AVIATION INVESTIGATION REPORT 

A05F0047 

LOSS OF RUDDER IN FLIGHT 

AIR TRANSAT 

AIRBUS A310-308 C-GPAT 

MIAMI, FLORIDA, 90 nm S 

06 MARCH 2005 

 

The Transportation Safety Board of Canada (TSB) investigated this occurrence for the purpose 

of advancing transportation safety. It is not the function of the Board to assign fault or 

determine civil or criminal liability. 

Aviation Investigation Report 

Loss of Rudder in Flight 

Air Transat 

Airbus A310-308 C-GPAT 

Miami, Florida, 90 nm S 

06 March 2005 

Report Number A05F0047

Synopsis 

At 0645 Coordinated Universal Time (UTC), Air Transat Flight 961 (TSC961), an Airbus 

A310-308 aircraft, departed Varadero, Cuba, for QuÈbec, Quebec, with 9 crew members and 

262 passengers on board. At approximately 0702 UTC, the aircraft was 90 nautical miles south 

of Miami, Florida, United States, and in level flight at flight level (FL) 350, when the flight crew 

heard a loud bang and felt some vibration. The aircraft entered a Dutch roll and the captain 

disconnected the autopilot to manually fly the aircraft. The aircraft climbed nearly 1000 feet 

while the captain tried to control the Dutch roll. The crew initiated a descent back to FL 350 and 

requested further descent and a possible diversion to Fort Lauderdale, Florida. During the 

descent, the Dutch roll intensity lessened and then stopped when the aircraft descended 

through FL 280. No emergency was declared. When the aircraft was abeam Miami, the crew 

decided to return to Varadero. 

During the landing flare, the rudder control inputs were not effective in correcting for a slight 

crab. The aircraft landed and taxied to the gate. After shutdown, it was discovered that the 

aircraft rudder was missing. Small pieces of the rudder were still attached to the vertical 

stabilizer. One flight attendant suffered a minor back injury during the event. 

Ce rapport est Ègalement disponible en franÁais. 
© Minister of Public Works and Government Services Canada 2007 

Cat. No. TU3-5/05-2E 

ISBN 978-0-662-46783-0 
TABLE OF CONTENTS

TRANSPORTATION SAFETY BOARD iii 

1.0 Factual Information ........................................................................ 1

1.1 History of the Flight ............................................................................................. 1

1.2 Injuries to Persons ................................................................................................ 2

1.3 Damage to the Aircraft ........................................................................................ 2

1.4 Personnel Information ......................................................................................... 3

1.4.1 Captain Information............................................................................................. 3

1.4.2 First Officer Information...................................................................................... 4

1.4.3 Flight Attendants.................................................................................................. 4

1.5 Aircraft Information............................................................................................. 5

1.5.1 General Information............................................................................................. 5

1.5.2 Aircraft History..................................................................................................... 5

1.5.3 Vertical Tail Plane Design ................................................................................... 5

1.5.4 Rudder Information ............................................................................................. 6

1.5.4.1 General................................................................................................................... 6

1.5.4.2 Rudder Design ...................................................................................................... 7

1.5.5 Rudder Manufacturing Method......................................................................... 9

1.5.6 Rudder Manufacturing Records......................................................................... 9

1.5.7 Rudder Modification Status.............................................................................. 10

1.5.8 Rudder Control System ..................................................................................... 10

1.5.8.1 Rudder Control System Components.............................................................. 10

1.5.8.2 Rudder Control System Operation .................................................................. 11

1.5.8.3 Dutch Roll Description ...................................................................................... 12

1.5.8.4 Dutch Roll Recovery Training .......................................................................... 13

1.5.8.5 On-Board Documentation ................................................................................. 13

1.5.9 Certification Information................................................................................... 13

1.5.9.1 Type Certificate................................................................................................... 13

1.5.9.2 Rudder Certification Tests ................................................................................ 13

1.5.10 Inspection Schedule ........................................................................................... 14

1.5.10.1 Scheduled Inspection Cycle .............................................................................. 14

1.5.10.2 Scheduled Rudder Inspections......................................................................... 15

1.5.10.3 Recently Completed Inspections...................................................................... 15

1.5.10.4 Rudder Damage Structural Repair Manual Limits........................................ 15

1.5.11 Maintenance Actions.......................................................................................... 16

1.5.11.1 General................................................................................................................. 16

1.5.11.2 Rudder Synchronization Check........................................................................ 16

1.5.11.3 Lightning Protection Plate Replacement......................................................... 16

1.5.11.4 Lightning Strike Repair ..................................................................................... 16

1.5.11.5 Miscellaneous Rudder Servo-Controls Maintenance .................................... 17

1.5.11.6 Maintenance Facilities........................................................................................ 17

1.6 Meteorological Information .............................................................................. 17

1.7 Aids to Navigation ............................................................................................. 17
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iv TRANSPORTATION SAFETY BOARD

1.8 Communications................................................................................................. 17

1.8.1 Air Traffic Control.............................................................................................. 17

1.8.2 Crew/Company Communications .................................................................. 17

1.8.3 Communication Between the Flight Deck and the Flight Attendants........ 18

1.9 Aerodrome Information .................................................................................... 19

1.10 Flight Recorders.................................................................................................. 19

1.10.1 Digital Flight Data Recorder ............................................................................. 19

1.10.2 Cockpit Voice Recorder ..................................................................................... 20

1.10.3 Direct Access Recorder ...................................................................................... 20

1.10.4 Data Sampling Rates .......................................................................................... 21

1.10.5 Filtering of Recorded Data ................................................................................ 21

1.10.6 Summary of Flight Recorder Data ................................................................... 21

1.11 Wreckage and Impact Information .................................................................. 22

1.11.1 Miscellaneous Damage ...................................................................................... 22

1.11.2 Vertical Tail Plane Damage............................................................................... 23

1.11.2.1 General................................................................................................................. 23

1.11.2.2 Main Attachment Fittings ................................................................................. 23

1.11.2.3 Hinge Arms ......................................................................................................... 24

1.11.3 Rudder Damage.................................................................................................. 26

1.11.3.1 General................................................................................................................. 26

1.11.3.2 Detailed Description of Rudder Damage........................................................ 26

1.11.4 Chemical Attack and Contamination .............................................................. 29

1.11.5 System Inspection and Testing......................................................................... 30

1.11.6 High-Intensity Radiated Fields Investigation ................................................ 31

1.11.7 Examination of Pre-Occurrence Photos........................................................... 31

1.12 Tests and Research ............................................................................................. 32

1.12.1 General................................................................................................................. 32

1.12.2 Examination of Two Sister Rudders ................................................................ 32

1.12.3 Elasticity Laminate Checker Test ..................................................................... 32

1.12.3.1 Elasticity Laminate Checker Test Description................................................ 32

1.12.3.2 ELCH 1óComplete Grid Examination of 24 Rudders of Similar Design.. 32

1.12.3.3 ELCH 2óFocused Examination of Rudders .................................................. 33

1.12.4 Examination of Other Rudders......................................................................... 33

1.12.5 Fluid Contamination Program ......................................................................... 34

1.12.6 Impact Damage Tests......................................................................................... 35

1.12.7 Laboratory Tests of Rudder Residuals ............................................................ 35

1.12.7.1 General................................................................................................................. 35

1.12.7.2 Bond Between Honeycomb and Face Sheets.................................................. 35

1.12.7.3 Bond at z-Section ................................................................................................ 36

1.12.7.4 Inter-laminar Bond Between Face Sheet Layers............................................. 36

1.12.7.5 Splice Bond .......................................................................................................... 36

1.12.7.6 Type of Resins..................................................................................................... 36
TABLE OF CONTENTS

TRANSPORTATION SAFETY BOARD v 

1.12.7.7 Cure of Resins ..................................................................................................... 37

1.12.7.8 Lightning Protection Plate Replacement......................................................... 37

1.12.7.9 z-Strut Paint Chips ............................................................................................. 37

1.12.7.10 Explosion Damage.............................................................................................. 37

1.12.8 Double Cantilever Beam Tests.......................................................................... 37

1.12.9 Disbond Growth ................................................................................................. 37

1.12.9.1 History of Earlier Design Aramid FibreñReinforced Plastic Rudders ........ 37

1.12.9.2 Temperature Effects ........................................................................................... 38

1.12.9.3 Vacuum Cycling Tests ....................................................................................... 38

1.12.10 Computer Simulation of Disbond Growth ..................................................... 39

1.12.11 Effects of Temperature and Moisture.............................................................. 40

1.12.12 Age-Related Structures and Materials Degradation ..................................... 40

1.12.13 Flight Dynamics Analysis ................................................................................. 41

1.12.13.1 Background ......................................................................................................... 41

