Clipper
Maid of the Seas
Flight 103 – Lockerbie, Scotland
Air Accidents Investigation Branch
Aircraft Accident Report No 2/90 (EW/C1094)
Report on the accident to Boeing 747-121. N739PA at Lockerbie,
Dumfriesshire, Scotland 12/21/88
Contents
- SYNOPSIS
- 1. FACTUAL INFORMATION
- 1.1 History of the flight
- 1.2 Injuries to persons
- 1.3 Damage to aircraft
- 1.4 Other damage
- 1.5 Personnel information
- 1.6 Aircraft information
- 1.7 Meteorological information
- 1.8 Aids to navigation
- 1.9 Communications
- 1.10 Aerodrome information
- 1.11 Flight recorders
- 1.12 Wreckage and impact information
- 1.13 Medical and pathological information
- 1.14 Fire
- 1.15 Survival aspects
- 1.16 Tests and research
- 1.17 Additional information
- 2. ANALYSIS
- 2.1 Introduction
- 2.2 Explosive destruction of the aircraft
- 2.3 Flight recorders
- 2.4 IED position within the aircraft
- 2.5 Engine evidence
- 2.6 Detachment of forward fuselage
- 2.7 Speed of initial disintegration
- 2.8 The manoeuvre following the explosion
- 2.9 Secondary disintegration
- 2.10 Impact speed of components
- 2.11 Sequence of disintegration
- 2.12 Explosive mechanisms and the structural
disintegration
- 2.13 Potential limitation of explosive damage
- 2.14 Summary
- 3. CONCLUSIONS
- 3.a Findings
- 3.b Cause
Operator: |
Pan American World Airways |
|
Aircraft Type: |
Boeing 747-121 |
|
Nationality: |
United States of America |
|
Registration: |
N 739 PA |
|
Place of Accident |
Lockerbie, Dumfries, Scotland |
|
|
Latitude |
55? 07′ N |
|
Longitude |
003? 21′ W |
Date and Time (UTC): |
21 December 1988 at 19.02:50 hrs |
|
|
All times in this report are UTC |
|
SYNOPSIS
The accident was notified to the Air Accidents
Investigation Branch at 19.40 hrs on the 21 December 1988 and the
investigation commenced that day. The members of the AAIB team are
listed at Appendix A.
The aircraft, Flight PA103 from London Heathrow to
New York, had been in level cruising flight at flight level 310 (31,000
feet) for approximately seven minutes when the last secondary radar
return was received just before 19.03 hrs. The radar then showed
multiple primary returns fanning out downwind. Major portions of the
wreckage of the aircraft fell on the town of Lockerbie with other large
parts landing in the countryside to the east of the town. Lighter debris
from the aircraft was strewn along two trails, the longest of which
extended some 130 kilometres to the east coast of England. Within a few
days items of wreckage were retrieved upon which forensic scientists
found conclusive evidence of a detonating high explosive. The airport
security and criminal aspects of the accident are the subject of a
separate investigation and are not covered in this report which
concentrates on the technical aspects of the disintegration of the
aircraft.
The report concludes that the detonation of an
improvised explosive device led directly to the destruction of the
aircraft with the loss of all 259 persons on board and 11 of the
residents of the town of Lockerbie. Five recommendations are made of
which four concern flight recorders, including the funding of a study to
devise methods of recording violent positive and negative pressure
pulses associated with explosions. The final recommendation is that
Airworthiness Authorities and aircraft manufacturers undertake a
systematic study with a view to identifying measures that might mitigate
the effects of explosive devices and improve the tolerance of the
aircraft’s structure and systems to explosive damage.
1. FACTUAL INFORMATION
1.1 History of the Flight
Boeing 747, N739PA, arrived at London Heathrow
Airport from San Francisco and parked on stand Kilo 14, to the
south-east of Terminal 3. Many of the passengers for this aircraft had
arrived at Heathrow from Frankfurt, West Germany on a Boeing 727, which
was positioned on stand Kilo 16, next to N739PA. These passengers were
transferred with their baggage to N739PA which was to operate the
scheduled Flight PA103 to New York Kennedy. Passengers from other
flights also joined Flight PA103 at Heathrow. After a 6 hour turnround,
Flight PA103 was pushed back from the stand at 18.04 hrs and was cleared
to taxy on the inner taxiway to runway 27R. The only relevant Notam
warned of work in progress on the outer taxiway. The departure was
unremarkable.
Flight PA103 took-off at 18.25 hrs. As it was
approaching the Burnham VOR it took up a radar heading of 350? and flew
below the Bovingdon holding point at 6000 feet. It was then cleared to
climb initially to flight level (FL) 120 and subsequently to FL 310. The
aircraft levelled off at FL 310 north west of Pole Hill VOR at 18.56
hrs. Approximately 7 minutes later, Shanwick Oceanic Control transmitted
the aircraft’s oceanic clearance but this transmission was not
acknowledged. The secondary radar return from Flight PA103 disappeared
from the radar screen during this transmission. Multiple primary radar
returns were then seen fanning out downwind for a considerable distance.
Debris from the aircraft was strewn along two trails, one of which
extended some 130 km to the east coast of England. The upper winds were
between 250? and 260? and decreased in strength from 115 kt at FL 320 to
60 kt at FL 100 and 15 to 20 kt at the surface.
Two major portions of the wreckage of the aircraft
fell on the town of Lockerbie; other large parts, including the flight
deck and forward fuselage section, landed in the countryside to the east
of the town. Residents of Lockerbie reported that, shortly after 19.00
hrs, there was a rumbling noise like thunder which rapidly increased to
deafening proportions like the roar of a jet engine under power. The
noise appeared to come from a meteor-like object which was trailing
flame and came down in the north-eastern part of the town. A larger,
dark, delta shaped object, resembling an aircraft wing, landed at about
the same time in the Sherwood area of the town. The delta shaped object
was not on fire while in the air, however, a very large fireball ensued
which was of short duration and carried large amounts of debris into the
air, the lighter particles being deposited several miles downwind. Other
less well defined objects were seen to land in the area.
1.2 Injuries to persons
Injuries |
Crew |
Passengers |
Others |
Fatal |
16 |
243 |
11 |
Serious |
– |
– |
2 |
Minor/None |
– |
– |
3 |
1.3 Damage to aircraft
The aircraft was destroyed
1.4 Other damage
The wings impacted at the southern edge of
Lockerbie, producing a crater whose volume, calculated from a
photogrammetric survey, was approximately 560 cubic metres. The weight
of material displaced by the wing impact was estimated to be well in
excess of 1500 tonnes. The wing impact created a fireball, setting fire
to neighbouring houses and carrying aloft debris which was then blown
downwind for several miles. It was subsequently established that
domestic properties had been so seriously damaged as a result of fire
and/or impact that 21 had to be demolished and an even greater number of
homes required substantial repairs. Major portions of the aircraft,
including the engines, also landed on the town of Lockerbie and other
large parts, including the flight deck and forward fuselage section,
landed in the countryside to the east of the town. Lighter debris from
the aircraft was strewn as far as the east coast of England over a
distance of 130 kilometres.
1.5 Personnel information
1.5.1 |
Commander: |
Male, aged 55 years |
|
Licence: |
USA Airline Transport Pilot’s
Licence |
|
Aircraft ratings: |
Boeing 747, Boeing 707, Boeing
720, Lockheed L1011 and Douglas DC3 |
|
Medical Certificate: |
Class 1,valid to April 1989, with
the limitation that the holder shall wear lenses that correct for
distant vision and possess glasses that correct for near vision |
Flying experience: |
|
Total all types: |
10,910 hours |
Total on type: |
4,107 hours |
Total last 28 days |
82 hours |
Duty time: |
Commensurate with company
requirements |
Last base check: |
11 November 1988 |
Last route check: |
30 June 1988 |
Last emergencies check: |
8 November 1988 |
1.5.2 |
Co-pilot: |
Male, aged 52 years |
|
Licence: |
USA Airline Transport Pilot’s
Licence |
|
Aircraft ratings: |
Boeing 747, Boeing 707, Boeing
727 |
|
Medical Certificate: |
Class 1, valid to April 1989,
with the limitation that the holder shall possess correcting glasses
for near vision |
|
Flying experience: |
|
|
Total all types: |
11,855 hours |
|
Total on type: |
5,517 hours |
|
Total last 28 days: |
51 hours |
|
Duty time: |
Commensurate with company
requirements |
|
Last base check: |
30 November 1988 |
|
Last route check: |
Not required |
|
Last emergencies check: |
27 November 1988 |
1.5.3 |
Flight Engineer: |
Male, aged 46 years |
|
Licence: |
USA Flight Engineer’s Licence |
|
Aircraft ratings: |
Turbojet |
|
Medical certificate: |
Class 2, valid to June 1989, with
the limitation that the holder shall wear correcting glasses for
near vision |
|
Flying experience: |
|
|
Total all types: |
8,068 hours |
|
Total on type: |
487 hours |
|
Total last 28 days: |
53 hours |
|
Duty time: |
Commensurate with company
requirements |
|
Last base check: |
30 October 1988 |
|
Last route check: |
Not required |
|
Last emergencies check: |
27 October 1988 |
1.5.4 Flight Attendants: There were 13 Flight
Attendants on the aircraft, all of whom met company proficiency and
medical requirements
1.6 Aircraft information
1.6.1 |
Leading particulars |
|
|
Aircraft type: |
Boeing 747-121 |
|
Constructor’s serial number: |
19646 |
|
Engines: |
4 Pratt and Whitney JT9D-7A
turbofan |
1.6.2 General description
The Boeing 747 aircraft, registration N739PA, was
a conventionally designed long range transport aeroplane. A diagram
showing the general arrangement is shown at Appendix B, Figure B-1
together with the principal dimensions of the aircraft.
The fuselage of the aircraft type was of
approximately circular section over most of its length, with the forward
fuselage having a diameter of 21? feet where the cross-section was
constant. The pressurised section of the fuselage (which included the
forward and aft cargo holds) had an overall length of 190 feet,
extending from the nose to a point just forward of the tailplane. In
normal cruising flight the service pressure differential was at the
maximum value of 8.9 pounds per square inch. The fuselage was of
conventional skin, stringer and frame construction, riveted throughout,
generally using countersunk flush riveting for the skin panels. The
fuselage frames were spaced at 20 inch intervals and given the same
numbers as their stations, defined in terms of the distance in inches
from the datum point close to the nose of the aircraft [Appendix B,
Figure B-2]. The skin panels were joined using vertical butt joints and
horizontal lap joints. The horizontal lap joints used three rows of
rivets together with a cold bonded adhesive.
Accommodation within the aircraft was
predominately on the main deck, which extended throughout the whole
length of the pressurised compartment. A separate upper deck was
incorporated in the forward part of the aircraft. This upper deck was
reached by means of a spiral staircase from the main deck and
incorporated the flight crew compartment together with additional
passenger accommodation. The cross-section of the forward fuselage
differed considerably from the near circular section of the remainder of
the aircraft, incorporating an additional smaller radius arc above the
upper deck section joined to the main circular arc of the lower cabin
portion by elements of straight fuselage frames and flat skin.
In order to preserve the correct shape of the
aircraft under pressurisation loading, the straight portions of the
fuselage frames in the region of the upper deck floor and above it were
required to be much stiffer than the frame portions lower down in the
aircraft. These straight sections were therefore of very much more
substantial construction than most of the curved sections of frames
lower down and further back in the fuselage. There was considerable
variation in the gauge of the fuselage skin at various locations in the
forward fuselage of the aircraft.
The fuselage structure of N739PA differed from
that of the majority of Boeing 747 aircraft in that it had been modified
to carry special purpose freight containers on the main deck, in place
of seats. This was known as the Civil Reserve Air Fleet (CRAF)
modification and enabled the aircraft to be quickly converted for
carriage of military freight containers on the main deck during times of
national emergency. The effect of this modification on the structure of
the fuselage was mainly to replace the existing main deck floor beams
with beams of more substantial cross-section than those generally found
in passenger carrying Boeing 747 aircraft. A large side loading door,
generally known as the CRAF door, was also incorporated on the left side
of the main deck aft of the wing.
Below the main deck, in common with other Boeing
747 aircraft, were a number of additional compartments, the largest of
which were the forward and aft freight holds used for the storage of
cargo and baggage in standard air-transportable containers. These
containers were placed within the aircraft hold by means of a freight
handling system and were carried on a system of rails approximately 2
feet above the outer skin at the bottom of the aircraft, there being no
continuous floor, as such, below these baggage containers. The forward
freight compartment had a length of approximately 40 feet and a depth of
approximately 6 feet. The containers were loaded into the forward hold
through a large cargo door on the right side of the aircraft.
1.6.3 Internal fuselage cavities
Because of the conventional skin, frame and
stringer type of construction, common to all large public transport
aircraft, the fuselage was effectively divided into a series of ‘bays’.
Each bay, comprising two adjacent fuselage frames and the structure
between them, provided, in effect, a series of interlinking cavities
bounded by the frames, floor beams, fuselage skins and cabin floor
panels etc. The principal cavities thus formed were:
(i) |
A semi-circular cavity formed in
between the fuselage frames in the lower lobe of the hull, i.e. from
the crease beam (at cabin floor level) on one side down to the belly
beneath the containers and up to the opposite crease beam, bounded
by the fuselage skin on the outside and the containers/cargo liner
on the inside [Appendix B, Figure B-3, detail A]. |
|
|
(ii) |
A horizontal cavity between the
main cabin floor beams, the cabin floor panels and the cargo bay
liner. This extended the full width of the fuselage and linked the
upper ends of the lower lobe cavity [Appendix B, Figure B-3, detail
B]. |
|
|
(iii) |
A narrow vertical cavity between
the two containers [Appendix B, Figure B-3, detail C]. |
|
|
(iv) |
A further narrow cavity around
the outside of the two containers, between the container skins and
the cargo bay liner, communicating with the lower lobe cavity
[Appendix B, Figure B-3, detail D]. |
|
|
(v) |
A continuation of the
semi-circular cavity into the space behind the cabin wall liner
[Appendix B, Figure B-3, detail E]. This space was restricted
somewhat by the presence of the window assembly, but nevertheless
provided a continuous cavity extending upwards to the level of the
upper deck floor. Forward of station 740, this cavity was
effectively terminated at its upper end by the presence of
diaphragms which formed extensions of the upper deck floor panels;
aft of station 740, the cavity communicated with the ceiling space
and the cavity in the fuselage crown aft of the upper deck. |
All of these cavities were repeated at each
fuselage bay (formed between pairs of fuselage frames), and all of the
cavities in a given bay were linked together, principally at the crease
beam area [Appendix B, Figure B-3, region F]. Furthermore, each of the
set of bay cavities was linked with the next by the longitudinal
cavities formed between the cargo hold liner and the outer hull, just
below the crease beam [Appendix B, Figure B-3, detail F]; i.e. this
cavity formed a manifold linking together each of the bays within the
cargo hold.
