after the autothrottle was disengaged, and whilst the No 2 engine was running down, the No 1 engine began to stabilise. However, its indicated vibration remained at maximum for at least 3 minutes until this engine was throttled back for the descent.

    22. The evidence indicated that the timing of the sudden recovery of the No 1 engine from the compressor stalling was related to the autothrottle disengagement at a point when it had demanded a lower throttle lever angle than that required for rated climb, thereby allowing this engine to achieeve stabilised running at a slightly lower speed.

    23. During the descent, the No 1 engine respnded apparently normally at the idle/low throttle settings used, although its indicated vibration remained higher than normal.

    24. Fifty three seconds before ground impact, the No 1 engine abruptly lost thrust as a result of extensive secondary fan damage. This was accompanied by compressor stalling, heavy buffetting and the emission of pulsating flames. This damage was probably initiated by fan ingestion of the blade section released by the initial failure, which was considered to have partially penetrated, and temporarily lodged within, the acoustic lining panles of the intake casing before having been shaken-free during the period of high vibration following the increase in power on the final approach to land. Sections of fan blades were found below this point of the final approach, including two small fragments which were determined to be remnants of the blade section which detached initially.

    25. The No 1 engine fire warning, which occurred on the flight deck 36 seconds before ground impact, was initiated by a secondary fire which occurred on the outboard exterior of the engine fan casing. It was concluded that the prolonged period of running under conditions of excessive vibration had loosened fuel/oil system unions and seals on the exterior of the fan casing and that the inlet duct had probably been damaged sufficiently, by fan blade debris, to allow ignition of atomised fuel/oil sprays by titanium `sparks' and/or intake flame.

    26. This short duration in-flight fire on the No 1 engine was followed by a localised ground fire associated with this engine, which was successfully extinguished by the East Midland Airport Fire Service.

    27. The fan blade fatigue fracture initiated as a result of exposure of the blade to a vibratory stress level greater than that for which it was designed, due to the existence of a fan system vibratory mode, induced inder conditions of high corrected fan speed at altitude, which was not detected by engine certification testing.

    No 2 (right) engine

    28. The No 2 engine was running normally when it was throttled back to flight idle, and then shut down.

    29. This engine showed no evidence of power at impact, consistent with the evidence from the flight data recorder.

    30. Detailed strip inspection of this engine showed it to have been fully serviceable before ground impact.


    31. The No 2 (right) engine vibration reports which appeared in the aircraft Technical Log during December 1988 but had been correctly addressed by ground technicians [sic PBL]

    32. There were no malfunctions of the major airframe systems which contributed to this accident.

    33. No evidence was found of any cross-connection or similar obvious wiring errors associated with either the engine instrument system (EIS) or the fire detection system.

    34. The EIS fitted to the aircraft was serviceable at impact and tests indicated that it should have displayed those primary engine parameters recorded on the FDR, with close fidelity.

    35. The airborne vibration monitoring system (AVM) was serviceable at impact. Tests showed that the system was capable of tracking vibration caused by the massive fan imbalance and of outputting its maximum value approximately 2 seconds after the start of the vibration.

    36. Flight crew reports concerning the response of the AVM system during the two other cases on fan blade fracture on CFM56-3C engines which occurred subsequent to this accident supported the behaviour described above. Two cases of bird impact which resulted in fan damage generated crew reports of late indication on vibration gauges, although vibration was clearly felt by the flight crew. This was the result of the non-linear sensitivity of this engine type to small imbalances with changes of fan speed in the take-off and climb thrust range.

    37. The engine fire and overheat detection system contained a fault which could have rendered it incapable of providing warning of a fire in either engine. However, the CVR evidence indicated that it did, in fact, provide a warning of the fire in the No 1 engine 36 seconds before impact.

    Impact with the ground

    38. The aircraft suffered two distinct impacts with the ground, the first just before the eastern embankment of the M1 motorway and the second on the western edge of the northbound M1 carriageway, at the base of the western embankment.

    39. The first impact was at an airspeed of 113 knots CAS, with a rate of descent of between 8.5 feet/sec and 16 feet/sec. The pitch attitude was 13° nose up.

    40. The second and major impact occurred at a speed of between 80 and 100 knots, at an angle of approximately 16° below the horizontal and with the aircraft at a pitch attitude of between 9° and 14° nose down. The associatedd peak deceleration was of the order of 22 to 28g, predominantly longitudinal.

    41. In the second impact, the forward fuselage separated from the overwing sectiion of fuselage and the tail section buckled over, and to the right of, that section of fuselage just aft of the wing.

    42. The incidence of passenger fatality was highest where the floor had collapsed in the forward section of the passenger cabin and in the area just aft of the wing. The cabin floor and the passenger seating remained almost entirely intact within the overwing and tail sections.

    43. There was no major post impact fire, largely because the main landing gear legs and the engines separated from the wing without rupturing the wing fuel tanks. The separation of the landing gear legs was in accordance with their design. In the case of the engines, however, the separations occurred within the engine pylons themselves, leaving the fuse-pin bolts intact.


    44. Of the 8 crew and 118 passengers on board, all crew members survived but 39 passengers dies from impact injuries at the scene and a further 8 passengers dies later in hospital. A further 74 occupants were seriously injured.

    45. The decelerations generated in the second impact were greater than those specified in the Airworthiness Requirements to which the airframe and furnishings were designed and certificated. They were, however, within the physiological tolerance of a typical passenger.

    46. Passenger survivability was improved due to the passenger seats being of a design with impact tolerance in advance of the current regulatory requirements. This was most evident in the overwing and tail sections of the cabin, where the floor had remained intact.

    47. There is considerable potential for improving the survivability of passengers in this type of impact by improving the structural integrity of the cabin floor so as to retain the seats in their relative positions and by detail design improvements to the seats themselves.

    48. There is a need for a structured programme of research into alternative seating configurations, with particular emphasis on the provision of effective upper torso restraint or aft-facing seats.

    49. The injuries to the mother and child in seat 3F highlighted the advantages of infants being placed in child seats rather than in a loop-type supplementary belt.

    50. Although the overhead stowage bins met the appropriate Airworthiness Requirements for static loading, all but one of the 30 bins fell from their attachments, which did not withstand the dynamic loading conditions in this accident.

    51. Some of the doors on the overhead stowage bins opened during the last seconds of flight, demonstrating the need for some form of improved latching of the doors.

    (b) Cause

    The cause of the accident was that the operating crew shut down the No 2 engine after a fan blade had fractured in the No 1 engine. This engine subsequently suffered a major thrust loss due to secondary fan damage after power had been increased during the final approach to land.

    The following factors contributed to the incorrect response of the flight crew:
    1. The combination of heavy engine vibration, noise, shuddering and an associated smell of fire were outside their training and experience.
    2. They reacted to the initial engine problem prematurely and in a way that was contrary to their training.
    3. They did not assimilate the indications on the engine instrument display before they throttled back the No 2 engine.
    4. As the No 2 engine was throttled back, the noise and shuddering associated with the surging of the No 1 engine ceased, persuading them that they had correctly identified the defective engine.
    5. They were not informed of the flames which had emanated from the No 1 engine and which had been observed by many on board, including 3 cabin attendants in the aft cabin.

    4. Safety recommendations

    The following safety recommendations were made during the course of the investigation.

    Inspector of Air Accidents
    Air Accidents Investigation Branch
    Department of Transport

    August 1990

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    Peter B. Ladkin, 1999-02-08
    Last modification on 1999-06-15
    by Michael Blume