1.12.13.2 Initial Simulations .............................................................................................. 41

1.12.13.3 Simulations to Determine the Lateral Force at Tail During the Event........ 42

1.12.13.4 Simulations to Determine the Longitudinal/Vertical Force at Tail 

During the Event ................................................................................................ 42

1.12.13.5 Additional Simulations...................................................................................... 42

1.12.13.6 Yaw Damper Modelling .................................................................................... 43

1.12.14 Effect of Disbond Bubble on Static Aerodynamic Loads.............................. 43

1.12.15 Flutter Analysis................................................................................................... 44

1.12.15.1 Analysis of a Dynamic Event............................................................................ 44

1.12.15.2 Method of Analysis ............................................................................................ 44

1.12.15.3 Baseline Analysis................................................................................................ 44

1.12.15.4 Failure Scenarios................................................................................................. 45

1.12.15.5 Summary of Flutter Analyses ........................................................................... 47

1.12.16 Time-Domain Flutter Analysis ......................................................................... 48

1.12.16.1 General................................................................................................................. 48

1.12.16.2 Method of Analysis ............................................................................................ 48

1.12.16.3 Results .................................................................................................................. 48

1.12.17 Summary of Flutter and Time-Domain Flutter Analyses............................. 49

1.12.18 National Transportation Safety Board Radar Data Analysis Study............ 49

1.12.19 Other Aircraft in Vicinity .................................................................................. 49

1.12.20 Space Objects....................................................................................................... 49

1.13 Organizational and Management Information .............................................. 50

1.13.1 Company Operations......................................................................................... 50

1.13.2 Maintenance Organization................................................................................ 50

1.14 Additional Information ..................................................................................... 51

1.14.1 All Operators TelexóFleet Inspection of Rudder ......................................... 51

1.14.1.1 AOT-1óFleet Inspection of Rudder Exteriors ............................................... 51

1.14.1.2 AOT-2óFleet Inspection of Rudder Interiors ................................................ 52
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vi TRANSPORTATION SAFETY BOARD

2.0 Analysis .......................................................................................... 53

2.1 General Information........................................................................................... 53

2.2 Flight Control System ........................................................................................ 53

2.2.1 General................................................................................................................. 53

2.2.2 Ability to Diagnose the Source of Flight Control Difficulty......................... 53

2.2.3 Dutch Roll Recovery .......................................................................................... 53

2.2.4 Decision to Return to Varadero........................................................................ 53

2.2.5 Decision not to Declare an Emergency............................................................ 54

2.2.6 Crew Communication and Decision Making................................................. 54

2.3 Maintenance ........................................................................................................ 55

2.3.1 Maintenance Program........................................................................................ 55

2.3.2 Maintenance Records......................................................................................... 55

2.3.3 Hinge Bearing Condition .................................................................................. 55

2.3.4 Adequacy of Rudder Inspection Program ...................................................... 55

2.4 Recorders ............................................................................................................. 56

2.4.1 Cockpit Voice Recorder Duration .................................................................... 56

2.4.2 Digital Flight Data Recorder Data Sampling.................................................. 56

2.4.3 Digital Flight Data Recorder Filtered Data ..................................................... 56

2.4.4 Preservation of Recorder Information............................................................. 57

2.5 Analysis of Rudder Failure Mechanism.......................................................... 57

2.5.1 General................................................................................................................. 57

2.5.2 Static Loading Phenomenon ............................................................................. 57

2.5.2.1 Large Rudder Deflection ................................................................................... 57

2.5.2.2 High Static Load ................................................................................................. 57

2.5.3 Dynamic Loading Phenomenon....................................................................... 57

2.5.3.1 High-Frequency Control Input......................................................................... 57

2.5.3.2 Flutter................................................................................................................... 58

2.5.4 Possible Causes of Flutter.................................................................................. 58

2.5.4.1 Flutter without Prior Structural Deviation ..................................................... 58

2.5.4.2 Flutter Following Structural Deviation ........................................................... 58

2.5.5 Growth of Rudder Damage............................................................................... 58

2.5.6 Possible Causes of Rudder Damage ................................................................ 58

2.5.6.1 Manufacturing Process ...................................................................................... 58

2.5.6.2 Material Degradation......................................................................................... 59

2.5.6.3 Mechanical Damage........................................................................................... 59

2.5.7 Failure Scenarios................................................................................................. 60

2.5.7.1 Summary of Important Points .......................................................................... 60

2.5.7.2 Most Likely Failure Scenario ............................................................................ 61
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TRANSPORTATION SAFETY BOARD vii 

3.0 Conclusions.................................................................................... 63

3.1 Findings as to Causes and Contributing Factors ........................................... 63

3.2 Findings as to Risk ............................................................................................. 63

3.3 Other Findings .................................................................................................... 64

4.0 Safety Action.................................................................................. 65

4.1 Action Taken ....................................................................................................... 65

4.1.1 Transportation Safety Board of Canada.......................................................... 65

4.1.1.1 TSB RecommendationsóAirbus Composite Rudder Inspection 

Program ............................................................................................................... 65

4.1.1.2 TSB Safety AdvisoryóCockpit Voice Recorder Duration............................ 66

4.1.1.3 TSB Safety AdvisoryóDigital Flight Data Recorder Recording of 

Filtered Data........................................................................................................ 67

4.1.1.4 TSB Safety AdvisoryóDigital Flight Data Recorder Low Recording 

Rates ..................................................................................................................... 67

4.1.1.5 TSB Safety AdvisoryóDutch Roll Recovery Procedure............................... 67

4.1.2 National Transportation Safety Board............................................................. 68

4.1.3 Bureau díEnquÍtes et díAnalyses pour la SÈcuritÈ de líAviation Civile .... 68

4.1.4 Airbus................................................................................................................... 68

4.1.4.1 All Operators Telex (AOT-1)............................................................................. 68

4.1.4.2 All Operators Telex (AOT-2)............................................................................. 69

4.1.5 Air Transat........................................................................................................... 69

4.1.5.1 Abnormal Situation............................................................................................ 69

4.1.5.2 Preservation of Recorder Information............................................................. 69

Appendices 

Appendix A ñ Direct Access Recorder/Digital Flight Data Recorder Data 

Comparison .................................................................................................71

Appendix B ñ Glossary.........................................................................................................73

Photos 

Photo 1. Right-Side View of Vertical Tail Plane and Rudder Residuals ......................... 3

Photo 2. Removal of the VTP and a View of the Position of the Main Attachment 

Fittings..................................................................................................................... 23

Photo 3. LeftñSide View of VTP and Rudder Residuals................................................... 26

Photo 4. Pre-Occurrence Photograph of Aircraft............................................................... 31

Photo 5. MSN 361 Honeycomb Showing Stain Caused by Hydraulic Fluid Ingress .. 33

Photo 6. Fluid Accumulation in the Nose of the Rudder of Aircraft MSN 701............. 34

Photo 7. Damage Growth After one Vacuum Cycle (Rudder of Aircraft MSN 361).... 39
TABLE OF CONTENTS

viii TRANSPORTATION SAFETY BOARD

Figures 

Figure 1. Schematic of the Vertical Tail Plane...................................................................... 5

Figure 2. Schematic of Hinge Arm Details ........................................................................... 6

Figure 3. Schematic of the Rudder......................................................................................... 7

Figure 4. Rudder Side Panel Construction ........................................................................... 8

Figure 5. Schematic of Rudder Hinge Fitting Details ......................................................... 9

Figure 6. Low Bonding Pressure Next to the z-Section .................................................... 36

Figure 7. Disbond Scenarios Studied in Flutter Analysis................................................. 45

Figure 8. Typical Load Response from Time-Domain Flutter Analysis......................... 48

Figure 9. Schematic of AOT-1 Areas of Inspection............................................................ 51

Figure 10. Schematic of AOT-2 Areas of Inspection.......................................................... 52
FACTUAL INFORMATION

TRANSPORTATION SAFETY BOARD 1 

1.0 Factual Information 

1.1 History of the Flight 

The pre-flight inspection was carried out by the captain before departure from Varadero; no 

damage was observed on the rudder. The inspection was conducted at night, the logo light was 

on, and the pilot was using a flashlight. However, it was difficult to see the entire rudder, 

especially the bottom part, which is partially concealed by the elevators. To see the bottom part, 

the pilot has to step back from the aircraft, thus reducing the acuity of the observation. 