The main passenger cabin formed a large chamber
which communicated directly with each of the sub floor bays, and also
with the longitudinal manifold cavity, via the air conditioning and
cabin/cargo bay de-pressurisation vent passages in the crease beam area.
(It should be noted that a similar communication did not exist between
the upper and lower cabins because there were no air conditioning/depressurisation
passages to bypass the upper deck floor.)
1.6.4 Aircraft weight and centre of gravity
The aircraft was loaded within its permitted
centre of gravity limits as follows:
Loading: |
lb |
kg |
Operating empty weight |
366,228 |
166,120 |
Additional crew |
130 |
59 |
243 passengers (1) |
40,324 |
18,291 |
Load in compartments: |
|
|
1 |
11,616 |
5,269 |
2 |
20,039 |
9,090 |
3 |
15,057 |
6,830 |
4 |
17,196 |
7,800 |
5 |
2,544 |
1,154 |
Total in compartments (2) |
66,452 |
30,143 |
Total traffic load |
106,776 |
48,434 |
Zero fuel weight |
472,156 |
214,554 |
Fuel (Take-off) |
239,997 |
108,862 |
Actual take-off weight(4) |
713,002 |
323,416 |
Maximum take-off weight |
733,992 |
332,937 |
Note 1:
Calculated at standard weights and including cabin
baggage.
Note 2:
Despatch information stated that the cargo did not
include dangerous goods, perishable cargo, live animals or known
security exceptions.
1.6.5 Maintenance details
N739PA first flew in 1970 and spent its whole
service life in the hands of Pan American World Airways Incorporated.
Its Certificate of Airworthiness was issued on 12 February 1970 and
remained in force until the time of the accident, at which time the
aircraft had completed a total of 72,464 hours flying and 16,497 flight
cycles. Details of the last 4 maintenance checks carried out during the
aircraft’s life are shown below:
DATE |
SERVICE |
HOURS |
CYCLES |
27 Sept 88 |
C Check (Interior upgrade) |
71,502 |
16,347 |
2 Nov 88 |
B Service Check |
71,919 |
16,406 |
27 Nov 88 |
Base 1 |
72,210 |
16,454 |
13 Dec 88 |
Base 2 |
72,374 |
16,481 |
The CRAF modification programme was undertaken in
September 1987. At the same time a series of modifications to the
forward fuselage from the nose back to station 520 (Section 41) were
carried out to enable the aircraft to continue in service without a
continuing requirement for structural inspections in certain areas.
All Airworthiness Directives relating to the
Boeing 747 fuselage structure between stations 500 and 1000 have been
reviewed and their applicability to this aircraft checked. In addition,
Service Bulletins relating to the structure in this area were also
reviewed. The applicable Service Bulletins, some of which implement the
Airworthiness Directives are listed below together with their subjects.
The dates, total aircraft times and total aircraft cycles at which each
relevant inspection was last carried out have been reviewed and their
status on aircraft N739PA at the time of the accident has been
established.
N739PA Service Bulletin compliance:
SB 53-2064 |
Front Spar Pressure Bulkhead
Chord Reinforcement and Drag Splice Fitting Rework. |
|
Modification accomplished on 6
July 1974. |
|
Post-modification repetitive
inspection IAW (in accordance with) AD 84-18-06 last accomplished on
19 November 1985 at 62,030 TAT hours (Total Aircraft Time) and
14,768 TAC (Total Aircraft Cycles). |
SB 53-2088 |
Frame to Tension Tie Joint
Modification – BS760 to 780. |
|
Repetitive inspection IAW AD
84-19-01 last accomplished on 19 June 1985 at 60,153 hours TAT and
14,436 TAC. |
SB 53-2200 |
Lower Cargo Doorway Lower Sill
Truss and Latch Support Fitting Inspection Repair and Replacement. |
|
Repetitive inspection IAW AD
79-17-02 R2 last accomplished 2 November 1988 at 71,919 hours TAT
and 16,406 TAC. |
SB 53-2234 |
Fuselage – Auxiliary Structure –
Main Deck Floor – BS 480 Floor Beam Upper Chord Modification. |
|
Repetitive inspection per SB
53A2263 IAW AD 86-23-06 last accomplished on 26 September 1987 at
67,376 hours TAT and 15,680 TAC. |
SB 53-2237 |
Fuselage – Main Frame – BS 540
thru 760 and 1820 thru 1900 Frame Inspection and Reinforcement. |
|
Repetitive inspection IAW AD
86-18-01 last accomplished on 27 February 1987 at 67,088 hours TAT
and 15,627 TAC. |
SB 53-2267 |
Fuselage – Skin – Lower Body
Longitudinal Skin Lap Joint and Adjacent Body Frame Inspection and
Repair. |
|
Terminating modification
accomplished 100% under wing-to-body fairings and approximately 80%
in forward and aft fuselage sections on 26 September 1987 at 67,376
hours TAT and 15,680 TAC. |
|
Repetitive inspection of
unmodified lap joints IAW AD 86-09-07 R1 last accomplished on 18
August 1988 at 71,043 hours TAT and 16,273 TAC. |
SB 53A2303 |
Fuselage – Nose Section – station
400 to 520 Stringer 6 Skin Lap Splice Inspection, Repair and
Modification. |
|
Repetitive inspection IAW AD
89-05-03 last accomplished on 26 September 1987 at 67,376 hours TAT
and 15,680 TAC. |
This documentation, when viewed together with the
detailed content of the above service bulletins, shows the aircraft to
have been in compliance with the requirements laid down in each of those
bulletins. Some maintenance items were outstanding at the time the
aircraft was despatched on the last flight, however, none of these items
relate to the structure of the aircraft and none had any relevance to
the accident.
1.7 Meteorological Information
1.7.1 General weather conditions
An aftercast of the general weather conditions in
the area of Lockerbie at about 19.00 hrs was obtained from the
Meteorological Office, Bracknell. The synoptic situation included a warm
sector covering northern England and most of Scotland with a cold front
some 200 nautical miles to the west of the area moving eastwards at
about 35 knots. The weather consisted of intermittent rain or showers.
The cloud consisted of 4 to 6 oktas of stratocumulus based at 2,200 feet
with 2 oktas of altocumulus between 15,000 and 18,000 feet. Visibility
was over 15 kilometers and the freezing level was at 8,500 feet with a
sub-zero layer between 4,000 and 5,200 feet.
1.7.2 Winds
There was a weakening jet stream of around 115
knots above Flight Level 310. From examination of the wind profile (see
below), there appeared to be insufficient shear both vertically and
horizontally to produce any clear air turbulence but there may have been
some light turbulence.
Flight Level |
Wind |
320 |
260?/115 knots |
300 |
260?/ 90 knots |
240 |
250?/ 80 knots |
180 |
260?/ 60 knots |
100 |
250?/ 60 knots |
050 |
260?/ 40 knots |
Surface |
240?/ 15 to 20 gusting 25 to 30
knots |
1.8 Aids to navigation
Not relevant.
1.9 Communications
The aircraft communicated normally on London
Heathrow aerodrome, London control and Scottish control frequencies.
Tape recordings and transcripts of all radio telephone (RTF)
communications on these frequencies were available.
At 18.58 hrs the aircraft established two-way
radio contact with Shanwick Oceanic Area Control on frequency 123.95
MHz. At 19.02:44 hrs the clearance delivery officer at Shanwick
transmitted to the aircraft its oceanic route clearance. The aircraft
did not acknowledge this message and made no subsequent transmission.
1.9.1 ATC recording replay
Scottish Air Traffic Control provided copy tapes
with time injection for both Shanwick and Scottish ATC frequencies. The
source of the time injection on the tapes was derived from the British
Telecom "TIM" signal.
The tapes were replayed and the time signals
corrected for errors at the time of the tape mounting.
1.9.2 Analysis of ATC tape recordings
From the cockpit voice recorder (CVR) tape it was
known that Shanwick was transmitting Flight PA103’s transatlantic
clearance when the CVR stopped. By synchronising the Shanwick tape and
the CVR it was possible to establish that a loud sound was heard on the
CVR cockpit area microphone (CAM) channel at 19.02:50 hrs ?1 second.
As the Shanwick controller continued to transmit
Flight PA103’s clearance instructions through the initial destruction of
the aircraft it would not have been possible for a distress call to be
received from N739PA on the Shanwick frequency. The Scottish frequency
tape recording was listened to from 19.02 hrs until 19.05 hrs for any
unexplained sounds indicating an attempt at a distress call but none was
heard.
A detailed examination and analysis of the ATC
recording together with the flight recorder, radar, and seismic
recordings is contained in Appendix C.
1.10 Aerodrome information
Not relevant
1.11 Flight recorders
The Digital Flight Data Recorder (DFDR) and the
Cockpit Voice Recorder (CVR) were found close together at UK Ordnance
Survey (OS) Grid Reference 146819, just to the east of Lockerbie, and
recovered approximately 15 hours after the accident. Both recorders were
taken directly to AAIB Farnborough for replay. Details of the
examination and analysis of the flight recorders together with the
radar, ATC and seismic recordings are contained in Appendix C.
1.11.1 Digital flight data recorder
The flight data recorder installation conformed to
ARINC 573B standard with a Lockheed Model 209 DFDR receiving data from a
Teledyne Controls Flight Data Acquisition Unit (FDAU). The system
recorded 22 parameters and 27 discrete (event) parameters. The flight
recorder control panel was located in the flight deck overhead panel.
The FDAU was in the main equipment centre at the front end of the
forward hold and the flight recorder was mounted in the aft equipment
centre.
Decoding and reduction of the data from the
accident flight showed that no abnormal behaviour of the data sensors
had been recorded and that the recorder had simply stopped at 19.02:50
hrs ?1 second.
1.11.2 Cockpit voice recorder
The aircraft was equipped with a 30 minute
duration 4 track Fairchild Model A100 CVR, and a Fairchild model A152
cockpit area microphone (CAM). The CVR control panel containing the CAM
was located in the overhead panel on the flight deck and the recorder
itself was mounted in the aft equipment centre.
The channel allocation was as follows:-
Channel 1 |
Flight Engineer’s RTF. |
Channel 2 |
Co-Pilot’s RTF. |
Channel 3 |
Pilot’s RTF. |
Channel 4 |
Cockpit Area Microphone. |
The erase facility within the CVR was not
functioning satisfactorily and low level communications from earlier
recordings were audible on the RTF channels. The CAM channel was
particularly noisy, probably due to the combination of the inherently
noisy flight deck of the B747-100 in the climb and distortion from the
incomplete erasure of the previous recordings. On two occasions the crew
had difficulty understanding ATC, possibly indicating high flight deck
noise levels. There was a low frequency sound present at irregular
intervals on the CAM track but the source of this sound could not be
identified and could have been of either acoustic or electrical origin.
The CVR tape was listened to for its full duration
and there was no indication of anything abnormal with the aircraft, or
unusual crew behaviour. The tape record ended, at 19.02:50 hrs ?1
second, with a sudden loud sound on the CAM channel followed almost
immediately by the cessation of recording whilst the crew were copying
their transatlantic clearance from Shanwick ATC.
1.12 Wreckage and impact information
1.12.1 General distribution of wreckage in the
field
The complete wing primary structure, incorporating
the centre section, impacted at the southern edge of Lockerbie. Major
portions of the aircraft, including the engines, also landed in the
town. Large portions of the aircraft fell in the countryside to the east
of the town and lighter debris was strewn to the east as far as the
North Sea. The wreckage was distributed in two trails which became known
as the northern and southern trails respectively and these are shown in
Appendix B, Figure B-4. A computer database of approximately 1200
significant items of wreckage was compiled and included a brief
description of each item and the location where it was found
Appendix B, Figures B-5 to B-8 shows photographs
of a model of the aircraft on which the fracture lines forming the
boundaries of the separate items of structure have been marked. The
model is colour coded to illustrate the way in which the wreckage was
distributed between the town of Lockerbie and the northern and southern
trails.
1.12.1.1 The crater
The aircraft wing impacted in the Sherwood
Crescent area of the town leaving a crater approximately 47 metres (155
feet) long with a volume calculated to be 560 cubic metres.
The projected distance, measured parallel from one
leading edge to the other wing tip, of the Boeing 747-100 was
approximately 143 feet, whereas the span is known to be 196 feet. This
suggests that impact took place with the wing structure yawed. Although
the depth of the crater varied from one end to the other, its widest
part was clearly towards the western end suggesting that the wing
structure impacted whilst orientated with its root and centre section to
the west.
The work carried out at the main crater was
limited to assessing the general nature of its contents. The total
absence of debris from the wing primary structure found remote from the
crater confirmed the initial impression that the complete wing box
structure had been present at the main impact.
The items of wreckage recovered from or near the
crater are coloured grey on the model at Appendix B, Figures B-5 to B-8.
