azza boeing 707-330c, st-akw-interim report

AIRCRAFT ACCIDENT INTERIM REPORT 10/2009 DATED JANUARY 20th

, 2011

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INTERIM

FACTUAL AIRCRAFT ACCIDENT REPORT 10/2009

General Civil Aviation Authority

Investigation and Regulation Section

Abu Dhabi, UAE

Boeing 707-330C, ST-AKW Near Sharjah International Airport

October 21st, 2009 United Arab Emirates


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OBJECTIVE

This investigation is performed in accordance with the UAE Federal Act No 20/1991, promulgating the Civil Aviation Law, Chapter VII, Aircraft Accidents, Article 48, and in conformity to ICAO Annex 13 to the Chicago Convention.

The sole objective of this investigation is to prevent aircraft accidents and incidents. It is not the purpose of this activity to apportion blame or liability.


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AIRCRAFT ACCIDENT BRIEF GCAA R&I Report No.: 10/2009 Operator: Sudan Airways (“lessee”) under a wet lease agreement with Azza Air

Transport Co. (“lessor”) Aircraft Type and Registration: Boeing 707-330C (Cargo), ST-AKW No. and Type of Engines: Four Pratt and Whitney JT3D-3B Turbofan Engines Date and Time (UTC): 21st October, 2009, 11:31 Location: 1.6 kilometres from the end of runway 30 (threshold of RWY 12)

Latitude 25 degrees 21 minutes 0.19 seconds North Longitude 55 degrees 29 minutes 33.98 seconds East

Type of Flight: Cargo Transport Persons on Board: 6 crewmembers Injuries: 6 Fatal Nature of Damage: Aircraft completely destroyed and consumed by fire

The accident, involving a Boeing 707-330C (cargo) aircraft, registration number ST-AKW, was notified to the General Civil Aviation Authority (GCAA), on October 21st, 2009 at about 1131 UTC. An Investigation Team was formed and reached the accident site within minutes after being notified by the Sharjah Airport and coordinated with all Authorities on site by initiating the accident investigation process according to the already prepared and exercised plans. The Regulation and Investigation Department (R&I) of the GCAA is leading the Investigation as the United Arab Emirates (“UAE”) is the State of Occurrence. Notes:

1 The word (“Aircraft”) in this report implies the accident aircraft.

2 The word (“Team”) in this report implies the Accident Investigation Team lead by an Investigator-In-Charge assigned by the GCAA of the UAE and encompassed investigators from the GCAA, accredited representative from Sudan Civil Aviation Authority and his advisor, and accredited representative from the National Transportation Safety Board (“NTSB”) of the United States of America and his technical advisors from the Federal Aviation Administration (“FAA”), the Boeing Company and Pratt and Whitney.

3 LT is the Local Time of the UAE which is +4 hours of the Coordinated Universal Time.

4 All directional references to front and rear, right and left, top and bottom, and clockwise and counterclockwise are made aft looking forward (ALF) as is the convention. The direction of rotation of the engine is clockwise and the propeller is counterclockwise. All numbering in the circumferential direction starts with the No. 1 position at the 12:00 o’clock position, or immediately clockwise from the 12:00 o’clock position and progresses sequentially clockwise ALF.


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ABBREVIATIONS AND DEFINITIONS USED IN THIS REPORT

AD Airworthiness Directive BOAS Blade Outer Air Seal

C.G. Centre of Gravity

C/O Carried Out

CPCP Corrosion Prevention and Control Programmes

CVR Cockpit Voice Recorder

CSN Cycles Since New

EPR Engine Pressure Ratio

E.W. Empty Weight

FDR Flight Data Recorder

GCAA General Civil Aviation Authority (UAE)

hrs Hours

ICAO The International Civil Aviation Organisation

ID Inside Diameter

IMC Intermediate Case

Investigation Team GCAA/the Regulation and Investigation Team

kt Knot(s)

LG Landing Gear

LE Leading Edge

L/H Left Hand

LT Local time of the United Arab Emirates

m Metre(s)

MAC Mean Aerodynamic Centre

MSN Manufacturer serial number MLG Main Landing Gear

N1 Identifies the low pressure rotor section of a jet engine and its speed in revolutions per minute is normally expressed as a percentage (%)

NLG Nose Landing Gear

NS Nacelle Station (Stations referring to a certain datum

identified along the aircraft in inches

No. Number

NRC Non-Routine Card

NTSB National Transportation Safety Board

OD Outside Diameter

OUTBD Outboard

PIC Pilot In Command of the accident Aircraft

P/N Part Number

PPC Pilot Proficiency Check

R/H Right Hand

RPM Revolution Per Minute

RWY Runway


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s Second(s)

S/N Serial Number

SSID Supplemental Structural Inspection Document

T/R Thrust Reverser

TSN Time Since New-flight hours

TWY Taxiway

UAE The United Arab Emirates

UTC Coordinated Universal Time

V2 Takeoff safety speed. The speed at which the aircraft may

safely become airborne with one engine inoperative.

V/C Visual Check


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TABLE OF CONTENTS

FACTUAL INFORMATION

History of the Flight 1

Injuries to Persons 1

Damage to Aircraft 1

Other Damage 1

Aircraft Information 1

Meteorological Information 8

Flight Recorders 9

Wreckage and Impact Information 10

Fire 10

Survival Aspects 11

Tests and Research 11

Ongoing Investigation Activities 19

Safety Concerns and Actions

19


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FACTUAL INFORMATION

The information contained in this interim report is derived from the factual information gathered during

the ongoing investigation of the occurrence. Later interim reports or final report may contain altered

information in case new evidences appear during the ongoing investigation that requires changes to the

information depicted in this report.

HISTORY OF FLIGHT

On October 21st, 2009, at approximately 1130 UTC, a Boeing 707-330C (cargo) Aircraft, registration number ST-AKW departed Sharjah International Airport with no reported mechanical anomalies. The Aircraft took off from runway 30 to Khartoum, Sudan, with a total of six persons onboard: three flight crewmembers (Captain, First Officer and Flight Engineer), a Ground Engineer and two Load Masters, along with cargo of air conditioning units, auto parts, computers, personal effects, and some tools.