The crew engaged autopilot system No. 2 on departure from Varadero. The flight progressed 

normally until the aircraft reached flight level (FL)1 350, its assigned altitude. At approximately 

0702 Coordinated Universal Time (UTC),2 the flight crew heard a loud bang immediately 

followed by several seconds of vibration. Cabin crew members located in the back of the aircraft 

were thrown to the floor and unsecured galley carts moved freely. The aircraft started to Dutch 

roll,3 and the captain took control and disconnected the autopilot. The aircraft was difficult to 

control in the lateral axis. In an attempt to better manage the cockpit workload, the other 

autopilot system (No. 1) was engaged. As the Dutch roll movement started to intensify, 

autopilot No. 1 was disengaged and the aircraft was hand-flown. 

During these actions, the aircraft climbed to about FL 359. The flight crew requested a descent 

and informed air traffic control (ATC) that they had experienced an autopilot problem and had 

reverted to flying manually. While descending, the crew cycled through the electronic 

centralized aircraft monitor (ECAM) system pages in an attempt to diagnose the problem. 

Throughout the event, there was no ECAM message relating to the control problem that the 

aircraft had experienced, and there were no warning lights or cockpit indications of an aircraft 

malfunction. Even with limited clues as to the cause of the Dutch roll, the crew knew that 

descending to a lower altitude might lessen or stop the Dutch roll motion. Initial indications led 

to the possibility of the loss of both yaw dampers (YD) but both YD switches were engaged. 

Had a dual YD failure occurred, the flight warning computer would have triggered appropriate 

warnings and messages, and the autopilot would have disconnected. 

 

1 See Glossary at Appendix B for all abbreviations and acronyms. 

2 All times are Coordinated Universal Time (UTC) unless otherwise noted. 

3 Dutch roll is a motion of an aircraft that consists of simultaneous oscillations of the bank 

(or roll) angle, the sideslip angle, and the heading angle. The roll manifests itself as an 

out-of-phase combination of ìtail waggingî and rocking from side to side. The motion is 

normally well damped in most light aircraft, though some aircraft with well-damped 

Dutch roll modes can experience a degradation of damping as airspeed and altitude 

increase. Dutch roll stability can be artificially increased by the installation of a yaw 

damper, as is the case with most swept-wing aircraft (see also Section 1.5.8.3). 
FACTUAL INFORMATION

2 TRANSPORTATION SAFETY BOARD

The Dutch roll gradually decreased in the descent and ceased when the aircraft passed FL 280. 

The crew continued the descent to 10 000 feet above sea level (asl) in preparation for a landing 

in Fort Lauderdale. The captain returned control of the aircraft to the first officer and called the 

flight director (FD) to provide the standard briefing to the cabin crew for emergency or 

abnormal situations. 

The crew contacted company dispatch to discuss the situation and elected to return to 

Varadero, where the company was better equipped to deal with the aircraft and the passengers. 

At 0739, the flight was cleared to Varadero at FL 190. 

During the climb to FL 190, the crew engaged autopilot No. 1 and disengaged it during the final 

portion of the visual approach to Runway 06 at Varadero. During the landing flare, nose wheel 

steering was used for directional control on the runway. An uneventful landing was completed 

at 0819. 

The crew conducted a flight control check after landing and the ECAM indicated that 

everything was normal. The aircraft was taxied to the gate where the passengers were deplaned 

normally through the main door. After shutdown, a visual inspection revealed that the aircraft 

rudder had broken and most of it was missing. 

1.2 Injuries to Persons 

Crew Passengers Others Total 

Fatal ñ ñ ñ ñ 

Serious ñ ñ ñ ñ 

Minor/None 9 262 ñ 271 

Total 9 262 ñ 271 

1.3 Damage to the Aircraft 

The rudder was substantially damaged (see Photo 1), and the rear attachment fittings of the 

vertical tail plane (VTP) were delaminated locally. There was minor damage to the tail cone. 
FACTUAL INFORMATION

TRANSPORTATION SAFETY BOARD 3 

Photo 1. Right-side view of vertical tail plane 

and rudder residuals 

1.4 Personnel Information 

Captain First Officer 

Licence Airline Transport Commercial 

Medical Expiry Date 01 September 2005 01 July 2005 

Total Flying Hours 10 795 11 305 

Hours on Type 450 500 

Hours Last 90 days 75 200 

Hours on Type Last 90 Days 75 200 

Hours on Duty Prior to Landing 4.6 4.6 

Hours off Duty Prior to Work Period 60 60 

1.4.1 Captain Information 

The captain held a Canadian airline transport pilot licence (ATPL) ñ aeroplane, endorsed for 

single- and multi-engine land aeroplanes, with type ratings on Boeing 727, Boeing 737, 

Boeing 757, Convair 580, Airbus A310, Fokker 100, and Lockheed 1011 aircraft. His licence was 

endorsed with a Group 1 instrument rating valid until 01 September 2005. 
FACTUAL INFORMATION

4 TRANSPORTATION SAFETY BOARD

The captain started working for the company as a captain on the Boeing 757 on 18 March 1996. 

In 1997, he qualified as captain on the Boeing 737-400 and flew it for about six months before 

returning as captain on the Boeing 757. In 2003, he began his conversion to the A310, and under 

the supervision of an Air Transat instructor, completed the A310 computer-based ground 

school. The flight simulator portion of the initial A310 training was conducted by Air Transat 

instructors at a training centre in Miami, Florida, from 12 to 27 August 2003. All training was in 

accordance with the company A310 training program. 

The captain passed his initial pilot proficiency check (PPC) as an A310 captain on 27 August 

2003, and his last line check was performed on 17 September 2004. His last PPC was performed 

on 14 December 2004. Company training records indicate that he had successfully completed all 

required recurrent training. 

1.4.2 First Officer Information 

The first officer held a commercial pilot licence ñ aeroplane, endorsed for single- and 

multi-engine land aeroplanes, with type ratings on Convair 580, Airbus A310, and 

Lockheed 1011 aircraft. His licence was endorsed with a Group 1 instrument rating valid until 

01 December 2005. 

The first officer started working for the company on 15 February 1988 as a flight engineer on the 

Lockheed 1011 aircraft, accumulating 8500 hours of flight time. He was qualified as first officer 

on the Lockheed 1011 on 18 June 2002. In 2004, he began conversion to the A310. Under the 

supervision of an Air Transat instructor, he completed the A310 computer-based training. The 

flight simulator portion of the initial A310 training was conducted by Air Transat instructors at 

a training centre in MontrÈal, Quebec, from 25 May to 15 June 2004. All training was in 

accordance with the company A310 training program. 

The first officer passed his initial PPC as an A310 first officer on 15 June 2004, and his last line 

check was performed on 07 July 2004. His last PPC was performed on 08 October 2004. 

Company training records indicate that he had successfully completed all required recurrent 

training. 

1.4.3 Flight Attendants 

The cabin crew comprised seven flight attendants (FAs), including a flight director (FD) and an 

assistant flight director (AFD), all of whom had 10 to 16 years of service. They were qualified 

and trained in accordance with the requirements of Transport Canada and Air Transat. 
FACTUAL INFORMATION

TRANSPORTATION SAFETY BOARD 5 

1.5 Aircraft Information 

1.5.1 General Information 

The occurrence aircraft was an Airbus A310-308, manufacturerís serial number (MSN) 597. 

Transport Canada issued the certificate of registration on 16 May 2001 and the certificate of 

airworthiness on 16 June 2001, both valid at the time of the occurrence. 

1.5.2 Aircraft History 

The occurrence aircraft had its first flight in September 1991 and was delivered to a Middle 

Eastern airline in August 1992, where it remained until acquired by Air Transat in May 2001. At 

the time of the occurrence, the aircraft had accumulated 49 224 flight hours and 13 444 flight 

cycles. By comparison, the flight hour and flight cycle fleet leader aircraft for this aircraft type 

had accumulated 75 675 hours and 34 384 cycles respectively. 

1.5.3 Vertical Tail Plane Design 

The VTP consists of a spar box, 

leading edge fairing, trailing edge 

panels, and tip (see Figure 1). The 

spar box consists of left and right 

side panels each composed of solid 

carbon fibreñreinforced plastic 

(CFRP) laminate skin and interior 

stiffeners. At the bottom of each 

side panel, there are three large 

integrally constructed CFRP lugs, 

known as the main attachment 

fittings, which attach to the 

fuselage. At the front and rear of 

the box, there are solid CFRP 

laminate spars running the length 

of the VTP, joining the left and right 

skin panels, forming the front and 

rear faces of the spar box. 