1.12.1.2 The Rosebank Crescent site
A 60 feet long section of fuselage between frame
1241 (the rear spar attachment) and frame 1960 (level with the rear edge
of the CRAF cargo door) fell into a housing estate at Rosebank Crescent,
just over 600 metres from the crater. This section of the fuselage was
that situated immediately aft of the wing, and adjoined the wing and
fuselage remains which produced the crater. It is colour coded yellow on
the model at Appendix B, Figures B-5 to B-8. All fuselage skin structure
above floor level was missing except for the following items:
Section containing 3 windows between door 4L and
CRAF door;
The CRAF door itself (latched) apart from the top
area containing the hinge;
Window belt containing 8 windows aft of 4R door
aperture
Window belt containing 3 windows forward of 4R door
aperture;
Door 4R.
Other items found in the wreckage included both
body landing gears, the right wing landing gear, the left and right
landing gear support beams and the cargo door (frames 1800-1920) which
was latched. A number of pallets, luggage containers and their contents
were also recovered from this site.
1.12.1.3 Forward fuselage and flight deck section.
The complete fuselage forward of approximately
station 480 (left side) to station 380 (right side) and incorporating
the flight deck and nose landing gear was found as a single piece
[Appendix B, Figure B-9] in a field approximately 4 km miles east of
Lockerbie at OS Grid Reference 174808. It was evident from the nature of
the impact damage and the ground marks that it had fallen almost flat on
its left side but with a slight nose-down attitude and with no
discernible horizontal velocity. The impact had caused almost complete
crushing of the structure on the left side. The radome and right nose
landing gear door had detached in the air and were recovered in the
southern trail.
Examination of the torn edges of the fuselage skin
did not indicate the presence of any pre-existing structural or material
defects which could have accounted for the separation of this section of
the fuselage. Equally so, there were no signs of explosive blast damage
or sooting evident on any part of the structure or the interior
fittings. It was noted however that a heavy, semi-eliptical scuff mark
was present on the lower right side of the fuselage at approximately
station 360. This was later matched to the intake profile of the No 3
engine.
The status of the controls and switches on the
flight deck was consistent with normal operation in cruising flight.
There were no indications that the crew had attempted to react to rapid
decompression or loss of control or that any emergency preparations had
been actioned prior to the catastrophic disintegration.
1.12.1.4 Northern trail
The northern trail was seen to be narrow and
clearly defined, to emanate from a point very close to the main impact
crater and to be orientated in a direction which agreed closely with the
mean wind aftercast for the height band from sea level to 20,000 ft.
Also at the western end of the northern trail were the lower rear
fuselage at Rosebank Crescent, and the group of Nos. 1, 2 and 4 engines
which fell in Lockerbie.
The trail contained items of structure distributed
throughout its length, from the area slightly east of the crater, to a
point approximately 16 km east, beyond which only items of low weight /
high drag such as insulation, interior trim, paper etc, were found. For
all practical purposes this trail ended at a range of 25 km.
The northern trail contained mainly wreckage from
the rear fuselage, fin and the inner regions of both tailplanes together
with structure and skin from the upper half of the fuselage forward to
approximately the wing mid-chord position. A number of items from the
wing were also found in the northern trail, including all 3 starboard
Kreuger flaps, most of the remains of the port Kreuger flaps together
with sections of their leading edge attachment structures, one portion
of outboard aileron approximately 10 feet long, the aft ends of the
flap-track fairings (one with a slide raft wrapped around it), and
fragments of glass reinforced plastic honeycombe structure believed to
be from the flap system, i.e. fore-flaps, aft-flaps, mid-flaps or
adjacent fairings. In addition, a number of pieces of the engine
cowlings and both HF antennae (situated projecting aft from the
wing-tips) were found in this trail.
All items recovered from the northern trail, with
the exception of the wing, engines, and lower rear fuselage in Rosebank
Crescent, are coloured red on the model of the aircraft in Appendix B,
Figures B-5 to B-8.
1.12.1.5 Southern trail
The southern trail was easily defined, except
within 12 km of Lockerbie where it tended to merge with the northern
trail. Further east, it extended across southern Scotland and northern
England, essentially in a straight band as far as the North Sea. Most of
the significant items of wreckage were found in this trail within a
range of 30 km from the main impact crater. Items recovered from the
southern trail are coloured green on the model of the aircraft at
Appendix B, Figures B-5 to B-8.
The trail contained numerous large items from the
forward fuselage. The flight deck and nose of the aircraft fell in the
curved part of this trail close to Lockerbie. Fragments of the whole of
the left tailplane and the outboard portion of the right tailplane were
distributed almost entirely throughout the southern trail. Between 21
and 27 km east of the main impact point (either side of Langholm)
substantial sections of tailplane skin were found, some bearing
distinctive signs of contact with debris moving outwards and backwards
relative to the fuselage. Also found in this area were numerous isolated
sections of fuselage frame, clearly originating from the crown region
above the forward upper deck.
1.12.1.6 Datum line
All grid references relating to items bearing
actual explosive evidence, together with those attached to heavily
distorted items found to originate immediately adjacent to them on the
structure, were plotted on an Ordnance Survey (OS) chart. These
references, 11 in total, were all found to be distributed evenly about a
mean line orientated 079?(Grid) within the southern trail and were
spread over a distance of 12 km. The distance of each reference from the
line was measured in a direction parallel to the aircraft’s track and
all were found to be within 500 metres of the line, with 50% of them
being within 250 metres of the line. This line is referred to as the
datum line and is shown in Appendix B, Figure B-4.
1.12.1.7 Distribution of wreckage within the
southern trail
North of the datum line and parallel to it were
drawn a series of lines at distances of 250, 300, 600 and 900 metres
respectively from the line, again measured in a direction parallel to
the aircraft’s track. The positions on the aircraft structure of
specific items of wreckage, for which grid references were known with a
high degree of confidence, within the bands formed between these lines,
are shown in Appendix B, Figures B-10 to 13. In addition, a separate
assessment of the grid references of tailplane and elevator wreckage
established that these items were distributed evenly about the 600 metre
line.
1.12.1.8 Area between trails
Immediately east of the crater, the southern trail
converged with the northern trail such that, to an easterly distance of
approximately 5 km, considerable wreckage existed which could have
formed part of either trail. Further east, between 6 and 11 km from the
crater, a small number of sections and fragments of the fin had fallen
outside the southern boundary of the northern trail. Beyond this a large
area existed between the trails in which there was no wreckage.
1.12.2 Examination of wreckage at CAD Longtown
The debris from all areas was recovered by the
Royal Air Force to the Army Central Ammunition Depot Longtown, about 20
miles from Lockerbie. Approximately 90% of the hull wreckage was
successfully recovered, identified, and laid out on the floor in a
two-dimensional reconstruction [Appendix B, Figure B-14]. Baggage
container material was incorporated into a full three-dimensional
reconstruction. Items of wreckage added to the reconstructions was given
a reference number and recorded on a computer database together with a
brief description of the item and the location where it was found.
1.12.2.1 Fuselage
The reconstruction revealed the presence of damage
consistent with an explosion on the lower fuselage left side in the
forward cargo bay area. A small region of structure bounded
approximately by frames 700 & 720 and stringers 38L & 40L, had clearly
been shattered and blasted through by material exhausting directly from
an explosion centred immediately inboard of this location. The material
from this area, hereafter referred to as the ‘shatter zone’, was mostly
reduced to very small fragments, only a few of which were recovered,
including a strip of two skins [Appendix B, Figure B-15] forming part of
the lap joint at the stringer 39L position.
Surrounding the shatter zone were a series of much
larger panels of torn fuselage skin which formed a ‘star-burst’ fracture
pattern around the shatter zone. Where these panels formed the boundary
of the shatter zone, the metal in the immediate locality was ragged,
heavily distorted, and the inner surfaces were pitted and sooted –
rather as if a very large shotgun had been fired at the inner surface of
the fuselage at close range. In contrast, the star-burst fractures,
outside the boundary of the shatter zone, displayed evidence of more
typical overload tearing, though some tears appeared to be rapid and, in
the area below the missing panels, were multi-branched. These
surrounding skin panels were moderately sooted in the regions adjacent
to the shatter zone, but otherwise were lightly sooted or free of soot
altogether. (Forensic analysis of the soot deposits on frame and skin
material from this area confirmed the presence of explosive residues.)
All of these skin panels had pulled away from the supporting structure
and had been bent and torn in a manner which indicated that, as well as
fracturing in the star burst pattern, they had also petalled outwards
producing characteristic, tight curling of the sheet material.
Sections of frames 700 and 720 from the area of
the explosion were also recovered and identified. Attached to frame 720
were the remnants of a section of the aluminium baggage container (side)
guide rail, which was heavily distorted and displayed deep pitting
together with very heavy sooting, indicating that it had been very close
to the explosive charge. The pattern of distortion and damage on the
frames and guide rail segment matched the overall pattern of damage
observed on the skins.
The remainder of the structure forming the cargo
deck and lower hull was, generally, more randomly distorted and did not
display the clear indications of explosive processes which were evident
on the skin panels and frames nearer the focus of the explosion.
Nevertheless, the overall pattern of damage was consistent with the
propagation of explosive pressure fronts away from the focal area
inboard of the shatter zone. This was particularly evident in the
fracture and bending characteristics of several of the fuselage frames
ahead of, and behind station 700.
The whole of the two-dimensional fuselage
reconstruction was examined for general evidence of the mode of
disintegration and for signs of localised damage, including overpressure
damage and pre-existing damage such as corrosion or fatigue. There was
some evidence of corrosion and dis-bonding at the cold-bond lap joints
in the fuselage. However, the corrosion was relatively light and would
not have compromised significantly the static strength of the airframe.
Certainly, there was no evidence to suggest that corrosion had affected
the mode of disintegration, either in the area of the explosion or at
areas more remote. Similarly, there were no indications of fatigue
damage except for one very small region of fatigue, involving a single
crack less than 3 inches long, which was remote from the bomb location.
This crack was not in a critical area and had not coincided with a
fracture path.
No evidence of overpressure fracture or distortion
was found at the rear pressure bulkhead. Some suggestion of ‘quilting’
or ‘pillowing’ of skin panels between stringers and frames, indicative
of localised overpressure, was evident on the skin panels attached to
the larger segments of lower fuselage wreckage aft of the blast area. In
addition, the mode of failure of the butt joint at station 520 suggested
that there had been a rapid overpressure load in this area, causing the
fastener heads to ‘pop’ in the region of stringers 13L to 16L, rather
than producing shear in the fasteners. Further evidence of localised
overpressure damage remote from the source of the explosion was found
during the full three-dimensional reconstruction, detailed later in
paragraph 1.12.3.2.
An attempt was made to analyse the fractures, to
determine the direction and sequence of failure as the fractures
propagated away from the region of the explosion. It was found that the
directions of most of the fractures close to the explosion could be
determined from an analysis of the fracture surfaces and other features,
such as rivet and rivet hole distortions. However, it was apparent that
beyond the boundary of the petalled region, the disintegration process
had involved multiple fractures taking place simultaneously – extremely
complex parallel processes which made the sequencing of events not
amenable to conventional analysis.
1.12.2.2 Wing structure and adjacent fuselage area
On completion of the initial layout at Longtown it
became evident that, in the area from station 1000 to approximately
station 1240 the only identifiable fuselage structure consisted of
elements of fuselage skin, stringers and frames from above the cabin
window belts. The wreckage from in and around the crater was therefore
sifted to establish more accurately what sections of the aircraft had
produced the crater. All of the material was highly fragmented, but it
was confirmed that the material comprised mostly wing structure, with a
few fragments of fuselage sidewall and passenger seats. The badly burnt
state of these fragments made it clear that they were recovered from the
area of the main impact crater, the only scene of significant ground
fire. Amongst these items a number of cabin window forgings were
recovered with sections of thick horizontal panelling attached having a
length equivalent to the normal window spacing/frame pitch. This
arrangement, with skins of this thickness, is unique to the area from
station 1100 to 1260. It is therefore reasonable to assume that these
fragments formed parts of the missing cabin sides from station 1000 to
station 1260, which must have remained attached to the wing centre
section at the time of its impact. Because of the high degree of
fragmentation and the relative insignificance of the wing in terms of
the overall explosive damage pattern, a reconstruction of the wing
material was not undertaken. The sections of the aircraft which went
into the crater are colour coded grey in Appendix B, Figures B-5 to B-8.
1.12.2.3 Fin and aft section of fuselage
Examination of the structure of the fin revealed
evidence of in-flight damage to the leading edge caused by the impact of
structure or cabin contents. This damage was not severe or extensive and
the general break-up of the fin did not suggest either a single readily
defined loading direction, or break-up due to the effects of leading
edge impact. A few items of fin debris were found between the northern
and southern trails.
A number of sections of fuselage frame found in
the northern trail exhibited evidence of plastic deformation of skin
attachment cleats and tensile overload failure of the attachment rivets.
This damage was consistent with that which would occur if the skin had
been locally subjected to a high loading in a direction normal to its
plane. Although this was suggestive of an internal overpressure
condition, the rear fuselage revealed no other evidence to support this
possibility. Examination of areas of the forward fuselage known to have
been subjected to high blast overpressures revealed no comparable
evidence of plastic deformation in the skin attachment cleats or rivets,
most skin attachment failures appearing to have been rapid.
Calculations made on the effects of internal
pressure generated by an open ended fuselage descending at the highest
speed likely to have been experienced revealed that this could not
generate an internal pressure approaching that necessary to cause
failure in an intact cabin structure.
1.12.2.4 Baggage containers
During the wreckage recovery operation it became
apparent that some items, identified as parts of baggage containers,
exhibited damage consistent with being close to a detonating high
explosive. It was therefore decided to segregate identifiable container
parts and reconstruct any that showed evidence of explosive damage. It
was evident, from the main wreckage layout, that the explosion had
occurred in the forward cargo hold and, although all baggage container
wreckage was examined, only items from this area which showed the
relevant characteristics were considered for the reconstruction.
Discrimination between forward and rear cargo hold containers was
relatively straightforward as the rear cargo hold wreckage was almost
entirely confined to Lockerbie, whilst that from the forward hold was
scattered along the southern wreckage trail.