After takeoff, and during initial climb, the cowling of engine No. 4 separated from the Aircraft and fell down onto the departure runway.

The Aircraft continued in a shallow climbing with level wings when the pilot informed the Air Traffic Control that he lost engine No. 4. A few seconds later, the Aircraft veered to the right in comparison to the initial takeoff flight path and finally went into a right turn associated with approximately 70 degrees right bank, dived and impacted the ground at approximately 1.6 kilometres from the end of the Sharjah Airport runway about 20 seconds after takeoff.

The high impact forces caused an explosion, subsequent fire, and completely destroyed the Aircraft.

INJURIES TO PERSONS

Injuries Flight Crew Cabin Crew Passengers Other Total

Fatal 3 - 3 - 6

Serious - - - - -

Minor - - - - -

None - - - - -

Total 3 - 3 - 6

DAMAGE TO AIRCRAFT

The Aircraft was destroyed due to significant impact forces and subsequent fire.

OTHER DAMAGE

None

AIRCRAFT INFORMATION

General Information


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Manufacturer Boeing Company

Type and model B707-330C (Cargo)

MSN 20123

Date of delivery 28/02/1969

Registration mark ST-AKW

TSN 77484 hrs as of 10/10/2009

CSN 26888 cycles as of 10/10/2009

Certificate of Airworthiness

Issuing Authority Sudan Civil Aviation Authority

Issue date September 10th

, 2008

Valid till February 24th

, 2010

Certificate of Registration

Issuing Authority Sudan Civil Aviation Authority

Issue date June 21st

, 1998

Last maintenance checks

Check B-Check

Date 25/07/2009

Time since last check 136 hrs as of 10/10/2009

Last heavy check C3-Check, complete on 02/02/2009

Engine Description

The JT3D-3B engine is a dual-spool, axial flow, low bypass, turbofan engine having a multistage split compressor, an eight can (can-annular) combustion chamber and a split four stage reaction-impulse turbine.

Empty weight 134, 094 lbs

E.W. C. G in % MAC 25.71%

Last Weight and Balance Done on 30/05/2009

Engines Four Pratt & Whitney JT3D-3B turbofan engines

S/N

Engine No. 1 668597

Engine No. 2 668411

Engine No. 3 644103

Engine No. 4 644495


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Figure 2- Thrust Reverser Controls

Figure 1- Thrust Reverser Controls

The inlet diameter of the engine is 53 inches and the engine itself is 136 inches long. The engine is flat-rated1 to 84 °F and the maximum thrust is 18,000 pounds.

The engine consists of a front compressor (Low Pressure Compressor “LPC”), rear compressor (High Pressure Compressor “HPC”), combustion section, rear compressor drive turbine (High Pressure Turbine “HPT”), front compressor drive turbine (Low Pressure Turbine “LPT”), and exhaust section. The front compressor contains two fan stages and six LPC stages. The rear compressor contains seven HPC stages and drives the accessory gearbox through a towershaft. The HPC 15th stage provides bleed air for aircraft use.

Each of the eight combustion cans is fed by a ring of six dual orifice fuel nozzles. The HPT is a single- stage that drives the rear compressor through the HPT drive shaft. The LPT is a three-stage that drives the front compressor through the LPT drive shaft. The engine exhaust section consists of a Turbine Exhaust Case “TEC” and tail plug and has provisions for the rear engine mount. Stage numbering convention in the compressor section is as follows: the fan stages are stages 1 and 2, the LPC stages are 4 – 9 and the HPC stages are 10 – 16. There is no stage designated as stage 3 in the compressor. Stage numbering convention in the turbine section is as follows: the HPT is stage 1 and the LPT is stages 2 – 4. Together the fan, LPC, and LPT are considered the low (N1) rotor, while the HPC and HPT are considered the high (N2) rotor.

The engine has eight bearings designated as No. 1, 2, 2½, 3, 4, 4½, 5, and 6. The No. 2 and 4 bearings are duplex ball construction and are the thrust bearings for the low and high rotors respectively. The No. 2, 2½, and 3 bearings share a common compartment in the Intermediate Case (IMC). The bearings are positioned on the shafts in such a way that both the low and high rotors are straddle mounted (bearings on the front and rear of each rotor).

An accessory gearbox, driven by the engine high rotor through a towershaft, has provisions for, among other things, the engine fuel pump, hydro-mechanical fuel control, hydraulic pump, air turbine starter, and alternator.

A nacelle provides an aerodynamic fairing around the outside of the engine. The nacelle consists of an

1 Flat-rated to a specific temperature indicates that the engine is capable of producing the rated power up to

the specific inlet temperature.


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inlet cowl, a L/H and R/H fan cowl, a L/H and R/H engine core cowl, and an aft thrust reverser outer sleeve. The fan and engine cowls are hinged at the pylon and latch on the bottom side of the engine and are capable of being opened for the purpose of performing maintenance on the engine.

Thrust Reversers

The thrust reversers on the JT3D engine consist of a forward fan thrust reverser, and an aft core thrust reverser.

Flow reversing components on the forward thrust reverser are located circumferentially around the first stage compressor case. During forward thrust operation, the exhaust air from the fan section of the first stage compressor is discharged in the aft direction from a duct created by the powerplant diaphragm and engine cowl ring assembly.

During reverse thrust operation, the forward thrust reverser actuators move the cowl ring aft and reposition the blocker doors, lower vane assemblies, and baffle assemblies to discharge the fan air in a forward direction.

The aft thrust reverser, during forward thrust operation, is an intermediate path for exhaust gas flow between the engine and tail pipe.

During reverser operation, the aft thrust reverser actuators move the aft thrust reverser sleeve rearward uncovering the cascade assemblies. Movement of the sleeve causes the clamshell doors to close through the action of a hinge drive mechanism connecting the sleeve and clamshell door hinge arm. Engine exhaust gases are diverted through the cascade vane assemblies.