In the centre of the box, there is a 

shorter solid CFRP laminate spar, 

which extends only up to rib 5. At 

the bottom of each of these three 

spars are two integrally constructed 

lugs, known as transverse load 

fittings, which attach to the fuselage. Within the box, there are a total of 18 solid CFRP laminate 

ribs, including closing ribs at the bottom and top. The leading edge and the tip are constructed 

of sandwich composite. Attached to each side of the rear spar, and extending aft, there is a flat 

trailing edge panel that acts as an aerodynamic fairing to fill the gap between the rear spar of 

the VTP and the leading edge of the rudder. 

Figure 1. Schematic of the vertical tail plane 
FACTUAL INFORMATION

6 TRANSPORTATION SAFETY BOARD

There are seven hinge positions along the VTP rear spar for the attachment of the rudder. These 

are numbered 1 through 7, from bottom to top. Figure 2 shows the design details at these hinge 

points. At each hinge position, there is a CFRP fitting attached to the rear spar. Each CFRP 

fitting has two lugs, one on the left and one on the right. The two front arms of each V-shaped 

metal hinge arm fit into these lugs on the rear of the VTP spar. 

Figure 2. Schematic of hinge arm details 

The hinge arms are attached to the CFRP fittings with spherical bearings, so they are free to 

pivot up and down. The rear of each hinge arm contains a hinge point for the attachment of the 

rudder. The hinge arm at hinge position 4 is supported in the vertical direction by a metal 

structural tube referred to as the z-strut. All the vertical loads from the rudder are transferred to 

the VTP through the z-strut. Rudder movement is controlled by three hydraulic actuators 

located inside the VTP at hinge positions 2, 3, and 4. The forward ends of the actuators are 

attached to CFRP fittings on the rear spar of the VTP, and the aft ends are attached to aluminum 

alloy fittings on the front spar of the rudder. 

1.5.4 Rudder Information 

1.5.4.1 General 

The occurrence rudder, serial number 1331, was of the part number series A55471500, which is 

in use on earlier production A310, A300-600, A330, and A340 aircraft. It was the same rudder 

that had been originally installed on the occurrence aircraft at the time of manufacture in 1991. 

This rudder was one of the first in a batch of five rudders whose side panels were manufactured 

by the company Soko in Mostar, former Yugoslavia. The side panels were shipped from Soko to 

Airbus in Stade, Germany, where they were assembled into rudders. 
FACTUAL INFORMATION

TRANSPORTATION SAFETY BOARD 7 

1.5.4.2 Rudder Design 

The rudder consists of a single spar at the front, two side panels that fasten together at the 

trailing edge, and top and bottom closing ribs (see Figure 3). The side panels are of single-piece 

construction and do not include any design features to mechanically arrest the growth of 

disbond damage. Each side panel is a sandwich composite constructed of a non-metallic 

NomexÆ aramid-based honeycomb core, with CFRP face sheets, and a glass fibreñreinforced 

plastic (GFRP) intermediate layer between the CFRP and the honeycomb as shown in Figure 4. 

The GFRP intermediate layer does not have a structural purpose. It is simply a carrier for the 

resin that bonds the CFRP to the honeycomb. There is a layer of TedlarÆ on the interior face to 

provide a moisture barrier, and a layer of film adhesive (AF 126) on the exterior face to provide 

aerodynamic smoothness. The density and thickness of the honeycomb and the number of face 

sheets vary with location because they are designed to react to applicable loads. 

Figure 3. Schematic of the rudder 

Different pieces of honeycomb are bonded together along their side edges by a splice bonding 

adhesive. This same adhesive is also used to bond the side edges of the honeycomb to the 

z-section. The forward and bottom edges of the side panels are made with a pre-cured CFRP 

z-section. The side panels are fastened to the spar and ribs using blind mechanical fasteners. 

There are three aluminum lightning protection plates (LPPs) running chordwise on each side 

panel. To avoid galvanic reaction between these metal plates and the CFRP, there is an 

intermediate insulating layer of GFRP. There is a single spar, located along the front edge of the 
FACTUAL INFORMATION

8 TRANSPORTATION SAFETY BOARD

rudder and running the entire length of the rudder. The spar is a sandwich composite 

constructed of a NomexÆ honeycomb core with CFRP face sheets. There are seven lightening 

holes distributed along the length of the spar. 

There are only two ribs within the rudder. 

Rib 0 is the closing rib at the bottom of the 

rudder and is a sandwich composite 

constructed of a NomexÆ honeycomb core 

with CFRP face sheets. Rib 54a, constructed 

of aluminum, is the closing rib at the top of 

the rudder. The leading edge fairing of the 

rudder is divided into multiple sections 

along its length, each constructed of 

sandwich composite (see Figure 4). The 

leading edge fairing sections are fastened to 

the side panel z-sections with threaded 

fasteners. There is an aluminum alloy strip 

along this row of fasteners as part of the 

lightning-protection system. Attached to the 

z-section at the bottom of each side panel is a rubber weatherstrip that covers the gap between 

the bottom of the rudder and the top of the tail cone. The weatherstrip is attached with threaded 

fasteners, and a metal strip is used as a washer plate along this row of fasteners. The side panels 

attach together at the rear of the rudder by a row of mechanical fasteners running parallel to the 

trailing edge, roughly 30 cm ahead of the trailing edge. A metal protective strip runs down the 

entire length of the rudder trailing edge, which is also attached using mechanical fasteners. 

There are three hoisting points on each side panel. 

There are seven hinge positions, numbered 1 through 7, from bottom to top. Figure 5 shows the 

design details at these hinge points. At each hinge position, aluminum alloy fittings are attached 

to solid GFRP blocks integrated locally into the side panels and to the spar web by mechanical 

fasteners. The core of the spar web, where the fasteners pass through, is filled by core filler and 

reinforced by an aluminum backing plate. The three control actuators attach to the rudder at 

hinge positions 2, 3, and 4. The metal hinge fittings at these locations have two lugs, one to act 

as the hinge point, and one to attach to a hydraulic actuator. 

Figure 4. Rudder side panel construction
FACTUAL INFORMATION

TRANSPORTATION SAFETY BOARD 9 

Figure 5. Schematic of rudder hinge fitting details 

1.5.5 Rudder Manufacturing Method 

The rudder side panels, rudder spar, and rib 0 are manufactured and cured separately and then 

assembled with mechanical fasteners into a rudder. Each side panel is assembled in a mold, 

with the exterior face on the bottom against the face of the mold. During curing, the 

manufacturing process results in the lower (outer) skin having a stronger bond. Although both 

bonds exceed design requirements, the inner skin bond does so by a smaller margin. The three 

LPPs are integrally manufactured and co-cured with the side panel. 

1.5.6 Rudder Manufacturing Records 

Some manufacturing records for the side panels of the occurrence rudder were lost when the 

factory was bombed during the Yugoslavian war. Manufacturing records available at Airbus in 

Stade, Germany, and Toulouse, France, were reviewed for the occurrence rudder. This review 

found that non-conformities were detected by the quality assurance system, corrective actions 

were defined, rework was conducted, and the final product was inspected and released as 

airworthy. These non-conformities included such items as the position of hoisting points, the 

resistance of the anti-static paint, and various splice bond, skin and core filler re-works. The 

quality assurance of the Soko components was always under the responsibility of Airbus. The 

manufacturing records indicated that the rudder was in an airworthy condition at final 

assembly. 
FACTUAL INFORMATION

10 TRANSPORTATION SAFETY BOARD

1.5.7 Rudder Modification Status 

The following is the modification status of the occurrence rudder: 

• Modification 5844 (Glass Intermediate Layer). The occurrence rudder was a 

post-modification 5844 (Service Bulletin [SB] A310-55-2012) design, which 

incorporated a GFRP layer between the honeycomb and the CFRP skin, rather than 

aramid fibreñreinforced plastic (AFRP) as used in earlier design. 

• Modification 8408 (Change in Honeycomb Size). The occurrence rudder was a 

post-modification 8408 configuration, which incorporated increased density 

honeycomb at certain locations. 

• Modification 8827 (Change in Spar Construction). The occurrence rudder was 

pre-modification 8827, meaning its spar had the earlier design NomexÆ 

honeycomb/CFRP sandwich spar, rather than the solid CFRP spar of later design. 

• Modification 5185 (Single-Piece Side Panels). The occurrence rudder was postserial 1035, which means that the side panels were each constructed as a single panel. 