All immediately identifiable parts of the forward
cargo containers were segregated into areas designated by their serial
numbers and items not identified at that stage were collected into piles
of similar parts for later assessment. As a result of this, two adjacent
containers, one of metal construction the other fibreglass, were
identified as exhibiting damage likely to have been caused by the
explosion. Those parts which could be positively identified as being
from these two containers were assembled onto one of three simple wooden
frameworks, one each for the floor and superstructure of the metal
container and one for the superstructure of the fibreglass container.
From this it was positively determined that the explosion had occurred
within the metal container (serial number AVE 4041 PA), the direct
effects of this being evident also on the forward face of the adjacent
fibreglass container (serial number AVN 7511 PA) and on the local
airframe on the left side of the aircraft in the region of station 700.
It was therefore confirmed that this metal container had been loaded in
position 14L in agreement with the aircraft loading records. While this
work was in progress a buckled section of the metal container skin was
found by an AAIB Inspector to contain, trapped within its folds, an item
which was subsequently identified by forensic scientists at the Royal
Armaments Research and Development Establishment (RARDE) as belonging to
a specific type of radio-cassette player and that this had been fitted
with an improvised explosive device (IED).
The reconstruction of these containers and their
relationship to the aircraft structure is described in detail in
Appendix F. Examination of all other components of the remaining
containers revealed only damage consistent with ejection into the high
speed slipstream and/or ground impact, and that only one device had
detonated within the containers on board the aircraft.
1.12.3 Fuselage three-dimensional reconstruction
1.12.3.1 The reconstruction
The two-dimensional reconstruction successfully
established that there had been an explosion in the forward hold; its
location was established and the general damage characteristics in the
vicinity of the explosion were determined. However, the mechanisms by
which the failure process developed from local damage in the immediate
vicinity of the explosion to the complete structural break-up and
separation of the whole forward section of the fuselage, could not be
adequately investigated without recourse to a more elaborate
reconstruction.
To facilitate this additional work, wreckage
forming a 65 foot section of the fuselage (approximately 30 feet each
side of the explosion) was transported to AAIB Farnborough, where it was
attached to a specially designed framework to form a fully
three-dimensional reconstruction [Appendix B, Figures B-16 and B-17] of
the complete fuselage between stations 360 & 1000 (from the separated
nose section back to the wing cut out). The support framework was
designed to provide full and free access to all parts of the structure,
both internally and externally. Because of height constraints, the
reconstruction was carried out in two parts, with the structure divided
along a horizontal line at approximately the upper cabin floor level.
The previously reconstructed containers were also transported to AAIB
Farnborough to allow correlation of evidence with, and partial
incorporation into, the fuselage reconstruction.
Structure and skin panels were attached to the
supporting framework by their last point of attachment, to provide a
better appreciation of the modes and direction of curling, distortion,
and ultimate separation. Thus, the panels of skin which had petalled
back from the shatter zone were attached at their outer edges, so as to
identify the bending modes of the panels, the extent of the petalled
region, and also the size of the resulting aperture in the hull. In
areas more remote from the explosion, the fracture and tear directions
were used together with distortion and curling directions to determine
the mode of separation, and thus the most appropriate point of
attachment to the reconstruction. Cabin floor beam segments were
supported on a steel mesh grid and a plot of the beam fractures is shown
at Appendix B, Figure B-18.
The cargo container base elements were separated
from the rest of the container reconstruction and transferred to the
main wreckage reconstruction, where the re-assembled container base was
positioned precisely onto the cargo deck. To assist in the correlation
of the initial shatter zone and petalled-out regions with the position
of the explosive device, the boundaries of the skin panel fractures were
marked on a transparent plastic panel which was then attached to the
reconstruction to provide a transparent pseudo-skin showing the
positions of the skin tear lines. This provided a clear visual
indication of the relationship between the skin panel fractures and the
explosive damage to the container base, thus providing a more accurate
indication of the location of the explosive device.
1.12.3.2 Summary of explosive features evident
The three-dimensional reconstruction provided
additional information about the region of tearing and petalling around
the shatter zone. It also identified a number of other regions of
structural damage, remote from the explosion, which were clearly
associated with severe and rapidly applied pressure loads acting normal
to the skin’s internal surface. These were sufficiently sharp-edged to
pre-empt the resolution of pressure induced loads into membrane tension
stresses in the skin: instead, the effect was as though these areas of
skin had been struck a severe ‘pressure blow’ from within the hull.
The two types of damage, i.e. the direct
blast/tearing/petalling damage and the quite separate areas of ‘pressure
blow’ damage at remote sites were evidently caused by separate
mechanisms, though it was equally clear that each was caused by
explosive processes, rather than more general disintegration.
The region of petalling was bounded
(approximately) by frames 680 and 740, and extended from just below the
window belt down nearly to the keel of the aircraft [Appendix B, Figure
B-19, region A]. The resulting aperture measured approximately 17 feet
by 5 feet. Three major fractures had propagated beyond the boundary of
the petalled zone, clearly driven by a combination of hull
pressurisation loading and the relatively long term (secondary) pressure
pulse from the explosion. These fractures ran as follows:
(i) |
rearwards and downward in a
stepped fashion, joining the stringer 38L lap joint at around
station 840, running aft along stringer 38L to around station 920,
then stepping down to stringer 39L and running aft to terminate at
the wing box cut-out [Appendix B, Figure B-19, fracture 1]. |
|
|
(ii) |
downwards and forward to join the
stringer 44L lap joint, then running forward along stringer 44L as
far as station 480 [Appendix B, Figure B-19, fracture 2]. |
|
|
(iii) |
downwards and rearward, joining
the butt line at station 740 to run under the fuselage and up the
right side to a position approximately 18 inches above the cabin
floor level [Appendix B, Figures B-19 and B-20, fracture 3]. |
The propagation of tears upwards from the shatter
zone appeared to have taken the form of a series of parallel fractures
running upwards together before turning towards each other and closing,
forming large flaps of skin which appear to have separated relatively
cleanly.
Regions of skin separation remote from the site of
the explosion were evident in a number of areas. These principally were:
(i) |
A large section of upper fuselage
skin extending from station 500 back to station 760, and from around
stringers 15/19L up as far as stringer 5L [Appendix B, Figures B-19
and B-20, region B], and probably extending further up over the
crown. This panel had separated initially at its lower forward edge
as a result of a pressure blow type of impulse loading, which had
popped the heads from the rivets at the butt joint on frame 500 and
lifted the skin flap out into the airflow. The remainder of the
panel had then torn away rearwards in the airflow. |
|
|
|
A region of ‘quilting’ or
‘pillowing’, i.e. spherical bulging of skin panels between frames
and stringers, was evident on these panels in the region between
station 560 and 680, just below the level of the upper deck floor,
indicative of high internal pressurisation loading [Appendix B,
Figure B-19, region C]. |
|
|
(ii) |
A smaller section of skin between
stations 500 and 580, bounded by stringers 27L and 34L [Appendix B,
Figure B-19, region D], had also been ‘blown’ outwards at its
forward edge and torn off the structure rearwards. A characteristic
curling of the panel was evident, consistent with rapid, energetic
separation from the structure. |
|
|
(iii) |
A section of thick belly skin
extending from station 560, stringers 40R to 44R, and tapering back
to a point at stringer 45R/station720 [Appendix B, Figure B-19 and
B-20, region E], had separated from the structure as a result of a
very heavy ‘pressure blow’ load at its forward end which had popped
the heads off a large number of substantial skin fasteners. The
panel had then torn away rearwards from the structure, curling up
tightly onto itself as it did so – indicating that considerable
excess energy was involved in the separation process (over and above
that needed simply to separate the skin material from its supporting
structure). |
|
|
(iv) |
A panel of skin on the right side
of the aircraft, roughly opposite the explosion, had been torn off
the frames, beginning at the top edge of the panel situated just
below the window belt and tearing downwards towards the belly
[Appendix B, Figure B-20, region F]. This panel was curled downwards
in a manner which suggested significant excess energy. |
Appendix B, Figure B-21 shows a plot of the
fractures noted in the fuselage skins between stations 360 and 1000.
The cabin floor structure was badly disrupted,
particularly in the general area above the explosion, where the floor
beams had suffered localised upward loading sufficient to fracture them,
and the floor panels were missing. Elsewhere, floor beam damage was
mainly limited to fractures at the outer ends of the beams and at the
centreline, leaving sections of separated floor structure comprising a
number of half beams joined together by the Nomex honeycomb floor
panels.
1.12.3.3 General damage features not directly
associated with explosive forces.
A number of features appeared to be a part of the
general structural break-up which followed on from the explosive damage,
rather than being a part of the explosive damage process itself. This
general break-up was complex and, to a certain extent, random. However,
analysis of the fractures, surface scores, paint smears and other
features enabled a number of discreet elements of the break-up process
to be identified. These elements are summarised below.
(i) |
Buckling of the window belts on
both sides of the aircraft was evident between stations 660 and 800.
That on the left side appeared to be the result of in-plane bending
in a nose up sense, followed by fracture. The belt on the right side
had a large radius curve suggesting lateral deflection of the
fuselage possibly accompanied by some longitudinal compression. This
terminated in a peeling failure of the riveted joint at station
800. |
|
|
(ii) |
On the left side three fractures,
apparently resulting from in-plane bending/buckling distortion, had
traversed the window belt [Appendix B, Figure B-21, detail G]. Of
these, the forward two had broken through the window apertures and
the aft fracture had exploited a rivet line at the region of
reinforcement just forward of the L2 door aperture. On the right
side, the window belt had peeled rearwards, after buckling had
occurred, separating from the rest of the fuselage, following rivet
failure, at the forward edge of the R2 door aperture. |
|
|
(iii) |
All crown skins forward of frame
840 were badly distorted and a number of pieces were missing. It was
clearly evident that the skin sections from this region had struck
the empennage and/or other structure following separation. |
|
|
(iv) |
The fuselage left side lower lobe
from station 740 back to the wing box cut-out, and from the window
level down to the cargo deck floor (the fracture line along stringer
38L), had peeled outwards, upwards and rearwards – separating from
the rest of the fuselage at the window belt. The whole of this
separated section had then continued to slide upwards and rearwards,
over the fuselage, before being carried back in the slipstream and
colliding with the outer leading edge of the right horizontal
stabiliser, completely disrupting the outer half. A fragment of
horizontal stabiliser spar cap was found embedded in the fuselage
structure adjacent to the two vent valves, just below, and forward
of, the L2 door [Appendix B, Figure B-22]. |
|
|
(v) |
A large, clear, imprint of semi-eliptical
form was apparent on the lower right side at station 360 which had
evidently been caused by the separating forward fuselage section
striking the No 3 engine as it swung rearwards and to the right
(confirmed by No 3 engine fan cowl damage). |
1.12.3.4 Tailplane three-dimensional
reconstruction
The tailplane structural design took the form of a
forward and an aft torque box. The forward box was constructed from
light gauge aluminium alloy sheet skins, supported by closely pitched,
light gauge nose ribs but without lateral stringers. The aft torque box
incorporated heavy gauge skin/stringer panels with more widely spaced
ribs. The front spar web was of light gauge material. Leading edge
impacts inflicted by debris would therefore have had the capacity to
reduce the tailplane’s structural integrity by passing through the light
gauge skins and spar web into the interior of the aft torque box,
damaging the shear connection between top and bottom skins in the
process and thereby both removing the bending strength of the box and
opening up the weakened structure to the direct effects of the airflow.
Examination of the rebuilt tailplane structure at
AAIB Farnborough left little doubt that it had been destroyed by debris
striking its leading edges. In addition, the presence on the skins of
smear marks indicated that some unidentified soft debris had contacted
those surfaces whilst moving with both longitudinal and lateral velocity
components relative to the aircraft.
The reconstructed left tailplane [Appendix B,
Figure B-23] showed evidence that disruption of the inboard leading
edge, followed respectively by the forward torque box, front spar web
and main torque box, occurred as a result of frontal impact by the base
of a baggage container. Further outboard, a compact object appeared to
have struck the underside of the leading edge and penetrated to the aft
torque box. In both cases, the loss of the shear web of the front spar
appeared to have permitted local bending failure of the remaining main
torque box structure in a tip downwards sense, consistent with the
normal load direction. For both events to have occurred it would be
reasonable to assume that the outboard damage preceded that occurring
inboard.
The right tailplane exhibited massive leading edge
impact damage on the outboard portion which also appeared to have
progressed to disruption of the aft torsion box. A fragment of right
tailplane spar cap was found embedded in the fuselage structure adjacent
to the two vent valves, just below, and forward of, the L2 door and it
is clear that this area of forward left fuselage had travelled over the
top of the aircraft and contributed to the destruction of the outboard
right tailplane.
1.12.4 Examination of engines
All four engines had struck the ground in
Lockerbie with considerable velocity and therefore sustained major
damage, in particular to most of the fan blades. The No 3 engine had
fallen 1,100 metres north of the other three engines, striking the
ground on its rear face, penetrating a road surface and coming to rest
without any further change of orientation i.e. with the front face
remaining uppermost. The intake area contained a number of loose items
originating from within the cabin or baggage hold. It was not possible
initially to determine whether any of the general damage to any of the
engine fans or the ingestion noted in No 3 engine intake occurred whilst
the relevant engines were delivering power or at a later stage.