The T/R control system directs pneumatic pressure to actuators of the forward and aft thrust reversers which position the flow reversing components. The system consists of the reverse thrust controls mounted on the thrust control shaft and on the engine controls strut bracket, forward thrust reverser follow-up linkage, aft reverser follow-up linkage, hinge-drive mechanism, forward and aft actuators, and miscellaneous system components (Figure 1).

The reverse thrust controls consist of a thrust reverser directional control valve, a rocker arm mechanism, a thrust lever actuating cam, forward thrust reverser interlock cam, aft thrust reverser interlock cam, aft thrust reverser follow-up pulley, forward thrust reverser follow-up crank, and two concentric shafts on which the interlock cams, follow-up crank, and follow-up pulley are mounted. The thrust lever actuating shaft is installed on the lower end of the thrust control shaft. The other reverse thrust controls are installed on the engine controls strut bracket, which is mounted over the lower end of the throttle and start shafts (Figure 2).

An interlock feature in the control system prevents the application of full forward, or full reverse, when either the forward or aft thrust reverser is not nearly in the commanded position. A forward thrust reverser interlock cam and an aft thrust reverser interlock cam limit rotation of the thrust control shaft to partial power until follow-up linkages connecting the cams to the forward thrust reverser cowl ring and aft reverser sleeve reposition the cams to allow full forward or reverse operation. An over-ride provision on the forward thrust fan-reverser interlock allows increased thrust to be commanded by overcoming a spring load

Engine No. 4 core T/R stow pneumatic line

Deploy line

Pt7 manifold line on its route to the engine pressure ratio transmitter

Figure 3- A photo taken on a sister aircraft showing the hinge support structure at NS 198.82 mounted on the pylon of engine No. 4.


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if additional bleed pressure is required to move the fan thrust reverser to the closed (“cruise”) position.

The thrust reverser directional control valve is a simple, two position valve that receives its position input from the thrust-lever actuating cam to provide the pneumatic signal for thrust reverser actuation.

The directional valve is positioned by a thrust-lever actuating cam through a rocker arm mechanism to route engine bleed air through either of two ports on the valve to the reverser or stow (“cruise”) pneumatic manifolds.

The thrust reverser directional control valve actuating cam is mounted on the thrust control shaft in the engine strut.

The valve assembly is on the engine controls strut bracket. The locking cams, follow-up crank and follow-up pulley are mounted on a shaft through the strut bracket (Figure 3).

The forward fan thrust reverser actuators operate the cowl ring assembly, blocker doors, and lower vane assemblies. There are fourteen actuators located circumferentially around the engine. Twelve of these actuators are identical and operate the blocker doors and the baffle assembly. The other actuators are smaller, have a shorter stroke and operate the vane assemblies. During forward thrust operation, pneumatic pressure enters the rod ports of the actuators pulling the blocker doors forward and out of the fan air duct to lay flat around the compressor case. The vane assemblies move forward also. The cowl ring assembly is repositioned forward to exhaust the fan air in the aft direction. For reverse thrust operation, air enters the head ports of the actuators to move the cowl ring aft, and reposition the blocker doors, vane assemblies and baffle assemblies to discharge the fan exhaust in the forward direction.

Two dual-cylinder pneumatic actuator assemblies operate the aft (core) thrust reverser. One of the actuator assemblies is mounted on the top of the thrust reverser frame. The other actuator assembly is mounted underneath the engine, just forward of the aft thrust reverser.

Each of the actuator assemblies includes two side-by-side joined pneumatic cylinders. The lower actuator assembly also includes a lock actuator and hook-type lock to prevent inflight deployment of the aft thrust reverser sleeve with the engine shutdown (no pneumatic pressure). This lock actuator, mounted below the dual-cylinder assembly, contains a spring loaded (toward lock) piston with the rod connected to the lock hook. When reverse thrust is commanded; the lower lock actuator receives pneumatic pressure to actuate, disengaging the lock and uncovering a port in the lock actuator to route the pneumatic air to the head ports of the upper and lower thrust reverser actuators causing reverse thrust actuation. During forward thrust operation, the four primary actuators are pressurized to the stowed position (rod end port), and the lock cylinder is pressurized to lock.

The core reverser consists of a left and right hand inner clam shell door halves that, when actuated, rotate into the core flow streaming out the turbine exhaust case. The inner clamshells are mechanically coupled to the outer translating sleeve and, when deployed, redirect airflow from the core of the engine through cascades that are exposed when the outer translating sleeve is in the aft position (deployed).

A position indicating switch senses if the core reverser is in the stowed or unlocked position. The position indicating switch is wired to an indicator light in the cockpit.


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Engine Pressure Ratio Indicating System

The engine pressure ration (EPR) indicating system provides the ratio of exhaust total pressure to the inlet total pressure (Pt7/Pt2) for each engine to the flight crew on the engine instrument panel. EPR is the primary parameter used to quantify engine thrust (power setting) for the JT3D engine.

The system for each engine consists of six exhaust (Pt7) sensing probes around the periphery of the engine exhaust case, one inlet pressure (Pt2) probe on the right hand side of the nacelle strut, a pressure ratio transmitter mounted in the nacelle strut and an indicator on the engine instrument panel in the cockpit.

Engine exhaust pressure (Pt7) is detected by six probes extending into the engine exhaust stream. These probes are connected to a common manifold.

Engine inlet pressure (Pt2) is sensed by a probe similar to a pitot tube. This probe is mounted on the right hand side of the nacelle strut so that the open end of the tube faces the air stream. The probe is heated by nose cowl anti-icing air when the engine anti-icing system is in operation.

The exhaust pressure sensing manifold is made up of two segments of tubing mounted around the periphery of the exhaust case. Three Pt7 sensing probes are connected to each manifold sections. The manifold assembly averages the pressure sensed by the probes.