Earlier side panels were constructed of two parts, top and bottom, with a chordwise 

joint. 

• Modification SRM (structural repair manual) 55-41-12 (Reinforcing Bolts in GFRP 

Blocks). The occurrence rudder had received modification SRM 55-41-12, 

Paragraph 27, during manufacture. This modification added reinforcing bolts 

through the GFRP blocks at the hinge point level. 

1.5.8 Rudder Control System 

1.5.8.1 Rudder Control System Components 

The following is a descriptive list of the Airbus A310 rudder control system components: 

• The rudder pedals, the rudder trim actuator, the two YD actuators, and the autopilot 

yaw actuator (APYA), which command the rudder to move. 

• The push rods, the bell cranks, and the tension regulator and cables, also referred to 

as linkage, which transmit rudder commands. 

• The three servo-controlsóupper, middle, and lowerówhich operate the rudder. (The 

maximum rudder actuation rate with no load is 60 ± 5° per second. The maximum 

rudder deflection is 30° either left or right.) 

• The differential unit, a mechanical device, which sends a command to the rudder 

servo-controls. This unit sums the pilot or the autopilot input and the YD input. 
FACTUAL INFORMATION

TRANSPORTATION SAFETY BOARD 11 

• The two rudder travel limiter (RTL) systems, which provide a variable stop, limiting 

the travel of the rudder mechanical linkage downstream of the differential unit, and 

thus the input to the three servo-controls as the airspeed increases. 

• The transmitter, located on the fin at rib 1 and connected to the rudder with a rod 

attached to fitting No. 1, which indicates the rudder surface position to the 

appropriate ECAM display unit. 

1.5.8.2 Rudder Control System Operation 

The YD actuators are electro-hydraulic mechanisms that operate the YD system. The YD system 

has three functions: Dutch roll damper; turn coordinator; and yaw compensator during an 

engine failure on take-off or go-around. The YD commands are limited by software in the flight 

augmentation computers to a maximum of 39° of rudder movement per second. The maximum 

allowable displacement of the rudder by the YD is ±10° at indicated airspeeds up to 165 knots. 

The maximum allowable displacement at indicated airspeeds greater than 165 knots is 

determined by a formula (10 x (165/knots indicated airspeed [KIAS])2). 

As the aircraft was flying at an indicated airspeed of 270 knots at the time of the occurrence, the 

maximum displacement of the rudder by the YD was of ±3.7°. The YD and the rudder pedals 

are not linked, so YD inputs do not result in pedal motion. Rudder pedal and YD commands are 

restricted to the limits imposed by the RTL system. Rudder position is determined by the sum 

of the pilot or autopilot input and the YD commands limited by the travel limitation unit. 

The APYA, which produces yaw autopilot commands, is a single unit that houses two 

electro-hydraulic actuators, each controlled by a flight control computer (FCC). The APYA has 

an output lever that is connected through a torque limiter to the main bell crank. The torque 

limiter allows a pilot to override autopilot output by applying about 65 decanewtons (daN) 

more than the rudder pedal feel forces. Autopilot yaw control commands are limited by 

software in the FCC to a maximum of 34° of rudder per second. The APYA and the rudder 

pedals are rigidly linked; therefore, autopilot yaw input results in pedal motion. 

The RTL system reduces the maximum allowable rudder deflection as airspeed increases. The 

limitation is such that the maximum deflection that can be achieved by the rudder remains 

lower than the deflection that would induce limit loads on the structure throughout the flight 

envelope. 
FACTUAL INFORMATION

12 TRANSPORTATION SAFETY BOARD

1.5.8.3 Dutch Roll Description 

The Airplane Upset Recovery Training Aid4 describes the Dutch roll as follows: 

Static directional stability is a measure of the tendency of an airplane to 

weathervane into the free stream air mass. The vertical fin and distribution 

of flat plate area aft of the CG [centre of gravity] tend to reduce sideslip and 

add to good directional stability. All conventional airplanes require 

positive static directional stability. In simple terms, an airplane with good 

directional stability always wants to point directly into the relative windó

zero sideslip. As directional stability increases, the speed at which the 

aircraft returns to zero sideslip after being disturbed increases (higher 

frequency). In order to minimize overshoots in sideslip, the damping in the 

directional axis must be increased as the directional stability is increased. 

An undesirable characteristic can develop when the directional damping is 

not adequate enough to prevent overshoots in sideslip. A phenomenon 

known as ìDutch rollî (based on the similarity with the motions of highspeed ice skaters) can occur. A Dutch roll occurs when yaw rates produce 

sideslips, which produce roll rates. If the sideslips are not adequately 

damped, the aircraft nose will swing back and forth with respect to the 

relative wind, and the aircraft will roll right and left due to the dihedral 

effect (the wingsweep results in asymmetric lift, depending on the relative 

wind). Airplanes designed to fly at higher Mach numbers have more 

wingsweep to control the critical Mach number (the speed at which shock 

waves begin to form on the wing). As wingsweep increases, the dihedral 

effect increases, and if the airplane is not adequately damped in the 

directional axis, a Dutch roll might occur if the airplane is upset 

directionally. Yaw dampers were designed to minimize yaw rates, which 

result in sideslip rates, and are very effective in modern transports in 

damping the Dutch roll. However, some transport airplanes have a neutral 

or slightly divergent Dutch roll if the yaw damper is off or inoperative.5

Conventional airplanes exhibit more of a Dutch roll tendency at higher 

altitude (less damping) and higher speed (more directional stability). 

Therefore, if a pilot encounters a Dutch roll condition, every effort should 

be made to ìslow down and go down.î With a properly functioning yaw 

damper, Dutch rolls will not occur in modern transport aircraft. Transport 

airplanes are certificated to demonstrate positively damped Dutch roll 

oscillations. The rudder should not be used to complement the yaw 

damper system. If the yaw damper system is inoperative, the rudder 

should not be used to dampen Dutch roll. 

 

4 The Airplane Upset Recovery Training Aid was developed by aircraft manufacturers, 

airlines, pilot associations, flight training organizations, and government and regulatory 

agencies to help pilots recover from unintentionally exceeding parameters normally 

experienced during line operations or training. 

5 This is not the case for A310 Dutch roll for which Dutch roll characteristics remain 

convergent in the whole flight envelope. 
FACTUAL INFORMATION

TRANSPORTATION SAFETY BOARD 13 

1.5.8.4 Dutch Roll Recovery Training 

During Air Transat initial training, pilots are exposed to Dutch roll recovery. The exercise is 

conducted with YDs engaged to demonstrate the automatic damping, and with the YDs 

disengaged to practice the recovery technique and to demonstrate the natural damping. During 

the exercise, at the request of the pilot flying, the pilot not flying rapidly applies rudder until 40°

of bank is achieved and then releases the rudder pedal. The rudder should not be used during 

recovery and the rudder control should remain in the neutral position. Transferring fuel 

forward will improve Dutch roll characteristics, and flying at or below FL 310 will improve 

aircraft directional stability. 

1.5.8.5 On-Board Documentation 

The A310 quick reference handbook (QRH) does not include procedures for abnormal flight 

conditions related to Dutch roll. However, the expanded checklist in the Flight Crew Operating 

Manual provides information to control Dutch roll in case of a yaw damper fault. A yaw 

damper fault was not the problem in this event. 

1.5.9 Certification Information 

1.5.9.1 Type Certificate 

This model of aircraft is covered by Transport Canada type certificate A-151. The data sheet 

provides the following information applicable to this occurrence: 

• maximum operating speed: 340 KIAS 

• maximum operating Mach: 0.84 

• flight load factor with flaps up: -1.0 to +2.5 

1.5.9.2 Rudder Certification Tests 

The manufacturer conducted the following structural and flutter tests during the original 

certification of the rudder: 

Static Load TestóThe rudder was tested and sustained 1.6 times the limit 

load at high temperature/high humidity conditions. 

Fatigue TestóThe rudder was cycled for three lifetimes (3 x 48 000 flight 

cycles) between -35°C and 70°C and varying moisture content. 

Structural Tests 

Damage Tolerance TestsóFollowing the fatigue test, artificial damage was 

introduced to the specimen and it was subjected to one further lifetime of 

fatigue. No damage growth was observed. Following this fourth fatigue 

life of cycling, the specimen was subjected to 1.8 times the limit load 

without failure. The specimen was then loaded several times to 2.3 times 

the limit load without failure. The specimen was then subjected to 

increased damage sizes and finally failed at 2.15 times the limit load. The 

applied load during these tests was mechanical only and did not include 

vacuum cycling tests. Vacuum cycling tests were not specifically 

demanded for certification. 
FACTUAL INFORMATION

14 TRANSPORTATION SAFETY BOARD

 Sub-Component TestsóTests were conducted on the anchor nuts used to 

attach the leading edge to the z-section, the trailing edge connection, the 

connection of the side panel to the front spar, the load introduction area at 

the hydraulic actuators, and the front spar. 