Numbers 1, 2 and 3 engines were taken to British
Airways Engine Overhaul Limited for detailed examination under AAIB
supervision in conjunction with a specialist from the Pratt and Whitney
Engine Company. During this examination the following points were noted:
(i) |
No 2 engine (situated closest to
the site of the explosion) had evidence of blade "shingling" in the
area of the shrouds consistent with the results of major airflow
disturbance whilst delivering power. (This effect is produced when
random bending and torsional deflection occurs, permitting the
mid-span shrouds to disengage and repeatedly strike the adjacent
aerofoil surfaces of the blades). The interior of the air intake
contained paint smears and other evidence suggesting the passage of
items of debris. One such item of significance was a clear
indentation produced by a length of cable of diameter and strand
size similar to that typically attached to the closure curtains on
the baggage containers. |
|
|
(ii) |
No 3 engine, identified on site
as containing ingested debris from within the aircraft, nonetheless
had no evidence of the type of shingling seen on the blades of No 2
engine. Such evidence is usually unmistakable and its absence is a
clear indication that No 3 engine did not suffer a major intake
airflow disturbance whilst delivering significant power. The intake
structure was found to have been crushed longitudinally by an impact
on the front face although, as stated earlier, it had struck the
ground on its rear face whilst falling vertically. |
|
|
(iii) |
All 3 engines had evidence of
blade tip rubs on the fan cases having a combination of
circumference and depth greater than hitherto seen on any
investigation witnessed on Boeing 747 aircraft by the Pratt and
Whitney specialists. Subsequent examination of No 4 engine confirmed
that it had a similar deep, large circumference tip rub. These
tip-rubs on the four engines were centred at slightly different
clock positions around their respective fan cases. |
The Pratt and Whitney specialists supplied
information which was used to interpret the evidence found on the blades
and fan cases including details of engine dynamic behaviour necessary to
produce the tip rub evidence. This indicated that the depth and
circumference of tip rubs noted would have required a marked nose down
change of aircraft pitch attitude combined with a roll rate to the left.
Pratt and Whitney also advised that:
(i) |
Airflow disruption such as that
presumed to have caused the shingling observed on No 2 engine fan
blades was almost invariably the result of damage to the fan blade
aerofoils, resulting from ingestion or blade failure. |
|
|
(ii) |
Tip rubs of a depth and
circumference noted on all four engines could be expected to reduce
the fan rotational energy on each to a negligible value within
approximately 5 seconds. |
|
|
(iii) |
Airflow disruption sufficient to
cause the extent of shingling noted on the fan blades of No 2 engine
would also reduce the rotational fan energy to a negligible value
within approximately 5 seconds. |
1.13 Medical and pathological information
The results of the post mortem examination of the
victims indicated that the majority had experienced severe multiple
injuries at different stages, consistent with the in-flight
disintegration of the aircraft and ground impact. There was no
pathological indication of an in-flight fire and no evidence that any of
the victims had been injured by shrapnel from the explosion. There was
also no evidence which unequivocally indicated that passengers or cabin
crew had been killed or injured by the effects of a blast. Although it
is probable that those passengers seated in the immediate vicinity of
the explosion would have suffered some injury as a result of blast, this
would have been of a secondary or tertiary nature.
Of the casualties from the aircraft, the majority
were found in areas which indicated that they had been thrown from the
fuselage during the disintegration. Although the pattern of distribution
of bodies on the ground was not clear cut there was some correlation
with seat allocation which suggested that the forward part of the
aircraft had broken away from the rear early in the disintegration
process. The bodies of 10 passengers were not recovered and of these, 8
had been allocated seats in rows 23 to 28 positioned over the wing at
the front of the economy section. The fragmented remains of 13
passengers who had been allocated seats around the eight missing persons
were found in or near the crater formed by the wing. Whilst there is no
unequivocal proof that the missing people suffered the same fate, it
would seem from the pattern that the missing passengers remained
attached to the wing structure until impact.
1.14 Fire
Of the several large pieces of aircraft wreckage
which fell in the town of Lockerbie, one was seen to have the appearance
of a ball of fire with a trail of flame. Its final path indicated that
this was the No 3 engine, which embedded itself in a road in the
north-east part of the town. A small post impact fire posed no hazard to
adjacent property and was later extinguished with water from a hosereel.
The three remaining engines landed in the Netherplace area of the town.
One severed a water main and the other two, although initially on fire,
were no risk to persons or property and the fires were soon
extinguished.
A large, dark, delta shaped object was seen to
fall at about the same time in the Sherwood area of the town. It was not
on fire while in the air, however, a fireball several hundred feet
across followed the impact. It was of relatively short duration and
large amounts of debris were thrown into the air, the lighter particles
being carried several miles downwind, while larger pieces of burning
debris caused further fires, including a major one at the Townfoot
Garage, up to 350 metres from the source. It was determined that the
major part of both wings, which included the aircraft fuel tanks, had
formed the crater. A gas main had also been ruptured during the impact.
At 19.04 hrs the Dumfries Fire Brigade Control
received a call from a member of the public which indicated that there
had been a "huge boiler explosion" at Westacres, Lockerbie, however,
subsequent calls soon made it clear that it was an aircraft which had
crashed. At 19.07 hrs the first appliances were mobile and at 1910 hrs
one was in attendance in the Rosebank area. Multiple fires were
identified and it soon became apparent that a major disaster had
occurred in the town and the Fire Brigade Major Incident Plan was
implemented. During the initial phase 15 pumping appliances from various
brigades were deployed but this number was ultimately increased to 20.
At 22.09 hrs the Firemaster made an assessment of
the situation. He reported that there was a series of fires over an area
of the town centre extending 1? by ? mile. The main concentration of the
fire was in the southwest of the town around Sherwood Park and Sherwood
Crescent. Appliances were in attendance at other fires in the town,
particularly in Park Place and Rosebank Crescent. Water and electricity
supplies were interrupted and water had to be brought into the town.
By 02.22 hrs on 22 December, all main seats of
fire had been extinguished and the firemen were involved in turning over
and damping down. At 04.42 hrs small fires were still occurring but had
been confined to the Sherwood Crescent area.
1.15 Survival aspects
1.15.1 Survivability
The accident was not survivable.
1.15.2 Emergency services
A chronology of initial responses by the emergency
services is listed below:-
Time |
Event |
19.03 hrs |
Radio message from Police patrol
in Lockerbie to Dumfries and Galloway Constabulary reporting an
aircraft crash at Lockerbie. |
19.04 hrs |
Emergency call to Dumfries and
Galloway Fire Brigade. |
19.37 hrs |
First ambulances leave for
Dumfries and Galloway Royal Infirmary with injured town residents.
(2- serious; 3- minor) |
19.40 hrs |
Sherwood Park and Sherwood
Crescent residents evacuated to Lockerbie Town Hall. |
20.25 hrs |
Nose section of N739PA discovered
at Tundergarth (approximately 4 km east of Lockerbie). |
During the next few days a major emergency
operation was mounted using the guidelines of the Dumfries and Galloway
Regional Peacetime Emergency Plan. The Dumfries and Galloway
Constabulary was reinforced by contingents from Strathclyde and Lothian
& Borders Constabularies. Resources from HM Forces were made available
and this support was subsequently authorised by the Ministry of Defence
as Military Aid to the Civil Power. It included the provision of
military personnel and a number of helicopters used mainly in the search
for and recovery of aircraft wreckage. It was apparent at an early stage
that there were no survivors from the aircraft and the search and
recovery of bodies was mainly a Police task with military assistance.
Many other agencies were involved in the provision
of welfare and support services for the residents of Lockerbie,
relatives of the aircraft’s occupants and personnel involved in the
emergency operation.
1.16 Tests and research
An explosive detonation within a fuselage, in
reasonably close proximity to the skin, will produce a high intensity
spherically propagating shock wave which will expand outwards from the
centre of detonation. On reaching the inner surface of the fuselage
skin, energy will partially be absorbed in shattering, deforming and
accelerating the skin and stringer material in its path. Much of the
remaining energy will be transmitted, as a shock wave, through the skin
and into the atmosphere but a significant amount of energy will be
returned as a reflected shock wave, which will travel back into the
fuselage interior where it will interact with the incident shock to
produce Mach stem shocks – re-combination shock waves which can have
pressures and velocities of propagation greater than the incident shock.
The Mach stem phenomenon is significant because it
gives rise (for relatively small charge sizes) to a geometric limitation
on the area of skin material which the incident shock wave can shatter,
irrespective of charge size, thus providing a means of calculating the
standoff distance of the explosive charge from the fuselage skin.
Calculations suggest that a charge standoff distance of aproximately 25
inches would result in a shattered region approximately 18 to 20 inches
in diameter, comparable to the size of the shattered region evident in
the wreckage. This aspect is covered in greater detail in [Appendix G].
1.17 Additional information
1.17.1 Recorded radar information
Recorded radar information on the aircraft was
available from 4 radar sites. Initial analysis consisted of viewing the
recorded information as it was shown to the controller on the radar
screen from which it was clear that the flight had progressed in a
normal manner until secondary surveillance radar (SSR) was lost.
The detailed analysis of the radar information
concentrated on the break-up of the aircraft. The Royal Signals and
Radar Establishment (RSRE) corrected the radar returns for fixed errors
and converted the SSR returns to latitude and longitude so that an
accurate time and position for the aircraft could be determined. The
last secondary return from the aircraft was recorded at 19.02:46.9 hrs,
identifying N739PA at Flight Level 310, and at the next radar return
there is no SSR data, only 4 primary returns. It was concluded that the
aircraft was, by this time, no longer a single return and, considering
the approximately 1 nautical mile spread of returns across track, that
items had been ejected at high speed probably to both right and left of
the aircraft.
Each rotation of the radar head thereafter showed
the number of returns increasing, with those first identified across
track having slowed down very quickly and followed a track along the
prevailing wind line. The radar evidence then indicated that a further
break-up of the aircraft had occurred and formed a parallel wreckage
trail to the north of the first. From the absence of any returns
travelling along track it was concluded that the main wreckage was
travelling almost vertically downwards for much of the time.
A detailed analysis of the recorded radar
information, together with the radar, ATC and seismic recordings is
contained in Appendix C.
1.17.2 Seismic data
The British Geological Survey has a number of
seismic monitoring stations in Southern Scotland. Stations close to
Lockerbie recorded a seismic event measuring 1.6 on the Richter scale
and, with appropriate corrections for the times of the waves to reach
the sensors, it was established that this occurred at 19.03:36.5 hrs ?1
second. A further check was made by triangulation techniques from the
information recorded by the various sensors.
An analysis of the seismic recording, together
with the radar, ATC and radar information is contained in Appendix C.
1.17.3 Trajectory analysis
A detailed trajectory analysis was carried out by
Cranfield Institute of Technology in an effort to provide a sequence for
the aircraft disintegration. This analysis comprised several separate
processes, including individual trajectory calculations for a limited
number of key items of wreckage and mathematical modelling of trajectory
paths adopted by a series of hypothetical items of wreckage encompassing
the drag/weight spectrum of the actual wreckage.
The work carried out at Cranfield enabled the
reasons for the two separate trails to be established. The narrow
northern trail was shown to be created by debris released from the
aircraft in a vertical dive between 19,000 and 9,000 feet overhead
Lockerbie. The southern trail, longer and straight for most of its
length, appeared to have been created by wreckage released during the
initial disintegration at altitude whilst the aircraft was in level
flight. Those items falling closest to Lockerbie would have been those
with higher density which would travel a significant distance along
track before losing all along-track velocity, whilst only drifting a
small distance downwind, owing to the high speed of their descent. The
most westerly items thus showed the greatest such effect. The southern
trail therefore had curved boundaries at its western end with the
curvature becoming progressively less to the east until the wreckage
essentially fell in a straight band. Thus wreckage in the southern trail
positioned well to the east could be assumed to have retained negligible
velocity along aircraft track after separation and the along-track
distribution could be used to establish an approximate sequence of
initial disintegration.
The analysis calculated impact speeds of 120 kts
for the nose section weighing approximately 17,500 lb and 260 kts for
the engines and pylons which each weighed about 13,500 lb. Based on the
best available data at the time, the analysis showed that the wing
(approximately 100,000 lb of structure containing an estimated 200,000
lb of fuel) could have impacted at a speed, in theory, as high as 650
kts if it had ‘flown’ in a streamlined attitude such that the drag
coefficient was minimal. However, because small variations of wing
incidence (and various amounts of attached fuselage) could have resulted
in significant increases in drag coefficient, the analysis also
recognized that the final impact speed of the wing could have been
lower.
1.17.4 Space debris re-entry
Four items of space debris were known to have
re-entered the Earth’s atmosphere on 21 December 1988. Three of these
items were fragments of debris which would not have survived re-entry,
although their burn up in the upper atmosphere might have been visible
from the Earth’s surface. The fourth item landed in the USSR at 09.50
hrs UTC.
2 ANALYSIS
2.1 Introduction
The airport security and criminal aspects of the
destruction of Boeing 747 registration N739PA near Lockerbie on 21
December 1988 are the subjects of a separate investigation and are not
covered in this report. This analysis discusses the technical aspects of
the disintegration of the aircraft and considers possible ways of
mitigating the effects of an explosion in the future.
2.2 Explosive destruction of the aircraft
The geographical position of the final secondary
return at 19.02:46.9 hrs was calculated by RSRE to be OS Grid Reference
15257772, annotated Point A in Appendix B, Figure B-4, with an accuracy
considered to be better than ?300 metres This return was received 3.1?1
seconds before the loud sound was recorded on the CVR at 19.02:50 hrs.
By projecting from this position along the track of 321?(Grid) for 3.1?1
seconds at the groundspeed of 434 kts, the position of the aircraft was
calculated to be OS Grid Reference 14827826, annotated Point B in
Appendix B, Figure B-4, within an accuracy of ?525 metres. Based on the
evidence of recorded data only, Point B therefore represents the
geographical position of the aircraft at the moment the loud sound was
recorded on the CVR.
The datum line, discussed at paragraph 1.12.1.6,
was derived from a detailed analysis of the distribution of specific
items of wreckage, including those exhibiting positive evidence of a
detonating high performance plastic explosive. The scatter of these
items about the datum line may have been due partly to velocities
imparted by the force of the detonating explosive and partly by the
difficulty experienced in pinpointing the location of the wreckage
accurately in relatively featureless terrain and poor visibility.
However, the random nature of the scatter created by these two effects
would have tended to counteract one another, and a major error in any
one of the eleven grid references would have had little overall effect
on the whole line. There is, therefore, good reason to have confidence
in the validity of the datum line.