The engine pressure ratio transmitter converts the exhaust pressure (Pt7) and the inlet pressure (Pt2) into a ratio, and generates three-phase electrical signals corresponding to pressure changes in the engine. The EPR transmitter consists of two bellows (multicell diaphragms), a sensing mechanism, an amplifier, a motor-gear train, and a synchronous generator. The EPR transmitter is mounted in the centre section of the nacelle strut. On engine No. 4 installations, the flex line from the Pt7 manifold is just aft of NS 198.82 hinge support structure (Figure 3).

The Pt7 and Pt2 pressures are applied to either side of the bellows assembly of the transmitter. A change in either of these pressures cause differential bellows movement. The bellows movement effects the sensing mechanism which, with the aid of the amplifiers and the motor-gear train, causes the generator rotor to rotate and generate three-phase electrical signals.

The engine pressure ratio indicator is located on the engine instrument panel. It contains a synchronous receiver which is actuated by the electrical signal received from the engine pressure ratio transmitter. The indicator shows the ratio between the exhaust and inlet pressures (Pt7 and Pt2).

Maintenance Records

Approved Maintenance Schedule

The Approved Maintenance Schedule “AMS” No. AZ/AMS/B.707/01 was approved by letter Reference No. CAA/7/AW/ENO/AZZA AIR/B.707 dated 04/09/2008 issued by Airworthiness Directorate of the Sudanese Civil Aviation Authority.

According to the AMS, a Pre-flight Check is to be carried out prior to first flight of the day, Transit Check is to be carried out prior to every flight, A-, B- and C-Checks are to be done on times not to exceed 30 days, 120 days, and 12 months, respectively.

In addition, Structural Inspections were to be performed according to Boeing Document D6-7552, Supplemental Structural Inspection Document (SSID) according to Document 44860, Corrosion Inspections and Aging Inspections are to be done according to Documents D6-54928 and D6-54996, respectively.


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Last C3-Check 2

The Investigation Team focused it’s attention to the maintenance records relating to engine No. 4, the engine No. 4 cowls and the No. 4 thrust reverser due to facts that the engine No. 4 cowls departed the Aircraft shortly after takeoff, the pilot reported that he lost engine No. 4, and that the Aircraft rolled sharply to the right into the reported lost engine.

A Certificate of Release to Service (CRS) was issued by the Egyptian Company for Aircraft Maintenance “ECAM” for the Aircraft on February 2nd, 2009. The CRS showed that in addition to the C3-Check, maintenance was performed on Corrosion Prevention and Control Programme “CPCP”, Airworthiness Directives “AD” and Non-Routine Cards ”NRC”.

Table 1 below shows Non-routine cards “NRCs” related to engine No. 4 cowls and T/R. The table also illustrates the maintenance corrective action for each NRC. Those NRCs were included in the NCR index contained in the work order that was submitted by Azza Air Transport to ECAM:

Table 1- C-Check NRCs 3

NRC S/N Discrepancy/Customer

Request Maintenance Corrective Action Date

011 All Eng T/R to be checked Engine No. 4 aft and engine No. 2 wire repair check during engine G.R [ground run] O.K.

January 20th

2009

012 All eng. T/R to be check All T/R thoroughly checked, cleaned and lubricated carried out tested with external pneumatic pressure and also during engines ground run found operating normally and satisfactory

Not recorded

081 Pls check engine No. 4 cowl very difficult to open and close

Engine No. 4 cowl found slightly twisted and need to be adjusted. Repaired carried out.

October 15th

2008

Table 2 shows the Routine Inspection Cards as contained in the work order handed by Azza Air Transport to ECAM:

Table 2- C-Check Job Instruction Cards Job Card

No. Job Instruction Action [taken] Date

015 Check the following (V/C for 4 Engines) Left and right engine cowl panels

Left & right engines cowl V/C carried out refer to N.R.C No. 081

October 15th

2008

2 Last C3-check was carried out at the Egyptian Company for Aircraft Maintenance “ECAM” which place of

maintenance facilities is Sherm El-Shaikh, Egypt. A maintenance contract agreement No. ECAM/AZZA Company/001/B707-30C was signed by Azza Air Transport and ECAM in August 2008. A term related to the last C3-Check was contained in that agreement.

3 The words in this table are written in the same language of the pertinent document.


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249

Check the following (V/C) Engine cowling and panels

A. Side cowl panels. Panel hinge fittings “F

2. u-bolts 3. support rods

B. Fan cowl panels 1. Panel hinge fittings 2. bolts 3. panel skin

All above items V/C carried out O.K

October 16th

2008

275

Check the following (V/C) Thrust reverser A. Cowl ring assembly B. Blocker doors C. Cascade vane assemblies D. Track and carriage assemblies E. Aft T/R sleeve F. Aft T/R Exhaust plug

All the items above V/C carried out O.K.

October 30th

2008

276

Check the following (V/C) T/R control system

A. T/R directional control valve B. T/R locking cam C. T/R rocker arm shaft control D. T/R forward follow-up linkage E. T/R aft follow-up linkage F. T/R directional control valve filter (Clean)

V/C carried out for items above O.K.

October 31st

2008

Last B-Check

The last B-Check was carried at Azza Air Transport maintenance facilities and completed on 25/07/2009. The Check was conducted by ECAM staff and a CRS was issued on the same date.

Table 3 below shows, the Routine Inspection Cards as contained in the work order package, no NRCs were raised in that B-Check.

In addition to the B-Check package, the work order contained CPCP, AD and NRC works.

Table 3- Routine Inspection Cards in the last B-Check Package Job Card

No. Job Instruction Action [taken] Date

5

Eng. No. 4 Check: L/R [L/H and R/H] engine cowl panel, hook latch fasteners cowl panel support rod for condition, missing items & security

Checked & necessary repaired C/O. July 19th 2009

METEOROLOGICAL INFORMATION

There was no significant weather at the time of the accident.


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FLIGHT RECORDERS 4

The investigation team located and recovered the CVR away from its installed location on the Aircraft on the day of the accident. The FDR was still installed in the Aircraft when it was found two days after the accident.5

The Aircraft was equipped with a Sundstrand Data Control, Inc. model FA-542 (300-hour scratch foil) FDR (P/N 101035-1, S/N 1598), and a Fairchild model A100A (30-minute continuous tape) CVR (P/N 93-A100-80, S/N 54853).