Theoretical Investigation (Normal Cases)óA theoretical dynamic model 

was constructed. The vibration behaviour was calculated using the 

MSC Nastran finite element software program, and the model was 

adjusted to match ground vibration test (GVT) results. A more complex 

finite element model was subsequently developed for the multi-role 

tanker transport (MRTT) conversion. This model was accurate to within 3 

per cent of the GVT results. Analysis found the aircraft to have satisfactory 

flutter margins up to the certification limit of 1.2 times the operating dive 

speed at less than Mach 1.0, meeting certification requirements. 

Theoretical Investigation (Failure Cases)óFlutter analysis was conducted 

for a number of failure cases, including cracks in spar or skin, failure of 

two of three hydraulic circuits, ice accretion on leading edges, and water 

ingress into honeycomb core. Satisfactory flutter margins were found for 

all failure conditions. 

Ground Vibration TestsóGVTs were conducted on the A310-200, 

A310-300, and A310-300 MRTT versions to calibrate the computer models. 

Flutter Tests 

Flight Vibration TestsóFlight vibration tests were conducted on the 

A310-200, A310-300, and A310-300 MRTT versions. No flutter or critical 

damping reduction occurred up to the demonstrated flight diving speed 

of 410 knots calibrated airspeed (KCAS), Mach 0.9. 

1.5.10 Inspection Schedule 

1.5.10.1 Scheduled Inspection Cycle 

The scheduled aircraft inspection cycle is as follows: 

Transit check Before each flight 

Daily check 36 hours of lapse time 

Weekly check 8 calendar days 

A-check (1 through 12) 450 flight hours 

C-check (1 through 8) 15 months 

Note: Aircraft utilization is approximately 300 hours per month (3600 hours per year). 
FACTUAL INFORMATION

TRANSPORTATION SAFETY BOARD 15 

1.5.10.2 Scheduled Rudder Inspections 

The rudder is inspected during the following inspections: 

General visual inspection (GVI) from the ground of 

empennageóGVI (G) 

daily/transit check 

GVI at armís length of empennageóGVI (A) 2-C check 

GVI (A) of rudder 2-C check 

Rudder hinge free-play measurement 4-C check 

Detailed visual inspection (DVI) of rudder hinge arms 5 years 

DVI of rudder hinge fittings 5 years 

DVI of rudder front spar 5 years 

Non-destructive inspection (NDI) of rudder side panels 5 years 

1.5.10.3 Recently Completed Inspections 

The most recently completed major inspections before the occurrence were the following: 

May 2001 5-year rudder NDI ñ 34 415 hours and 10 037 cycles 

May 2004 2-C check (at TAP Portugal) ñ 46 198 hours and 

12 809 cycles 

01 March 2005 A-11 check ñ 49 156 hours and 13 429 cycles 

05 March 2005 daily/transit check (before departure from QuÈbec to 

Varadero) ñ 49 197 hours and 13 439 cycles 

1.5.10.4 Rudder Damage Structural Repair Manual Limits 

Chapter 55-41-00, Figure 105, of the structural repair manual (SRM) specifies that damage to the 

rudder side panels of the type ìimpact and delamination without visible cracks or holesî is to 

be repaired according to the following requirements: 

• below 1000 mm2: allowable damage 

• 1000 to 10 000 mm2: monitor damage and repair if it grows 

• 10 000 to 40 000 mm2: monitor damage and repair within 2500 hours in accordance 

with the SRM 

• above 40 000 mm2: repair immediately and refer to manufacturer6

 

6 These allowable values are in the process of being reviewed as a result of damage 

propagation studies that were conducted during the course of this investigation. 
FACTUAL INFORMATION

16 TRANSPORTATION SAFETY BOARD

1.5.11 Maintenance Actions 

1.5.11.1 General 

All inspection and maintenance work reports were analyzed from the date of the aircraftís first 

flight in September 1991 until the time of the occurrence. All records of structural repairs were 

examined, including all maintenance activities reported for components of the rudder control 

surface and system components, as well as special inspections. The investigation determined 

that the aircraft was maintained in airworthy condition in accordance with the 

Transport Canada (TC)ñapproved maintenance program. Significant rudder-related 

maintenance actions are described below. 

1.5.11.2 Rudder Synchronization Check 

There is a requirement every 1300 flight hours to conduct a rudder synchronization check as 

specified in Airbus SB A310-27-2082. This inspection requires the technicians to access the area 

at the base of the rudder. Although it does not include a structural inspection of the rudder, any 

significant external damage would be visible. This inspection had been carried out concurrently 

with the A-11 inspection on 01 March 2005, five days before the occurrence. No abnormalities 

were reported. 

1.5.11.3 Lightning Protection Plate Replacement 

On 20 May 2004, less than one year before the occurrence and during the aircraft 2-C inspection, 

the rudder lower right-side LPP was found to be corroded in the aft attachment area. It was 

subsequently replaced, and tap tests7 of the affected area following the replacement showed no 

indications of inadequate bonding. Because this was one of the few rudder maintenance 

activities that were recorded, the complete replacement process of the LPP was investigated. No 

anomalies were found that could have contributed to the occurrence. 

1.5.11.4 Lightning Strike Repair 

On 12 August 1997, during the aircraft 4-C inspection, a non-routine inspection card was raised 

to address suspected lightning strike damage. The defect was written as ìupper corner of 

rudder, lightning strike mark,î and the corrective action was written as ìrudder upper corner 

lightning strike area repaired in accordance with SRM 51-73-10.î This was a minor repair within 

SRM limits; the manufacturer was not advised. No photos or other records of the damage were 

available. This damage occurred more than seven years before the occurrence, and the aircraft 

was subject to all regular inspections in the intervening time. 

 

7 A tap test is a non-destructive inspection technique that involves gently striking the 

inspected component with a hand-held mass and evaluating the resulting sound to 

identify the presence of damage. 
FACTUAL INFORMATION

TRANSPORTATION SAFETY BOARD 17 

1.5.11.5 Miscellaneous Rudder Servo-Controls Maintenance 

In December 1999, the number 7 rudder hinge arm was found to have excessive play and it was 

repaired. In May 2004, the rudder servos were modified according to SB A310-27-2091. 

1.5.11.6 Maintenance Facilities 

Inspection of the operator maintenance base facility in MontrÈal showed no indication that the 

aircraft rudder suffered an impact against crew lifting devices, other devices on the ramp, or 

hangar door frame. The investigation also determined that the tail of the aircraft could not have 

been affected by the heating or lighting systems in place or at the previous location of the 

company in Mirabel, Quebec. 

1.6 Meteorological Information 

The reported weather at the time of departure from Varadero (0600) was as follows: winds 

variable at two knots, visibility 8000 m, few clouds (less than 2/8 sky coverage) at 1800 feet 

above ground level (agl), temperature 14°C, dew point 12°C, altimeter setting 1021 millibars. 

The weather at Varadero at the time of landing (0800) was reported as follows: winds variable 

at two knots, visibility 7000 m, few clouds at 1800 feet agl, temperature 12°C, dew point 11°C, 

altimeter setting 1020 millibars. 

At the time of the occurrence, the flight crew was in night visual flight conditions, and no 

turbulence was reported. 

1.7 Aids to Navigation 

There were no reported problems with navigational aids. 

1.8 Communications 

1.8.1 Air Traffic Control 

TSC961 levelled off at FL 350 at 0701. As a result of the in-flight problem, TSC961 climbed 

nearly 1000 feet, but there was no other traffic in the area; this altitude incursion did not result 

in a loss of separation. TSC961 was initially being guided for an approach in Fort Lauderdale, 

but the aircraft subsequently returned to Varadero. 

The crew was in contact with the controllers of four separate ATC sectors between the time of 

the occurrence and landing at Varadero. An emergency was not declared. 