The items used to define the datum line, included
those exhibiting positive evidence of a detonating high performance
plastic explosive, would have been the first pieces to have been
released from the aircraft. The datum line was projected westwards until
it intersected the known radar track of the aircraft in order to derive
the position of the aircraft along track at which the explosive items
were released and therefore the position at which the IED had detonated.
This position was OS grid reference 146786 and is annotated Point C in
Appendix B, Figure B-4. Point C was well within the circle of accuracy
(?525 metres) of the position at which the loud noise was heard on the
CVR (Point B). There can, therefore, be no doubt that the loud noise on
the CVR was directly associated with the detonation of the IED and that
this explosion initiated the disintegration process and directly caused
the loss of the aircraft.
2.3 Flight recorders
2.3.1 Digital flight data recordings
A working group of the European Organisation for
Civil Aviation Electronics (EUROCAE) was, during the period of the
investigation, formulating new standards (Minimum Operational
Performance Requirement for Flight Data Recorder Systems, Ref:- ED55)
for future generation flight recorders which would have permitted delays
between parameter input and recording (buffering) of up to ? second.
These standards are intended to form the basis of new CAA specifications
for flight recorders and may be adopted worldwide.
The analysis of the recording from the DFDR fitted
to N739PA, which is detailed in Appendix C, showed that the recorded
data simply stopped. Following careful examination and correlation of
the various sources of recorded information, it was concluded that this
occurred because the electrical power supply to the recorder had been
interrupted at 19.02:50 hrs ?1 second. Only 17 bits of data were not
recoverable (less that 23 milliseconds) and it was not possible to
establish with any certainty if this data was from the accident flight
or was old data from a previous recording.
The analysis of the final data recorded on the
DFDR was possible because the system did not buffer the incoming data.
Some existing recorders use a process whereby data is stored temporarily
in a memory device (buffer) before recording. The data within this
buffer is lost when power is removed from the recorder and in currently
designed recorders this may mean that up to 1.2 seconds of final data
contained within the buffer is lost. Due to the necessary processing of
the signals prior to input to the recorder, additional delays of up to
300 milliseconds may be introduced. If the accident had occurred when
the aircraft was over the sea, it is very probable that the relatively
few small items of structure, luggage and clothing showing positive
evidence of the detonation of an explosive device would not have been
recovered. However, as flight recorders are fitted with underwater
location beacons, there is a high probability that they would have been
located and recovered. In such an event the final milliseconds of data
contained on the DFDR could be vital to the successful determination of
the cause of an accident whether due to an explosive device or other
catastrophic failure. Whilst it may not be possible to reduce some of
the delays external to the recorder, it is possible to reduce any data
loss due to buffering of data within the data acquisition unit.
It is, therefore, recommended that manufacturers
of existing recorders which use buffering techniques give consideration
to making the buffers non-volatile, and hence recoverable after power
loss. Although the recommendation on this aspect, made to the EUROCAE
working group during the investigation, was incorporated into ED55, it
is also recommended that Airworthiness Authorities re-consider the
concept of allowing buffered data to be stored in a volatile memory.
2.3.2 Cockpit voice recorders
The analysis of the cockpit voice recording, which
is detailed in Appendix C, concluded that there were valid signals
available to the CVR when it stopped at 19.02:50 hrs ?1 second because
the power supply to the recorder was interrupted. It is not clear if the
sound at the end of the recording is the result of the explosion or is
from the break-up of the aircraft structure. The short period between
the beginning of the event and the loss of electrical power suggests
that the latter is more likely to be the case. In order to respond to
events that result in the almost immediate loss of the aircraft’s
electrical power supply it was therefore recommended during the
investigation that the regulatory authorities consider requiring CVR
systems to contain a short duration (i.e. no greater than 1 minute)
back-up power supply.
2.3.3 Detection of explosive occurrences
In the aftermath of the Air India Boeing 747
accident (AI 182) in the North Atlantic on 23 June 1985, RARDE were
asked informally by AAIB to examine means of differentiating, by
recording violent cabin pressure pulses, between the detonation of an
explosive device within the cabin (positive pulse) and a catastrophic
structural failure (negative pulse). Following the Lockerbie disaster it
was considered that this work should be raised to a formal research
project. Therefore, in February 1989, it was recommended that the
Department of Transport fund a study to devise methods of recording
violent positive and negative pressure pulses, preferably utilising the
aircraft’s flight recorder systems. This recommendation was accepted.
Preliminary results from the trials indicate that,
if a suitable sensor can be developed, its output will need to be
recorded in real time and therefore it may require wiring to the CVR
installation. This will further strengthen the requirement for battery
back up of the CVR electrical power supply.
2.4 IED position within the aircraft
From the detailed examination of the reconstructed
luggage containers, discussed at paragraph 1.12.2.4 and in Appendix F,
it was evident that the IED had been located within a metal container
(serial number AVE 4041 PA), near its aft outboard quarter as shown in
Appendix F, Figure F-13. It was also clear that the container was loaded
in position 14L of the forward hold which placed the explosive charge
approximately 25 inches inboard from the fuselage skin at frame 700.
There was no evidence to indicate that there was more than one explosive
charge.
2.5 Engine evidence
To produce the fan blade tip rub damage noted on
all engines by means of airflow inclined to the axes of the nacelles
would have required a marked nose down change of aircraft pitch attitude
combined with a roll rate to the left while all of the engines were
attached to the wing.
The shingling damage noted on the fan blades of No
2 engine can only be attributed to airflow disturbance caused by
ingestion related fan blade damage occurring when substantial power was
being delivered. This is readily explained by the fact that No 2 engine
intake is positioned some 27 feet aft and 30 feet outboard of the site
of the explosion and that the interior of the intake exhibited a number
of prominent paint smears and general foreign object damage. This damage
included evidence of a strike by a cable similar to that forming part of
the closure curtain of a typical baggage container. It is inconceivable
that an independent blade failure could have occurred in the short time
frame of this event. By similar reasoning, the absence of such shingling
damage on blades of No 3 engine was a reliable indication that it
suffered no ingestion until well into the accident sequence.
The combination of the position of the explosive
device and the forward speed of the aircraft was such that significant
sized debris resulting from the explosion would have been available to
be ingested by No 2 engine within milliseconds of the explosion. In view
of the fact that the tip rub damage observed on the fan case of No 2
engine is of similar magnitude to that observed on the other three
engines it is reasonable to deduce that a manoeuvre of the aircraft
occurred before most of the energy of the No 2 engine fan was lost due
to the effect of ingestion (seen only in this engine). Since this
shingling effect could only readily be produced as a by-product of
ingestion whilst delivering considerable power, it is reasonable to
assume that this was also occurring before loss of major fan energy due
to tip rubbing took place. Hence both phenomena must have been occurring
simultaneously, or nearly so, to produce the effects observed and must
have occupied a time frame of substantially less than 5 seconds. The
onset of this time period would have been the time at which debris from
the explosion first inflicted damage to fan blades in No 3 engine and,
since the fan is only approximately 40 feet from the location of the
explosive device, this would have been an insignificant time interval
after the explosion.
It was therefore concluded from this evidence that
the wing with all of the engines attached had achieved a marked nose
down and left roll attitude change well within 5 seconds of the
explosion.
2.6 Detachment of forward fuselage
Examination of the three major structural elements
either side of the region of station 800 on the right side of the
fuselage makes it clear that to produce the curvature of the window belt
and peeling of the riveted joint at the R2 door aperture requires the
door pillar to be securely in position and able to react longitudinal
and lateral loads. This in turn requires the large section of fuselage
on the right side between stations 760 and 1000 (incorporating the right
half of the floor) to be in position in order to locate the lower end of
the door pillar. Thus both these sections must have been in position
until the section from station 560 to 800 (right side) had completed its
deflection to the right and peeled from the door pillar. Separation of
the forward fuselage must thus have been complete by the time all three
items mentioned above had fallen free.
2.7 Speed of initial disintegration
The distribution of wreckage in the bands between
the datum line and the 250, 300, 600 and 900 metre lines was examined in
detail. The positions of these items of structure on the aircraft are
shown in Appendix B, Figures B-10 to B-13. It should be noted that the
position on the ground of these items, although separated by small
distances when measured in a direction along aircraft track, were
distributed over large distances when measured along the wreckage trail.
All were recovered from positions far enough to the east to be in that
part of the southern trail which was sufficiently close, theoretically,
to a straight line for any curvature effect to be neglected.
The wreckage found in each of the bands enabled an
approximate sequence of break-up to be established. It was clear that as
the distance travelled from the datum line increased, items of wreckage
further from the station of the IED were encountered. The items shown on
the diagram as falling on the 250 metre band also include those
fragments of lower forward fuselage skin having evidence of explosive
damage and presumed to have separated as a direct result of the blast.
However, a few portions of the upper forward fuselage were also found
within the 250 metre band, suggesting that these items had also
separated as a result of the blast.
By the time the 300 metre line was reached much of
the structure from the right side in the region of the explosive device
had been shed. This included the area of window belt, referred to in
paragraph 2.6 above, which gave clear indications that the forward
structure had detached to the right and finally peeled away at station
800. It also included the areas of adjacent structure immediately to the
rear of station 800 about which the forward structure would have had to
pivot. By the time the 600 metre line was reached, there was clearly
insufficient structure left to connect the forward fuselage with the
remainder of the aircraft. Wreckage between the 600 and 900 metre lines
consisted of structure still further from the site of the IED.
There is evidence that a manoeuvre occurred at the
time of the explosion which would have produced a significant change of
the aircraft’s flight path, however, it is considered that the change in
the horizontal velocity component in the first few seconds would not
have been great. The original groundspeed of the aircraft was therefore
used in conjunction with the distribution of wreckage in the successive
bands to establish an approximate time sequence of break-up of the
forward fuselage. Assuming the original ground speed of 434 Kts, the
elapsed flight times from the datum to each of the parellel lines were
calculated to be:
Distance (metres) |
250 |
300 |
600 |
900 |
Time (seconds) |
1.1 |
1.3 |
2.7 |
4.0 |
Thus, there is little doubt that separation of the
forward fuselage was complete within 2 to 3 seconds of the explosion.
The separate assessment of the known grid
references of tailplane and elevator wreckage in the southern trail
revealed that those items were evenly distributed about the 600 metre
line and therefore that most of the tailplane damage occurred after
separation of the forward fuselage was complete.
2.8 The manoeuvre following the explosion
The engine evidence, timing and mode of
disintegration of the fuselage and tailplane suggests that the latter
did not sustain significant damage until the forward fuselage
disintegration was well advanced and the pitch/roll manoeuvre was also
well under way.
Examination of the three dimensional
reconstruction makes it clear that both main and upper deck floors were
disrupted by the explosion. Since pitch control cables are routed
through the upper deck floor beams and the roll control cables through
the main deck beams, there is a strong possibility that movement of the
beams under explosive forces would have applied inputs to the control
cables, thus operating control surfaces in both axes.
2.9 Secondary disintegration
The distribution of fin debris between the trails
suggests that disintegration of the fin began shortly before the
vertical descent was established. No single mode of failure was
identified and the debris which had struck the leading edge had not
caused major disruption. The considerable fragmentation of the thick
panels of the aft torque box was also very different from that noted on
the corresponding structure of the tailplanes. It was therefore
concluded that the mode of failure was probably flutter.
The finding, in the northern trail, of a slide
raft wrapped around a flap track fairing suggests that at a later stage
of the disintegration the rear of the aircraft must have experienced a
large angle of sideslip. The loss of the fin would have made this
possible and also subjected the structure to large side loads. It is
possible that such side loading would have assisted the disintegration
of the rear fuselage and also have caused bending failure of the pylon
attachments of the remaining three engines.
2.10 Impact speed of components
The trajectory analysis carried out by Cranfield
Institute of Technology calculated impact speeds of 120 kts for the nose
section, and 260 kts for the engines and pylons. These values were
considered to be reliable because the drag coefficients could be
estimated with a reasonable degree of confidence. Based on the best
available data at the time, the analysis also showed that the wing could
have impacted at a speed, in theory, as high as 650 kts if it had flown
in a streamlined attitude such that the drag coefficient was minimal.
However, it was also recognized that relatively small changes in the
angle of incidence of the wing would have produced a significant
increase in drag with a consequent reduction in impact speed. Refinement
of timing information and radar data subsequent to the Cranfield
analysis has enabled a revised estimate to be made of the mean speed of
the wing during the descent.
The engine evidence indicated that there had been
a large nose down attitude change of the aircraft early in the event.
The Cranfield analysis also showed that the rear fuselage had
disintegrated while essentially in a vertical descent between 19,000 and
9,000 feet over Lockerbie. Assuming that, following the explosion, the
wing followed a straight line descending flight profile from 31,000 feet
to 19,000 feet directly overhead Lockerbie and then descended vertically
until impact, the wing would have travelled the minimum distance
practicable. The ground distance between the geographical position at
which the disintegration started (Figure B-4, Point B) and the crater
made by the wing impact was 2997 ?525 metres (9833 ?1722 feet). The time
interval between the explosion and the wing impact was established in
Appendix C as 46.5 ?2 seconds. Based on the above times and distances
the mean linear speed achieved by the wing would have been about 440 kts.
The impact location of Nos 1, 2, and 4 engines
closely grouped in Lockerbie was consistent with their nearly vertical
fall from a point above the town. If they had separated at about 19,000
feet and the wing had then flown as much as one mile away from the
overhead position before tracking back to impact, the total flight path
length of the wing would not have required it to have achieved a mean
linear speed in excess of 500 kts.
Any speculation that the flight path of the wing
could have been longer would have required it to have undergone
manoeuvres at high speed in order to arrive at the 19,000 feet point.
The manoeuvres involved would almost certainly have resulted in failure
of the primary wing structure which, from distribution of wing debris,
clearly did not occur. Alternatively the wing could have travelled more
than one mile from Lockerbie after reaching the 19,000 feet point, but
this was considered unlikely. It is therefore concluded that the mean
speed of the wing during the descent was in the region of 440 to 500 kts.