CVR Examination

Externally, the CVR displayed extensive impact damage and partial fire and smoke soot. The outside cover was cut off the unit to gain access to the inside crash case and the tape storage reel.

The crash case was intact but showed evidence of fire and smoke damage. Removal of the crash case exposed the fire protection case, which showed no external evidence of fire or smoke damage. Similarly, there was no evidence of fire or smoke damage to the tape transport (Figure 4).

The tape was found on the reel but was not intact. The end of the tape that feeds out from the centre of the reel, adjacent to the rotating hub, had been tucked back into the centre of the reel, while the end of the tape from the outside of the reel was still within the confines of the reel. Thus, no tape was present along the tape path or over the recording heads. Figure 4- tape transport

Also, it was noted that the ends of the tape did not match each other (Figure 5). No other undamaged or usable segments of recording tape were found in the CVR.

On the subsequent playback of the tape, the recording, which was found to be unrelated to the accident flight, ended after 23 minutes and 46 seconds, indicating that at least 6 minutes and 15

seconds of recording, or 701.25 inches (1,781 centimetres) of tape, was missing. 6

4 The use of engraving metal foil FDRs was discontinued on 1 January 1995 (ICAO Annex 6 to the Convention on

International Civil Aviation).

5 Both flight recorders were kept under the custody of the investigation team, and later on were sent to the Air Accidents Investigation Branch (AAIB) of the United Kingdom on November 2

nd, 2009 for examination.

6 The CVR operates at a tape speed of tape speed is 1.875 inches/second and minimum duration of the

recording is 30 minutes.


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FDR Examination

The FDR was still within its cylindrical housing which was sealed at both ends. Externally, the housing displayed evidence of impact and extensive fire damage. Internally, the housing and FDR displayed evidence of heat damage (Figure 6).

The FDR was removed from the housing and the “HOURS REMAINING” indicator read zero hours.

It was also observed that the round, tamper-evident maintenance seals (one on top of the FDR and the other on the back) were broken. The top seal was also partially stuck over a Hunting service sticker with the year 1996 printed on it.

The foil cassette (Figure 7) was then removed from the FDR, and on inspection, showed no signs of damage. There was no foil on the supply spool (R/H spool in Figure 7) with all of the used foil on the take-up spool (L/H spool in Figure 7). An examination of the foil showed that it had been reused on numerous occasions. Given that all of the foil was located on the take-up spool, it is very unlikely that the FDR was recording at the time of the accident.

WRECKAGE AND IMPACT INFORMATION

The Aircraft was found completely destroyed, burnt, and scattered over an area of approximately 0.25 square kilometres.

The tail, wings, and fuselage were completely destroyed and burnt. The MLG and the NLG were also scattered and damaged.

There were two ground impact marks each was about 10 metres long, 3 metres wide and 1.5 metres deep. The majority of the wreckage pieces settled to the east of a service car road in the vicinity of the accident site. Engines No. 3 and No. 4 were found to the L/H side of the R/H wing with the largest fuselage wreckage piece. Engines No. 1 and No. 2 were found to the R/H side of the largest fuselage piece.

Figure 5- CVR tape ends

Figure 6- FDR

Figure 8- Engine No. 4 core T/R position as found at the accident site

Figure 7- FDR foil cassette


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Engine No. 4 found with a deployed core T/R. The outer translating sleeve was proximate to its full deployment position (Figure 8). The L/H T/R clamshell was found at stowed position whereas the R/H T/R clamshell was found in almost at its full deployment position.

FIRE There were no evidence of inflight fire. The Aircraft burnt due to the post impact fire and the significant amount of fuel onboard contributed to the amount of fire damage.

SURVIVAL ASPECTS

This accident was non-survivable.

TESTS AND RESEARCH

Engine No. 4 Examination7

The objective of the examination was to document the condition of the engine and to determine if the engine was operating at the time of impact.

The initial visual examination was conducted at the Sharjah International Airport in December 2009 followed by a detailed engine disassembly conducted at the Aviation Engine Services facility in Miami, Florida, USA in August 2010. The details of engine disassembly are below.

Engine General Condition

The fan and LPC hardware separated during the accident impact sequence and was not included with the engine hardware exposed to the engine examination.

On-scene documentation of the fan indicated that both the first and second stage disks were intact. Some blade slots were empty, however, those blades that remained were all fractured transversely at or near the blade platform. Those blades with some airfoil material remaining were bent in the direction opposite of rotation. On-scene documentation of the LPC indicated that all visible blade slots of the 6-9th stages’ disks were empty, had blades fractured at the platforms, or had full length blades that were bent in the direction opposite rotation. All the stator vanes that were visible were bent in the direction of rotation and exhibited trailing edge damage.

The engine No. 4 structure was complete from the front of the HPC (10th stage) through the exhaust nozzle. None of the cases from the diffuser case back to the Turbine Exhaust Case (TEC) exhibited any breaches or indications of external fire. The rear 14 inches of the intermediate case “IMC” rear shirt was still attached and no indications of any breaches or indications of external fire were noted. The majority of the externals, ancillary components, and nacelle (with the exception of the core thrust reverser) had separated from the engine during the crash sequence.

All the Pt7 and Exhaust Gas Temperature (EGT) probes (6 for each) were intact with no notable damage. The EGT harness appeared intact and undamaged and the Pt7 manifold was intact but bent and distorted in the forward direction from the 11:00 to 1:30 o’clock position. The Pt7 signal tap line was bent and distorted inboards (left) and was separated at the flex line-to-hard line connection (Figure 9).

7 The engine was almost completely disassembled and the parts were visually inspected. The combustion

chamber outer case was cut by cutoff wheel to get access to the combustion chamber and the adjacent parts.