1.8.2 Crew/Company Communications 

At 0717, a phone patch was initiated with Air Transat dispatch in MontrÈal through New York 

aeronautical radio incorporated (ARINC) using their high frequency radio. The flight problem 

was discussed with dispatch and maintenance. 
FACTUAL INFORMATION

18 TRANSPORTATION SAFETY BOARD

1.8.3 Communication Between the Flight Deck and the Flight Attendants 

After hearing the abnormal loud noise, the FD contacted the flight deck via the interphone. The 

flight crew was unable to respond at the time because of the control situation. Shortly thereafter, 

as per the companyís prescribed abnormal/emergency communication procedure, the captain 

called the FD and provided the TESTRA briefing: 

• T ó Type of problem: autopilot not responding, flight diverting to Fort Lauderdale 

• E ó Evacuation (land or ditch): no evacuation 

• S ó Signals (standard or alternate): standard signals 

• T ó Time available before landing: 10 minutes 

• R ó Relocation of passengers: not necessary 

• A ó Announcement to passengers done by (captain or FD): captain 

The captain did not ask the FD for a briefing with respect to the cabin environment and none 

was provided. In abnormal and emergency situations, it is neither the flight crew procedure nor 

practice to ask the FD if he/she has information to provide. It is assumed that any information 

that may assist in decision making will automatically be provided. Air Transatís procedure for 

communicating in abnormal situations calls for flight crew to ask if there are any questions 

following the TESTRA briefing, which they did. None of the FAs that were in the area of the aft 

galley contacted the FD or the flight crew to provide information in reference to the abnormal 

events encountered because they assumed that the flight crew and the FD were aware of the 

severity of what was felt in the back. 

In accordance with the applicable regulations and standards and as per the operatorís approved 

training program, all crew members, pilots and FAs had received training with respect to crew 

communication. As well, they attended crew resource management training, which also 

addresses crew communications. Such training is provided during initial and annual training.

During annual training, FAs and pilots also participate in joint crew communication training 

sessions during which communication skills and procedures are reviewed during simulated 

emergency situations. 

For initial FA training, the prescribed communication training objective is to teach the 

importance of, and the procedures for, effective communication in normal, abnormal, and 

emergency situations. Emphasis is placed on 

• the responsibility of FAs to provide complete and accurate information to the 

pilot-in-command to assist in decision making; 

• the potential hazards to flight safety if communication is not effective; and 

• the consequence of poor communication in aviation occurrences. 

FAs are taught that they must communicate any on-board safety concerns they may have 

witnessed or that may have been communicated to them by passengers. When communicating 

safety concerns during normal or abnormal operations, FAs are to adhere to the line of 

authority when possible. 
FACTUAL INFORMATION

TRANSPORTATION SAFETY BOARD 19 

However, if FAs notice an emergency situation developing, including unusual noises, they must 

contact the flight crew immediately via the interphone, stating their position and the nature of 

the problem. Training stresses that FAs should never assume that the flight crew is aware of 

everything that is happening. When information is not communicated, its potential value to 

flight safety is lost. 

There are procedures that set out the requirement for and the manner in which the FD must 

provide information to the FAs in abnormal and emergency situations. However, no such 

procedure or guideline was identified with respect to the FD collecting information from the 

FAs. As well, there is no requirement for the FD to provide flight crew with a structured 

briefing regarding the cabin environment in those situations. 

1.9 Aerodrome Information 

TSC961 used Runway 06 at Varadero/Juan Gualberto GÛmez International Airport (MUVR), 

Cuba, for the initial arrival, departure, and the subsequent return that night. Runway 06 is 

11 490 feet long and 148 feet wide, with an asphalt surface, and is served by an instrument 

landing system. Air Transat had maintenance personnel on site at the airport. 

The Fort Lauderdale/Hollywood International Airport (KFLL), Florida, has a set of parallel 

runways and a crossing runway. TSC961 was being guided to Runway 27R, which is 9000 feet 

long and 150 feet wide, with an asphalt surface, and is served by an instrument landing system. 

Air Transat had maintenance personnel available at KFLL, but customs services were not 

available at night. 

The Miami International Airport (KMIA), Miami, Florida, has four runways: 08/26L, 08/26R, 

09/27 and 12/30. Runways 12, 08R, 09, 26L and 27 are equipped with an instrument landing 

system. Air Transat did not have maintenance personnel on site in Miami. 

The Aircraft Rescue and Fire Fighting category of the three airports that could potentially have 

received TSC961 on the night of the incident exceeded the minimum response requirement for 

rescue and firefighting services for an aircraft the size of an Airbus A310. 

1.10 Flight Recorders 

1.10.1 Digital Flight Data Recorder 

The aircraft was fitted with a Honeywell/Sundstrand model universal flight data recorder 

(UFDR), part number 980-4100-DXUN, serial number 10623. The recorder used an eight-track 

Mylar tape. The recording system consisted of a data frame of 64 words per second, recording 

over 300 parameters, with a minimum capacity of 25 hours. The digital flight data recorder 

(DFDR) was received in very good condition. The recorder was disassembled and the tape was 

removed from the crash-protected memory cartridge for playback on an eight-track reel-to-reel 

instrumentation recorder at slower speed. A total of 25.3 hours of data were recovered from the 

recorder. 
FACTUAL INFORMATION

20 TRANSPORTATION SAFETY BOARD

1.10.2 Cockpit Voice Recorder 

The cockpit voice recorder (CVR) on the aircraft was a Loral Fairchild model A100-A, part 

number 93-A100-80, serial number 60662, and was received in very good condition. The 

recorder contained four 30-minute audio tracks. Tracks 1 and 2 contained the radio channels of 

the captain and co-pilot, track 3 was the cockpit area microphone channel, and track 4 contained 

public address/interphone and radio communications. The quality of the recording was good. 

The aircraft flew for 1 hour 17 minutes after the loss of the rudder. The CVR audio of the 

rudder-loss event was overwritten, resulting in the loss of information, including the noises 

heard in the cockpit during the rudder failure. The CVR recording started with the aircraft 

en route to Varadero, approximately 15 minutes before landing. The last 15 minutes were 

recorded on the ground in Varadero; the crew had not disabled the recorders. As a result of the 

TSB investigation into the Swissair Flight 111 accident in Nova Scotia, the Board, in 1999, made 

two recommendations that CVRs installed on aircraft be required to have a recording capacity 

of at least two hours (A99-01 and A99-02). As a result, aircraft manufactured after 31 December 

2002 must retain information recorded during the last two hours of aircraft operation. Aircraft 

manufactured before this date, however, continue to require CVRs with a minimum of 

30-minute recording capacity. 

There was no company procedure describing how to disable the recorders after landing. 

Current requirements in Canada are set out in TCís Aeronautical Information Manual (AIM) 

under General Information, Section 3.0, Transportation Safety Board of Canada, Subsection 

3.4.3, Protection of Occurrence Sites, Aircraft, Components and Documentation, which states in 

part 

Where a reportable incident occurs, the pilot-in-command, operator, owner 

and any crew member of the aircraft involved shall, as far as possible, 

preserve and protect: 

a) the flight data recorders and the information recorded thereon. . . 

The AIM is consistent with the Transportation Safety Board Regulations, Section 9 (1), Preservation 

of Evidence Respecting Reportable Accidents and Incidents. 

1.10.3 Direct Access Recorder 

On the aircraft, there is a direct access recorder (DAR) with an optical disk device having a 

storage capacity of 128 megabytes. The data frame had a configuration of 128 words per second, 

recording approximately 127 parameters, identical to that of the DFDR. Both the DAR and 

DFDR recorded flight data from identical sources; however, the recorded samples were not 

identical due to differing sample times. Data acquisition for both DAR and DFDR is handled by 

the digital flight data acquisition monitoring unit. The unit, manufactured by SAGEM, 

combines both the digital flight data acquisition unit function for the DFDR and the data 

management unit function for the DAR, feeding data to both recorders. The DAR optical disk 

was not originally requested by the TSB. Arrangements were subsequently made to transfer all 

applicable DAR data to the TSB. A total of 977 hours (not continuous) of DAR data applicable to 

the incident aircraft, including the incident flight, were obtained from Air Transat. 
FACTUAL INFORMATION

TRANSPORTATION SAFETY BOARD 21 

The DAR data were scanned for possible airborne and ground events. From the DAR data 

available, there were no significant events recorded that indicated lateral acceleration 

excursions, severe turbulence, or rudder doublets.8 Similarly, there were no significant ground 

events recorded that might indicate an impact to the rudder. 

1.10.4 Data Sampling Rates 

The DFDR and DAR data were manually time-synchronized and the data showed good 

correlation, with the exception of the lateral acceleration data for approximately two seconds at 

the start of the rudder-loss event (see Appendix A). The differing data were the result of a 

highly dynamic event. Both the DFDR and DAR sampled lateral acceleration at a rate of 4 Hz. 