2.11 Sequence of disintegration
Analysis of wreckage in each of the bands, taken
in conjunction with the engine evidence and the three-dimensional
reconstruction, suggests the following sequence of disintegration:
(i) |
The initial explosion triggered a
sequence of events which effectively destroyed the structural
integrity of the forward fuselage. Little more then remained between
stations 560 and 760 (approximately) than the window belts and the
cabin sidewall structure immediately above and below the windows,
although much of the cargo-hold floor structure appears to have
remained briefly attached to the aircraft. [Appendix B, Figure
B-24] |
(ii) |
The main portion of the aircraft
simultaneously entered a manoeuvre involving a marked nose down and
left roll attitude change, probably as a result of inputs applied to
the flying control cables by movement of structure. |
(iii) |
Failure of the left window belt
then occurred, probably in the region of station 710, as a result of
torsional and bending loads on the fuselage imparted by the
manoeuvre (i.e. the movement of the forward fuselage relative to the
remainder of the aircraft was an initial twisting motion to the
right, accompanied by a nose up pitching deflection). |
(iv) |
The forward fuselage deflected to
the right, pivoting about the starboard window belt, and then peeled
away from the structure at station 800. During this process the
lower nose section struck the No 3 engine intake causing the engine
to detach from its pylon. This fuselage separation was apparently
complete within 3 seconds of the explosion. |
(v) |
Structure and contents of the
forward fuselage struck the tail surfaces contributing to the
destruction of the outboard starboard tailplane and causing
substantial damage to the port unit. This damage occurred
approximately 600 metres track distance after the explosion and
therefore appears to have happened after the fuselage separation was
complete. |
(vi) |
Fuselage structure continued to
break away from the aircraft and the separated forward fuselage
section as they descended. |
(vii) |
The aircraft maintained a
steepening descent path until it reached the vertical in the region
of 19,000 feet approximately over the final impact point. Shortly
before it did so the tail fin began to disintegrate. |
(viii) |
The mode of failure of the fin is
not clear, however, flutter of its structure is suspected. |
(ix) |
Once established in the vertical
dive, the fin torque box continued to disintegrate, possibly
permitting the remainder of the aircraft to yaw sufficiently to
cause side load separation of Nos 1, 2 and 4 engines, complete with
their pylons. |
(x) |
Break-up of the rear fuselage
occurred during the vertical descent, possibly as a result of loads
induced by the yaw, leaving a section of cabin floor and baggage
hold from approximately stations 1241 to 1920, together with 3
landing gear units, to fall into housing at Rosebank Terrace. |
(xi) |
The main wing structure struck
the ground with a high yaw angle at Sherwood Crescent. |
2.12 Explosive mechanisms and the structural
disintegration
The fracture and damage pattern analysis was
mainly of an interpretive nature involving interlocking pieces of subtle
evidence such as paint smears, fracture and rivet failure
characteristics, and other complex features. In the interests of
brevity, this analysis will not discuss the detailed interpretation of
individual fractures or damage features. Instead, the broader ‘damage
picture’ which emerged from the detailed work will be discussed in the
context of the explosive mechanisms which might have produced the
damage, with a view to identifying those features of greatest
significance.
It is important to keep in mind that whilst the
processes involved are considered and discussed separately, the
timescales associated with shock wave propagation and the high velocity
gas flows are very short compared with the structural response
timescales. Consequently, material which was shattered or broken by the
explosive forces would have remained in place for a sufficiently long
time that the structure can be considered to have been intact throughout
much of the period that these explosive propagation phenomena were
taking place.
2.12.1 Direct blast effect
2.12.1.1 Shock wave propagation
The direct effect of the explosive detonation
within the container was to produce a high intensity spherically
propagating shock wave which expanded from the centre of detonation
close to the side of the container, shattering part of the side and base
of the container as it passed through into the gap between the container
and the fuselage skin. In breaking out of the container, some internal
reflection and Mach stem interaction would have occurred, but this would
have been limited by the absorptive effect of the baggage inboard,
above, and forward of the charge. The force of the explosion breaking
out of the container would therefore have been directed downwards and
rearwards.
The heavy container base was distorted and torn
downwards, causing buckling of the adjoining section of frame 700, and
the container sides were blasted through and torn, particularly in the
aft lower corner. Some of the material in the direct path of the
explosive pressure front was reduced to shrapnel sized pieces which were
rapidly accelerated outwards behind the primary shock front. Because of
the overhang of the container’s sloping side, fragments from both the
device itself and the container wall impacted the projecting external
flange of the container base edge member, producing micro cratering and
sooting. Metallurgical examination of the internal surfaces of these
craters identified areas of melting and other features which were
consistent only with the impact of very high energy particles produced
by an explosion at close quarters. Analysis of material on the crater
surfaces confirmed the presence of several elements and compounds
foreign to the composition of the edge member, including material
consistent with the composition of the sheet aluminium forming the
sloping face of the container.
On reaching the inner surface of the fuselage
skin, the incident shock wave energy would partially have been absorbed
in shattering, deforming and accelerating the skin and stringer material
in its path. Much of its energy would have been transmitted, as a shock
wave, through the skin and into the atmosphere [Appendix B, Figure
B-25], but a significant amount of energy would have been returned as a
reflected shock wave, back into the cavity between the container and the
fuselage skin where Mach stem shock waves would have been formed.
Evidence of rapid shattering was found in a region approximately bounded
by frames 700 & 720 and stringers 38L & 40L, together with the lap joint
at 39L.
The shattered fuselage skin would have taken a
significant time to move, relative to the timescales associated with the
primary shock wave propagation. Clear evidence of soot and small impact
craters were apparent on the internal surfaces of all fragments of
container and structure from the shatter zone, confirming that the this
material had not had time to move before it was hit by the cloud of
shrapnel, unburnt explosive residues and sooty combustion products
generated at the seat of the explosion.
Following immediately behind the primary shock
wave, a secondary high pressure wave – partly caused by reflections off
the baggage behind the explosive material but mainly by the general
pressure rise caused by the chemical conversion of solid explosive
material to high temperature gas – emerged from the container. The
effect of this second pressure front, which would have been more
sustained and spread over a much larger area, was to cause the fuselage
skin to stretch and blister outwards before bursting and petalling back
in a star-burst pattern, with rapidly running tear fractures propagating
away from a focus at the shatter zone. The release of stored energy as
the skin ruptured, combined with the outflow of high pressure gas
through the aperture, produced a characteristic curling of the skin
‘petals’ – even against the slipstream. For the most part, the skins
which petalled back in this manner were torn from the frames and
stringers, but the frames and stringers themselves were also fractured
and became separated from the rest of the structure, producing a very
large jagged hole some 5 feet longitudinally by 17 feet
circumferentially (upwards to a region just below the window belt and
downwards virtually to the centre line).
From this large jagged hole, three of the
fractures continued to propagate away from the hole instead of
terminating at the boundary. One fracture propagated longitudinally
rearwards as far as the wing cut-out and another forwards to station
480, creating a continuous longitudinal fracture some 43 feet in length.
A third fracture propagated circumferentially downwards along frame 740,
under the belly, and up the right side of the fuselage almost as far as
the window belt – a distance of approximately 23 feet.
These extended fractures all involved tearing or
related failure modes, sometimes exploiting rivet lines and tearing from
rivet hole to rivet hole, in other areas tearing along the full skin
section adjacent to rivet lines, but separate from them. Although the
fractures had, in part, followed lap joints, the actual failure modes
indicated that the joints themselves were not inherently weak, either as
design features or in respect of corrosion or the conditions of the
joints on this particular aircraft.
Note: The cold bond process carried out at
manufacture on the lap joints had areas of disbonding prior to the
accident. This disbonding is a known feature of early Boeing 747
aircraft which, by itself, does not detract from the structural
integrity of the hull. The cold bond adhesive was used to improve the
distribution of shear load across the joint, thus reducing shear
transfer via the fasteners and improving the resistance of the joint to
fatigue damage; the fasteners were designed to carry the full static
loading requirements of the joint without any contribution from the
adhesive. Thus, the loss of the cold bond integrity would only have been
significant if it had resulted in the growth of fatigue cracks, or
corrosion induced weaknesses, which had then been exploited by the
explosive forces. No evidence of fatigue cracking was found in the
bonded joints. Inter-surface corrosion was present on most lap joints
but only one very small region of corrosion had resulted in significant
material thinning; this was remote from the critical region and had not
played any part in the break-up.
The cracks propagating upwards as part of the
petalling process did not extend beyond the window line. The wreckage
evidence suggests that the vertical fractures merged, effectively
closing off the fracture path to produce a relatively clean bounding
edge to the upper section of the otherwise jagged hole produced by the
petalling process. There are at least two probable reasons for this.
Firstly the petalling fractures above the shattered zone did not
diverge, as they had tended to do elsewhere. Instead, it appears that a
large skin panel separated and peeled upwards very rapidly producing
tears at each side which ran upwards following almost parallel paths.
However, there are indications that by the time the fractures had run
several feet, the velocity of fracture had slowed sufficiently to allow
the free (forward) edge of the skin panel to overtake the fracture
fronts, as it flexed upwards, and forcibly strike the fuselage skin
above, producing clear witness marks on both items. Such a tearing
process, in which an approximately rectangular flap of skin is pulled
upwards away from the main skin panel, is likely to result in the
fractures merging. Secondly, this merging tendency would have been
reinforced in this particular instance by the stiff window belt ahead of
the fractures, which would have tended to turn the fractures towards the
horizontal.
It appears that the presence of this initial
(‘clean’) hole, together with the stiff window belt above, encouraged
other more slowly running tears to break into it, rather than
propagating outwards away from the main hole.
2.12.1.2 Critical crack considerations
The three very large tears extending beyond the
boundary of the petalled region resulted in a critical reduction of
fuselage structural integrity.
Calculations were carried out at the Royal
Aerospace Establishment to determine whether these fractures, growing
outwards from the boundary of the petalled hole, could have occurred
purely as a result of normal differential pressure loading of the
fuselage, or whether explosive forces were required in addition to the
pressurisation loads.
Preliminary calculations of critical crack
dimensions for a fuselage skin punctured by a 20 by 20 inches jagged
hole indicated that unstable crack growth would not have occurred unless
the skin stress had been substantially greater than the stress level due
to normal pressurisation loads alone. It was therefore clear that
explosive overpressure must have produced the gross enlargement of the
initially small shattered hole in the hull. Furthermore, it was apparent
from the degree of curling and petalling of the skin panels within the
star-burst region that this overpressure had been relatively long term,
compared with the shock wave overpressure which had produced the shatter
zone. A more refined analysis of critical crack growth parameters was
therefore carried out in which it was assumed that the long term
explosive overpressure was produced by the chemical conversion of solid
explosive material into high temperature gas.
An outline of the fracture propagation analysis is
given at Appendix D. This analysis, using theoretical fracture
mechanics, showed that, after the incident shock wave had produced the
shatter zone, significant explosive overpressure loads were needed to
drive the star-burst fractures out to the boundary of the petalled skin
zone. Thereafter, residual gas overpressure combined with fuselage
pressurisation loads were sufficient to produce the two major
longitudinal cracks and a single major circumferential crack, extending
from the window belt down to beyond the keel centreline.
2.12.1.3 Damage to the cabin floor structure
The floor beams in the region immediately above
the baggage container in which the explosive had detonated were
extensively broken, displaying clear indications of overload failure due
to buckling caused by localised upward loading of the floor structure.
No direct evidence of bruising was found on the
top panel of the container. It therefore appears that the container did
not itself impact the floor beams, but instead the floor immediately
above the container was broken through as a result of explosive
overpressure as gases emerged from the ruptured container and loaded the
floor panels. Data on floor strengths, provided by Boeing, indicated
that the cabin floor (with the CRAF modification) would fail at a
uniform static differential pressure of between 3.5 and 3.9 psi (high
pressure below the cabin floor), and that the floor panel to floor beam
attachments would not fail before the floor beams. Whilst there is no
direct evidence of the pressure loading on the floor structure
immediately following detonation, there can be no doubt that in the
region of station 700 it would have exceeded the ultimate failure load
by a large margin.
2.12.2 Indirect explosive damage (damage at remote
sites)
All of the damage considered in the foregoing
analysis, and the mechanisms giving rise to that damage, resulted from
the direct impact of explosive shock waves and/or the short-term
explosive overpressure on structure close to the source of the
explosion. However, there were several regions of skin separation at
sites remote from the explosion (see para 1.12.3.2) which were much more
difficult to understand. These remote sites formed islands of indirect
explosive damage separated from the direct damage by a sea of more
generalised structural failure characterised by the progressive
aerodynamic break-up of the weakened forward fuselage. All of these
remote damage sites were consistent with the impact of very localised
pressure impulses on the internal surfaces of the hull -effectively high
energy ‘pressure blows’ against the inner surfaces produced by explosive
shock waves and/or high pressure gas flows travelling through the
interior spaces of the hull.
The propagation of explosive shock waves and
supersonic gas flows within multiple, interlinking, cavities having
indeterminate energy absorption and reflection properties, and
ill-defined structural response, is extremely complex. Work has been
initiated in an attempt to produce a three-dimensional computer analysis
of the shock wave and supersonic flow propagation inside the fuselage,
but full theoretical analysis is beyond present resources.
Because of the complexity of the problem, the
following analysis will be restricted to a qualitative consideration of
the processes which were likely to have taken place. Whilst such an
approach is necessarily limited, it has identified a number of
propagation mechanisms which appear to have been of fundamental
importance to the break-up of Flight PA103, and which are likely to be
critical in any future incident involving the detonation of high
explosive inside an aircraft hull.