Figure 9- Pt7 manifold


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Engine No. 4 Examination Observations

In addition to exhibited marks of impact and overload failures, noticeable circumferential rubbing marks generated by metal-to-metal contact were observed on the following engine rotating and stationary parts:

1. The LE of LPT 4th stage blades airfoil at the OD.

2. All of the 4th blade tip shrouds at their OD surfaces. The forward Knife Edge (KE) was heavily rubbed and worn down. The rear KE also exhibited rolling and contact damage.

3. Two distinct rubs in the 1:00 to 3:00 and 7:00 to 9:00 o’clock locations of the LPT 4th stage.

4. All the LPT 4th vane airfoils on the ID LE. Approximately half of the vanes exhibited rubbing on the airfoil OD LE. Also there was a shiny rubbed area of the TE end of the OD shroud that was uniform around the circumference with the exception of the 3:30 - 5:00 o’clock location where the rub became less severe and was sporadic.

5. A 360-degree fresh heavy rub at the LPT 3-4 spacer.

6. The front and aft face of the LPT 3rd stage blades tip shrouds.

7. The LPT 3rd stage vanes airfoil OD LE, over the majority of the circumference.

8. The LPT 3rd stage vane inner support, light intermittent 360-degree rubbing on the aft face just behind the ID of the vanes. The 2-3 KE seal land exhibited a shiny rub over an estimated 75% of the circumference that was out of plane.

9. The forward KE of the LPT 2-3 spacer, a fresh rub over approximately 75% of its circumference. There were sporadic indications of a light rub on the aft KE.

10. Intermittent circumferential rub contact on the front face of the OD shrouds of the 2nd stage blades, similar contact was noted on the front and aft faces of the ID flowpath platforms. Uniform circumferential rub contact around on the tips of both shroud KE seals and on the aft face of the tip shrouds.

11. The 2nd stage vane inner support knife edge rotating seal land, which is part of the LPT shaft, exhibited 360 degree circumferential scoring.

12. A 360 degree circumferential sporadic rub on the LPT 2nd stage BOAS.

13. Intermittent rubbing on the 2nd stage vanes ID of flowpath platforms.

14. Sporadic rubbing at the entire circumference of the TE aft face of the HPT blades ID flowpath platforms. The TE OD aft face of the tip shrouds exhibited uniform circumferential rubbing. Many of the blades exhibited rubbing on the outer 2/3 span of the airfoil TEs.

15. The aft second stage inner vane support knife edge seal land exhibited uniform circumferential wear. Sporadic rubbing noted on the front face of inner flowpath platforms over the entire circumference. The OD of the rotating seal land on the outer front side of the HPT disk exhibited uniform 360 degree rubbing.

16. A uniform 360 degree rubbing at the HPT BOAS.

17. The LPT shaft exhibited numerous circumferential scoring areas (Figure 10).

Figure 10- LPT shaft


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Figure 12- The face of engine No. 4 N1 %RPM gage as removed from the panel.

18. The HPC module was all corn-crobbed with many empty blade slots. In the slots with blades remaining, the airfoils were fractured at the platform with some blades roots push aft in their respective blade slot. Sporadic groups of stator vanes in various stages were all bent in the direction of rotation. Many fragments of battered HPC stators, blades and shroud material were found in the combustion chamber.

19. Removal of the HPC module revealed that the HPC rear hub had fractured at the transition radius from the web to the shaft.

Instruments in the Pilot’s Centre Engine Instrument Panel 8

The objective of the examination was to document the engine instrument panel and possibly determine the engine power setting or the operability of each of the engines at the time of impact since no usable data was available on the FDR.

Panel’s General Condition

The recovered instrument panel revealed nine gauges that were identified by visible characters on their faces and by comparing it to the schematic of the engine instrument panel in the Boeing 707 operations manual, panel configuration in section C (Figure 11). There were also three empty instrument cases attached to the panel.

Engine No. 4, N1 %RPM Gage

Examination revealed that the hub of the centre needle remained attached to the gage but the pointer was missing (Blue arrow in Figure 12). The inset needle was intact and was pointing between 5 and 6 (Yellow arrow in Figure 12). Closer examination of the centre needle hub revealed that the needle was made of plastic. A red line was drawn from the hub centre in the direction of the fractured needle end. This red line corresponds to a gage reading of 92% N1 (Red arrow in Figure 12).

Manipulation of the needle hub revealed that it did not move and manipulation of the brass plate revealed that it did not rotate. Examination of the dial under ultra violet light revealed no indications of fluorescence.

Engine No. 4, Engine PRESS RATIO Gage

The as-removed engine No. 4 EPR gage from the instrument panel is illustrated on the left hand side of

8 The pilot’s centre panel was examined in the labs of the NTSB, Washington DC, USA.

Figure 11- The instruments panel


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Figure 13- The face of the engine No. 4 PRESS RATIO gauge as removed from the panel (left) and after cleaning (right).l.

Figure 14- Close up view of the needle imprint observed on the engine No. 4 PRESS RATIO gauge

Figure 13. Examination revealed that the inner glass cover was intact with only two pieces of the outer glass cover, indicated by the red arrows, still attached. The barely visible needle under the inner glass cover is indicated by the blue arrow and was estimated to be pointing in the vicinity of 2.7 EPR (left image in Figure 13). Due to the damage to the instrument case, the bezel could not be removed so the case was cut adjacent to the bezel. The bezel, with the glass, was folded upwards as illustrated in the right hand side of Figure 13 to reveal the dial. During the cutting operation, the needle moved and indicated a new value of 2.28 EPR. Examination of the dial under ultra violet light revealed that the last digit indicated by the yellow arrow fluoresced and no other fluorescence was observed.

Microscopic examination of the dial revealed a band of disturbed paint at the 1.05 EPR location contained within the red box in the right image in Figure 13. The area within the red box is illustrated in Figure 14 with the disturbed paint indicated by the red arrow. The needle was moved to the 1.06 position in order to illustrate that the needle tip indicated by the green arrow matches the outer end of the disturbed area indicated by the red arrow. It was noted that the edge of the disturbed area matched the adjacent edge of the needle and that the intensity of the disturbed area decreased as the distance from the tip increased, consistent with the needle impacting the dial surface during the impact sequence.

The number indicated in the window containing the fluorescing digit was found to agree with the location of the peripheral pointer indicated by the green arrow (left hand side of Figure

13).

Engine No. 4, N2 %RPM Gage

Examination of the dial under ultra violet light revealed no indications of fluorescence.

Microscopic examination of the dial revealed no marks or paint deposits that may have been produced by the needle impacting the dial.


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Figure 15- Engine No. core T/R sleeve and cascades

Figure 16- R/H Clamshell Figure 17- L/H Clamshell

Engine No. 4 Core T/R Examination 9

The entire core T/R was pushed outboards and upwards. The outer translating sleeve was intact and at the almost fully deployed position and the cascades exposed. (Figure 15)

The R/H T/R clamshell was found in the almost fully deployed position. The LE exhibited multiple tears, impact marks, and buckled material with some of the impact marks in-line with the cascade forward end attachment bolts. The Trailing Edge (TE) lip of the R/H clamshell exhibited localized areas with a buckled appearance characterized by kinked/wavy material deformation. There was no missing material or tears noted on the TE of the right-hand clamshell. Numerous hairline cracks were found in what appeared to be weld repaired areas of the R/H clamshell. The skin of the R/H door exhibited two compression buckles orthogonal to the centreline of the engine. These buckles were near the LE of the door at the 3 o’clock location on the door. Figure 16 depicts details of the right-hand clamshell door.

The aft seal assembly of the R/H clamshell half was damaged and partially missing. The aft seal assembly is attached to the outer surface of the clamshell half at its trailing edge. The seal’s leaves provide a seal between the clamshell half and the turbine inner sleeve when in the stow position. The seal leaves in the bottom 1/3 of the seal assembly were pushed up and away from the seal retainer. The leaves in the centre 1/2 the seal assembly were completely missing and the seal retainer was deformed forward. The upper 1/3 of the seal assembly appeared to be intact.

The L/H T/R clamshell was found beyond the normally stowed position. A portion of the LE was found riding over the aft T/R forward seal (normally the LE tucks under the aft T/R forward seal in the fully stowed position).

9 Core T/R examination was performed during the engine examination. The T/R sliding sleeve as well as

actuating cylinders, driving mechanism and cascades were all inspected and documented. The clamshells status was the only part contained in this report since actuating cylinders and some drive mechanisms were submitted to Boeing for future examination to determine if the T/R was stowed or deployed at the time of impact.

Buckles

Kinked/wavy material deformation

Buckles

Kinked/wavy material deformation

Translating sleeve

Cascades


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The TE lip of the L/H clamshell exhibited localized areas with a buckled appearance characterized by kinked/wavy material deformation.

There was no missing material or tears noted on the TE of the L/H clamshell. Numerous hairline cracks were found in what appeared to be weld repaired areas of the L/H clamshell. The skin of the L/H-hand door exhibited two compression buckles orthogonal to the centreline of the engine. These buckles were near the LE of the door at the 9 o’clock location on the door. Figure 17 depicts details of the left-hand clamshell door.

The aft lip of the T/R forward seal was deformed/crushed/buckled in the axial direction consistent with impact from the rear. This damage was most prevalent in the 2:00-4:00 o’clock and 8:00-10:00 o’clock areas of the seal.

Engine No. 4 Core Cowls and Cowls’ Support Structure Examination 10

Core Cowls

The L/H core cowl has six hinge fittings, four -1 and two -2. The R/H core cowl has five hinge fittings, three -1 and two -2. This configuration is per the production diagrams for the respective core cowls.

In L/H core cowl, the -1 hinge fitting at Nacelle Station (NS) 128.22 was relatively straight but angled aft approximately -8°. The -2 hinge at (NS) 139.87 had fractured away in the middle of the hinge, the hinge itself was angled aft approximately -50°. The -1 hinge at (NS) 151.55 is relatively straight but angled forward approximately +2°. The -1 hinge fitting at NS 174.46 was damaged and bent aft approximately -40° and the seal bracket was also bent in the aft direction. The -2 hinge fitting at NS 198.82 was bent aft approximately -40°. The -1 hinge fitting at (NS) 211.00 was bent aft approximately -2°. The cowl was damaged, torn and folded forward at the aft inboard corner at (NS) 221.25.

In the R/H cowl, the -1 hinge fitting at (NS) 128.22 was bent forward approximately +40°. According to the production drawing, there is no hinge fitting at (NS) 139.87. The -2 hinge fitting at (NS) 151.55 was bent forward approximately +15°. The -1 hinge fitting at (NS) 174.46 was bent slightly forward, approximately +5°. The -2 hinge fitting at (NS) 198.82 was bent slightly forward approximately +10°. The -1 hinge fitting at (NS) 211.00 was bent slightly forward approximately +5°. The cowl was damaged at the upper edge with multiple holes along the aft edge (NS) 221.25.

There are six latch assemblies at the latch line. The latch hooks are common to the R/H core cowl and the receptacles are common to the L/H core cowl. The latch hooks are spring loaded to engage the receptacle when the cowls are aligned and pushed close.

The latch hook in the R/H cowl at NS 128.22 was attached with a spring load. The latch hook at NS 139.87 was attached with no spring load. The latch hook at NS 151.55 was attached with a light spring load. The latch hook at NS 174.46 was attached and binding and the seal is broken in the upper section of the R/H cowl. The latch hook at NS 198.82 was attached with no spring load including severe damage

10 The cowls’ examination observations elaborated in this Interim Report are based on the field visual inspection

more detailed metallurgical testing is one of the ongoing investigation at Boeing labs.

Figure 18- Engine No. 4 cowls


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to the cowl frame near the hook from the alignment pin on the left cowl. The latch hook at NS 211.00 was attached with a spring load.

For the latch line in the left core cowl, the receptacle at NS 128.22 was attached and not damaged. The alignment pin was bent down. The receptacle at NS 139.87 was pulled and separated from the surrounding structure. The receptacle at NS 151.55 was attached and not damaged. The alignment pin was in place. The receptacle at NS 174.46 was attached and the alignment pin was loose and bent down. The two receptacles from NS 198.82 through NS 221.25 the frame structure was pulled-out and in photos provided by the GCAA the tear-out section was attached to the latch hooks on the right core cowl when the cowl was found on RWY 30.