2.12.2.1 Shock wave propagation through internal
cavities
When Mach stem shocks are produced not only are
the shock pressures very high but they propagate at very high velocity
parallel to the reflecting surface. In the context of the lower fuselage
structure in the region of Mach stem formation, it can readily be seen
that the Mach stem will be perfectly orientated to enter the narrow
cavity formed between the outer skin and the cargo liner/containers,
bounded by the fuselage frames [Appendix B, Figure B-25]. This cavity
enables the Mach stem shock wave to propagate, without causing damage to
the walls (due to the relatively low pressure where the Mach stem sweeps
their surface), and reach regions of the fuselage remote from the source
of the explosion. Furthermore, energy losses in the cavity are likely to
be less than would occur in the ‘free’ propagation case, resulting in
the efficient transmission of explosive energy. The cavity would tend to
act like a ‘shock tube’, used for high speed aerodynamic research,
confining the shock wave and keeping it running along the cavity axis,
with losses being limited to kinetic heating due to friction at the
walls.
Paragraph 1.6.3 contains a general description of
the structural arrangements in the area of the cargo hold. Before
proceeding further and considering how the shock waves might have
propagated through this network of cavities, it should be pointed out
that the timescale associated with the propagation of the shock waves is
very short compared with the timescale associated with physical movement
and separation of skin and structure fractured or damaged by the shock.
Therefore, for the purpose of assessing the shock propagation through
the cavities, the explosive damage to the hull can be ignored and the
structure regarded as being intact. A further simplification can
usefully be made by considering the structure to be rigid. This
assumption would, if the analysis were quantitative, result in
over-estimations of the shock strengths. However, for the purposes of a
purely qualitative assessment, the assumption should be valid, in that
the general trends of behaviour should not be materially altered.
It has already been argued that the shock wave
emerging from the container was, in part, reflected back off the inner
surface of the fuselage skin, forming a Mach stem shock wave which would
then have tended to travel into the semi-circular lower lobe cavity. The
Mach stem waves would have propagated away through this cavity in two
directions:
(i) |
under the belly, between the
frames [Appendix B, Figure B-3, detail A], and |
(ii) |
up the left side, expanding into
the cavity formed by the longitudinal manifold chamber where it
joins the lower lobe cavity. |
As the shock waves travelled along the cavity,
little attenuation or other change of characteristic was likely to have
occurred until the shocks passed the entrances to other cavities, or
impinged upon projections and other local changes in the cavity. A
review of the literature dealing with propagation of blast waves within
such cavities provides useful insights into some of the physical
mechanisms involved.
As part of a research program carried out into the
design of ventilation systems for blast hardened installations intended
to survive the long duration blast waves following the detonation of
nuclear weapons, the propagation of blast waves along the primary
passages and into the side branches of ventilation ducts was studied.
The research showed that 90? bends in the ducts produced very little
attenuation of shock wave pressure; a series of six right angle bends
produced only a 30% pressure attenuation, together with an extension of
the shock duration. It is therefore evident that the attenuation of
shock waves propagating through the fuselage cavities, all of which were
short with hardly any right angle turns, would have been minimal.
It was also demonstrated that secondary shock
waves develop within the entrance to any side branch from the main duct,
produced by the interaction of the primary shock wave with the geometric
changes in the duct walls at the side-branch location. These secondary
shock waves interact as they propagate into the side branch, combining
together within a relatively short distance (typically 7 diameters) to
produce a single, plane shock wave travelling along the duct axis. In a
rigid, smooth walled structure, this mechanism produces secondary shock
overpressures in the side branch of between 30% and 50% of the value of
the primary shock, together with a corresponding attenuation of the
primary shock wave pressure by approximately 20% to 25%.
This potential for the splitting up and
re-transmission of shock wave energy within the lower hull cavities is
of extreme importance in the context of this accident. Though the
precise form of the interactions is too complex to predict
quantitatively, it is evident that the lower hull cavities will serve to
convey the overpressure efficiently to other parts of the aircraft.
Furthermore, the cavities are not of serial form, i.e. they do not
simply branch (and branch again) in a divergent manner, but instead form
a parallel network of short cavities which reconnect with each other at
many different points, principally along the crease beams. Thus,
considerable scope exists for: the additive recombination of blast waves
at cavity junctions; for the sustaining of the shock overpressure over a
greater time period; and, for the generation of multiple shocks produced
by the delay in shock propagation inherent in the different shock path
(i.e. cavity) lengths.
Whilst it has not been possible to find a specific
mechanism to explain the regions of localised skin separation and
peel-back (i.e. the ‘pressure blow’ regions referred to in para 2.12.2),
they were almost certainly the result of high intensity shock
overpressures produced locally in those regions as a result of the
additive recombination of shock waves transmitted through the lower hull
cavities. It is considered that the relatively close proximity of the
left side region of damage just below floor level at station 500,
[Appendix B, Figure B-19, region D] to the forward end of the cargo hold
may be significant insofar as the reflections back from the forward end
of the hold would have produced a local enhancement of the shock
overpressure. Similarly, ‘end blockage effects’ produced by the cargo
door frame might have been responsible for local enhancements in the
area of the belly skin separation and curl-back at station 560 [Appendix
B, Figure B-19 and B-20, region E].
The separation of the large section of upper
fuselage skin [Appendix B, Figure B-19 and B-20, detail B] was almost
certainly associated with a local overpressure in the side cavities
between the main deck window line and the upper deck floor, where the
cavity is effectively closed off. It is considered that the most
probable mechanism producing this region of impulse overpressure was a
reflection from the closed end of the cavity, possibly combined with
further secondary reflections from the window assembly, the whole being
driven by reflective overpressures at the forward end of the
longitudinal manifold cavity caused by the forward end of the cargo
hold. The local overpressure inside the sidewall cavity would have been
backed up by a general cabin overpressure resulting from the floor
breakthrough, giving rise to an increased pressure acting on the inner
face of the cabin side liner panels. This would have provided pseudo
mass to the panels, effectively preventing them from moving inwards and
allowing them to react the impulse pressure within the cavity, producing
the region of local high pressure evidenced by the region of quilting on
the skin panels [Appendix B, Figure B-19, region C].
2.12.2.2 Propagation of shock waves into the cabin
The design of the air-conditioning/depressurisation-venting
systems on the Boeing 747 (and on most other commercial aircraft) is
seen as a significant factor in the transmission of explosive energy, as
it provides a direct connection between the main passenger cabin and the
lower hull at the confluence of the lower hull cavities below the crease
beam. The floor level air conditioning vents along the length of the
cabin provided a series of apertures through which explosive shock
waves, propagating through the sub floor cavities, would have radiated
into the main cabin.
Once the shock waves entered the cabin space, the
form of propagation would have been significantly different from that
which occurred in the cavities in the lower hull. Again, the precise
form of such radiation cannot be predicted, but it is clear that the
energy would potentially have been high and there would also
(potentially) have been a large number of shock waves radiating into the
cabin, both from individual vents and in total, with further potential
to recombine additively or to ‘follow one another up’ producing, in
effect, sustained shock overpressures.
Within the cabin, the presence of hard,
reflective, surfaces are likely to have been significant. Again, the
precise way in which the shock waves interacted is vastly beyond the
scope of current analytical methods and computing power, but there
clearly was considerable potential for additive recombination of the
many different shock waves entering at different points along the cabin
and the reflected shock waves off hard surfaces in the cabin space, such
as the toilet and galley compartments and overhead lockers. These
recombination effects, though not understood, are known phenomena.
Appendix B, Figure B-26 shows how shock waves radiating from floor level
might have been reflected in such a way as produce shock loading on a
localised area of the pressure hull.
2.12.2.3 Supersonic gas flows
The gas produced by the explosive would have
resulted in a supersonic flow of very high pressure gas through the
structural cavities, which would have followed up closely behind the
shock waves. Whilst the physical mechanisms of propagation would have
been different from those of the shock wave, the end result would have
been similar, i.e. there would have been propagation via multiple,
linked paths, with potential for additive recombination and successive
pressure pulses resulting from differing path lengths. Essentially, the
shock waves are likely to have delivered initial ‘pressure blows’ which
would then have been followed up immediately by more sustained pressures
resulting from the high pressure supersonic gas flows.
2.13 Potential limitation of explosive damage
Quite clearly the detonation of high explosive
material anywhere on board an aircraft is potentially catastrophic and
the most effective means of protecting lives is to stop such material
entering the aircraft in the first place. However, it is recognised that
such risks cannot be eliminated entirely and it is therefore essential
that means are sought to reduce the vulnerability of commercial aircraft
structures to explosive damage.
The processes which take place when an explosive
detonates inside an aircraft fuselage are complex and, to a large
extent, fickle in terms of the precise manner in which the processes
occur. Furthermore, the potential variation in charge size, position
within the hull, and the nature of the materials in the immediate
vicinity of the charge (baggage etc) are such that it would be
unrealistic to expect to neutralise successfully the effect of every
potential explosive device likely to be placed on board an aircraft.
However, whilst the problem is intractable so far as a total solution is
concerned, it should be possible to limit the damage caused by an
explosive device inside a baggage container on a Boeing 747 or similar
aircraft to a degree which would allow the aircraft to land
successfully, albeit with severe local damage and perhaps resulting in
some loss of life or injuries.
In Appendix E the problem of reducing the
vulnerability of commercial aircraft to explosive damage is discussed,
both in general terms and in the context of aircraft of similar size and
form to the Boeing 747. In that discussion, those damage mechanisms
which appear to have contributed to the catastrophic structural failure
of Flight PA103 are identified and possible ways of reducing their
damaging effects are suggested. These suggestions are intended to
stimulate thought and discussion by manufacturers, airworthiness
authorities, and others having an interest in finding solutions to the
problem; they are intended to serve as a catalyst rather than to lay
claim to a definitive solution.
2.14 Summary
It was established that the detonation of an IED,
loaded in a luggage container positioned on the left side of the forward
cargo hold, directly caused the loss of the aircraft. The direct
explosive forces produced a large hole in the fuselage structure and
disrupted the main cabin floor. Major cracks continued to propagate from
the large hole under the influence of the service pressure differential.
The indirect explosive effects produced significant structural damage in
areas remote from the site of the explosion. The combined effect of the
direct and indirect explosive forces was to destroy the structural
integrity of the forward fuselage, allow the nose and flight deck area
to detach within a period of 2 to 3 seconds, and subsequently allow most
of the remaining aircraft to disintegrate while it was descending nearly
vertically from 19,000 to 9,000 feet.
The investigation has enabled a better
understanding to be gained of the explosive processes involved in such
an event and to suggest ways in which the effects of such an explosion
might be mitigated, both by changes to future design and also by
retrospective modification of aircraft. It is therefore recommended that
Regulatory Authorities and aircraft manufacturers undertake a systematic
study with a view to identifying measures that might mitigate the
effects of explosive devices and improve the tolerance of the aircraft
structure and systems to explosive damage.
3. CONCLUSIONS
(a) Findings
(i) |
The crew were properly licenced
and medically fit to conduct the flight. |
(ii) |
The aircraft had a valid
Certificate of Airworthiness and had been maintained in compliance
with the regulations. |
(iii) |
There was no evidence of any
defect or malfunction in the aircraft that could have caused or
contributed to the accident. |
(iv) |
The structure was in good
condition and the minimal areas of corrosion did not contribute to
the in-flight disintegration. |
(v) |
One minor fatigue crack
approximately 3 inches long was found in the fuselage skin but this
had not been exploited during the disintegration. |
(vi) |
An improvised explosive device
detonated in luggage container serial number AVE 4041 PA which had
been loaded at position 14L in the forward hold. This placed the
device approximately 25 inches inboard from the skin on the lower
left side of the fuselage at station 700. |
(vii) |
The analysis of the flight
recorders, using currently accepted techniques, did not reveal
positive evidence of an explosive event. |
(viii) |
The direct explosive forces
produced a large hole in the fuselage structure and disrupted the
main cabin floor. Major cracks continued to propagate from the large
hole under the influence of the service pressure differential. |
(ix) |
The indirect explosive effects
produced significant structural damage in areas remote from the site
of the explosion. |
(x) |
The combined effect of the direct
and indirect explosive forces was to destroy the structural
integrity of the forward fuselage. |
(xi) |
Containers and items of cargo
ejected from the fuselage aperture in the forward hold, together
with pieces of detached structure, collided with the empennage
severing most of the left tailplane, disrupting the outer half of
the right tailplane, and damaging the fin leading edge structure. |
(xii) |
The forward fuselage and flight
deck area separated from the remaining structure within a period of
2 to 3 seconds. |
(xiii) |
The No 3 engine detached when it
was hit by the separating forward fuselage. |
(xiv) |
Most of the remaining aircraft
disintegrated while it was descending nearly vertically from 19,000
to 9,000 feet. |
(xv) |
The wing impacted in the town of
Lockerbie producing a large crater and creating a fireball. |
(b) Cause
The in-flight disintegration of the aircraft was
caused by the detonation of an improvised explosive device located in a
baggage container positioned on the left side of the forward cargo hold
at aircraft station 700.
4. SAFETY RECOMMENDATIONS
The following Safety Recommendations were made
during the course of the investigation :
4.1 |
That manufacturers of existing
recorders which use buffering techniques give consideration to
making the buffers non-volatile, and the data recoverable after
power loss. |
4.2 |
That Airworthiness Authorities
re-consider the concept of allowing buffered data to be stored in a
volatile memory. |
4.3 |
That Airworthiness Authorities
consider requiring the CVR system to contain a short duration, i.e.
no greater than 1 minute, back-up power supply to enable the CVR to
respond to events that result in the almost immediate loss of the
aircraft’s electrical power supply. |
4.4 |
That the Department of Transport
fund a study to devise methods of recording violent positive and
negative pressure pulses, preferably utilising the aircraft’s flight
recorder systems. |
4.5 |
That Airworthiness Authorities
and aircraft manufacturers undertake a systematic study with a view
to identifying measures that might mitigate the effects of explosive
devices and improve the tolerance of aircraft structure and systems
to explosive damage. |
M M Charles
Inspector of Accidents
Department of Transport
July 1990
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