Cowls’ Hinge Support Structure11

The cowl hinge support structure was found near the engine No. 4 cowls on the departure RWY 30. Comparing this part to the drawing for the Boeing 707 No. 4 pylon structure confirmed that this cowl hinge support structure was part of the No. 4 pylon structure.

The visual examination of the hinge support structure at NS 198.82 revealed contact markings on the top aft side of the hinge consistent with contact with the extend/retract pneumatic lines for the T/R and potential interference with the Pt7 sense line during separation from the pylon (Figure 19 “a” and “b”).

Preliminary Simulation with Radar Data 12

The objective of the simulation was recreating the flight in two different cases:

1. Typical engine (No. 4) failure with a loss of thrust.

2. Rapid transition to reverse thrust (both fan and core reversers) while at high power.

11

The cowls’ hinge support structure are under the ongoing investigation by metallurgical and metal-transfer analysis which are performed at Boeing labs. The examination is aimed to determine if a material was transferred from the Pt7 manifold tube. 12

Engineering simulation 707-300C was performed at Boeing facilities based after “initial” site observations of

T/R deployment and the PIC reporting of engine No. 4 loss. The mathematical models of the simulation represent what was provided for crew training simulators. The Aerodynamic model reflects what was published in a Boeing document, “Improved Aerodynamic Data for the 707-300B (ADV)/C Flight Simulator” released on 2/17/1977. Further simulation presuming more scenarios are within the ongoing investigation at Boeing facilities.

Figure 19a- cowl support structure overview Figure 19b- top side

AFT


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The data sources were limited to the following:

1. Six radar hits obtained from the tower surveillance radar.

2. Ground speed

3. Altitude

4. GPS coordinates of the first impact.

5. Weight and CG from the loading sheet.

Assumptions to the simulation were made based on the wreckage information and normal operational takeoff and single engine out emergency procedures:

1. Airspeed was assumed to be V2

2. Flaps setting was at 14 as revealed by the wreckage.

3. landing gears were retracted as revealed by the wreckage.

4. The Aircraft is trimmed for all engine climb

5. Eyewitnesses and surveillance video camera revealed that the Aircraft descended in a high bank angle before the impact. 13

6. Winds were utilised to reflect the tower winds at the time of the event.

7. the sideslip angle didn’t exceed 10 degrees.

For both cases, the simulation was started at the first radar hit. The engine No. 4 failure is initiated at 0.1 s and any pilot response is delayed for one second. The simulation showed that the Aircraft could continue down the runway heading. The Aircraft can produce the track of the incident by using wheel to roll right or insufficient rudder or wheel to control the thrust asymmetry. Since there are multiple means of recreating the track, given insufficient data, no attempt has been made to reproduce the actual Aircraft track.

The simulation showed a better climb gradient than the radar data. The results for the ‘typical’ engine failure show the pilot has the capability to utilize sufficient rudder to balance the outboard engine without sustained use of the wheel. Heading angle may be kept on the initial path and bank angle is maintained close to zero.

The simulation of reverse thrust at high power predicts that the airplane can be controlled for the worst case condition but only with timely input of large wheel and rudder. The simulation roughly maintained the runway heading. Full rudder was utilized along with full wheel to recover the initial upset after waiting one second. After the initial upset recovery, it took full rudder, approximately 50 degrees of wheel, and 9 degrees of opposite bank angle to maintain airplane heading in the worst case condition. Any less input would have the Aircraft heading turn right which is the direction the accident Aircraft took. It should also be noted that the simulation is predicting for this condition that the Aircraft sideslip angle approached 10 degrees. At sideslip angles greater than 10 degrees the linearity assumptions of the aerodynamic model begin to lose their fidelity.

No unusual control inputs were required to maintain the pitch control. For the reverse thrust case, the ground speed decayed below the radar data. If the airplane airspeed were started at a higher value, the resulting ground speed would match the event better. A higher airspeed would improve the airplanes ability to control the lateral-directional upset. Both simulations maintained angle-of-attack below the predicted stick shaker threshold.

13

Security video camera fixed at Sharjah Intl Airport which had captured the Aircraft during its climb until its explosion after impact.


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In a case of thrust reverser deployment for two second pilot delay in providing corrective wheel and rudder, the simulation exhibited that the Aircraft heading response away from the runway heading is considerably greater than indicated from the radar data.

In a case of forward thrust engine-out for one, two, and three second pilot delays. The one and two second delay conditions were still held fairly close to the runway heading. The three second delay had a greater deviation from runway heading, but the airplane was capable of returning heading to the runway heading.

ONGOING INVESTIGATION ACTIVITIES

1. “Human Factors” analysis is ongoing to determine the pilot’s reaction to the flght circumstances.

2. A more detailed simulation, taking into account more likely available flight parameters, is ongoing in Boeing facilities. A flight profile for the accident flight might be developed.

3. For more solid confirmation of the engine No. 4 core T/R status prior to the impact, certain T/R parts were sent to Boeing for autopsy lab examinations.

4. Engine No. 4 cowls and their support structure are under examination in Boeing labs.

SAFETY CONCERNS AND ACTIONS

As a result of the lab examination of the FDR and CVR, and taking under consideration the International Standards of Aircraft Operations for enhancing Flight Safety, the GCAA issued prompt safety recommendations to the Sudanese Civil Aviation Authority, Sudan Airways and Azza Air Transport:

1. Prompt safety recommendation no. (SR26/09) was issued to the Sudanese Civil Aviation Authority

to “Ensure that all flight recorder installations and operation comply with the appropriate International Standards”.

2. Prompt safety recommendation (SR27/09) was issued to Sudan Airways and Azza Air Transport to

review the maintenance procedures for the FDR and CVR installed on their aircraft, to ensure that their installation and operation meet the current International Regulations”.