At this rate, it was not possible to identify any lateral acceleration frequencies above 2 Hz.9 The 

determination of the specific frequencies involved in the rudder-loss event was not possible due 

to these low sampling rates of the recorded lateral accelerations. 

Under current regulations (Standard 625, Schedule 3, Aeroplane Digital Flight Data Recorder 

(DFDR) Specifications, of the Canadian Aviation Regulations [CARs], which are harmonized with 

Part 121, Appendix M, of the United States Federal Aviation Regulations), the sampling intervals 

for lateral and longitudinal acceleration are 4 Hz and vertical acceleration is 8 Hz. These rates 

meet the performance standards as recommended by the European Organisation for Civil 

Aviation Equipment (EUROCAE) minimum operational performance specifications (MOPS) for 

Crash Protected Airborne Recorder Systems (ED112). 

1.10.5 Filtering of Recorded Data 

The control surface position data recorded on the DFDR and DAR, including rudder position, 

were filtered by the system data analogue converter before recording. The filtered data are fed 

to the cockpit instrument displays, and the filtering process is designed to smooth out the data 

to remove unwanted spikes and prevent erratic indications. This same information is also 

recorded on the DFDR, and due to sampling and filtering, does not accurately represent the true 

control surface positions under dynamic conditions. Since the rudder loss on TSC961 was a 

dynamic event, critical information concerning the flight controls was potentially lost due to 

filtering. 

1.10.6 Summary of Flight Recorder Data 

At the time of the occurrence, the aircraft was in steady level flight at approximately 35 000 feet 

and 270 knots (Mach 0.795), with no significant control movements or turbulence. The aircraft 

had not exceeded any load or airspeed boundaries of its structural design envelope. 

 

8 Control input where one rudder pedal is depressed for a short period followed by 

immediate depression of the other rudder pedal for an equal period. 

9 Sampling theory indicates that, to measure certain frequency components, sampling 

must occur at a frequency that is twice that of the frequency components of interest 

(Shannonís theorem). 
FACTUAL INFORMATION

22 TRANSPORTATION SAFETY BOARD

Approximately 50 seconds after levelling off at FL 350, a dynamic oscillation in lateral 

acceleration occurred, lasting for approximately two seconds. This was the first indication in the 

DFDR/DAR data of the rudder-loss event. At the start of the oscillations, the lateral acceleration 

changed from +0.006 g to -0.073 g, indicative of a lateral force applied to the aircraft. Within one 

second of the dynamic oscillations in lateral acceleration, the heading decreased by 2° and the 

aircraft began to roll left from wings level. At the same time, the autopilot commanded aileron 

and spoiler deflections (right-hand spoilers 5, 6 and 7 extended) for right roll. The recorded 

rudder position indicated movement to the right from 1.2° left of neutral (0° with the 1.2° bias 

removed) to approximately 0.3° left of neutral (0.9° right of neutral, with the 1.2° bias removed). 

A pitch increase from 2 to 3° nose-up occurred, with a corresponding increase in vertical 

acceleration to +1.28 g. 

A yawing/rolling oscillatory mode, consistent with Dutch roll, commenced within two seconds 

of the rudder-loss event, as the dynamic oscillations in lateral acceleration decreased. At this 

time, a slight increase in altitude was followed by a decrease in pitch (from 3° to 2° nose-up), 

and a reduction in engine thrust (N1 decreased from 90 to 77 per cent). A gradual reduction in 

speed followed. The roll attitude reached 6° left-wing-low and then reversed direction. 

Approximately six seconds into the event, the recorded rudder position reached 6.2° right of 

neutral (approximately 7.4° right, with the bias removed). At the speed of 270 KIAS, the 

recorded rudder deflection was beyond the YD authority of ±3.7°. 

Approximately seven seconds into the event, the No. 2 autopilot was disengaged, followed 

immediately by disengagement of the auto thrust mode (manual throttle armed). The aircraft 

began to climb above FL 350 approximately 18 seconds into the event. With the autopilot 

disengaged, the oscillatory motion decreased in amplitude as the aircraft climbed through 

35 200 feet, and as airspeed decreased through 256 KIAS. The speed decreased to a minimum of 

248 KIAS. The altitude briefly peaked at 35 900 feet and the aircraft then began to descend. 

Autopilot No.1 command mode was briefly engaged as the aircraft descended through 

35 000 feet. With autopilot engagement, the yawing/rolling oscillatory motion increased in 

amplitude. After approximately 17 seconds, the autopilot was disengaged and the oscillations 

subsequently began to decrease in amplitude. As the aircraft descended through 27 900 feet and 

the speed decreased through 258 KIAS, the oscillatory motion ceased. 

1.11 Wreckage and Impact Information 

1.11.1 Miscellaneous Damage 

Some ceiling panels inside the passenger cabin had partially popped out of position. The 

displacement was very slight and did not impede passenger movement. The interior of the 

fuselage compartment behind the aft pressure bulkhead was inspected; there were no 

indications that the loads and vibrations associated with the rudder separation had caused any 

structural damage. 

The aircraft exterior was inspected, and there were no missing panels or structural components 

that may have come loose and struck the rudder. Apparent scrapes on the fuselage side, 

directed upward toward the tail, were determined to be poorly adhered, peeling paint and were 

not the result of foreign object damage (FOD). There was also blue-colour paint transfer visible 
FACTUAL INFORMATION

TRANSPORTATION SAFETY BOARD 23 

on the left side of the tail cone, just aft of the rudder, probably the result of a piece of the rudder 

striking the tail cone during separation. There was a series of puncture holes in the fuselage skin 

on the upper right side near the base of the VTP. These punctures were the result of the impact, 

during rudder breakup, of the mechanical fasteners that attach the rudder leading edge fairing 

to the rudder. 

1.11.2 Vertical Tail Plane Damage 

1.11.2.1 General 

Photo 2 shows the VTP and its rudder 

residuals being removed from the aircraft. 

The damage to the VTP trailing edge panels 

was generally limited to minor paint 

chipping. There was no damage suggesting 

that the rudder had been battered due to 

extreme travel from side to side. 

1.11.2.2 Main Attachment Fittings 

The VTP main attachment fittings were 

examined. On the fuselage side, these 

fittings are constructed of metal. Following 

the occurrence, they were subjected to visual 

and NDI, and no damage was found. The six 

CFRP main attachment fittings on the VTP 

side were subjected to ultrasonic NDI. Delamination damage was found in the two aft main 

attachment fittings. 

When the VTP is loaded in lateral bending, the two rear main attachment fittings are the most 

severely loaded. A full-scale test of the VTP conducted during the initial certification involved 

fatigue testing for three lifetimes followed by static testing, where the specimen main 

attachment fittings failed at over 1.9 times the limit load. In addition, three further static load 

tests conducted during an earlier investigation resulted in attachment fitting failures at greater 

than 1.8 times the limit load. It is noted that design ultimate load corresponds to 1.5 times the 

limit load and that, in order to meet certification requirements, a structure must withstand 

design ultimate load for at least three seconds. These tests demonstrated that the design exceeds 

certification load requirements. 

An analysis conducted in support of this occurrence investigation determined that, in order to 

cause the damage observed to the VTP main attachment fittings, the load experienced during 

the occurrence exceeded the design ultimate lateral fin bending load. However, it was not 

possible to quantify the precise load value attained. 

A 3D finite element analysis was conducted and included details of the main attachment fittings 

delamination on the occurrence aircraft, as reported by the NDI. This analysis was validated by 

a test on a damaged rear attachment fitting. It indicated that, when ultimate loads were applied 

to the model, strain levels varied only slightly from those of the undamaged model and were 

Photo 2. Removal of the VTP and a view of the 

position of the main attachment fittings 
FACTUAL INFORMATION

24 TRANSPORTATION SAFETY BOARD

well below the levels required to cause a fracture of the main attachment fittings. Therefore, the 

delamination had a minimal effect on the strength and stiffness of the main attachment fittings. 

Consequently, after the rudder-separation event, the aircraft was not in danger of losing the 

VTP during the flight either through loss of static strength or loss of stiffness. 

1.11.2.3 Hinge Arms 

Laboratory examination of hinge arm 1 found the right bolt under tolerance at the attachment of 

the hinge arm to the VTP. The fitting on the aft face of the VTP showed no visible signs of 

damage, such as cracks in the paint or sealant, but ultrasonic NDI found delamination around 

the mechanical fasteners. The forward ends of the hinge arms did not show any indication of 

upward travel as found at hinge arm 5. The rudder residuals were still attached. All three 

electrical grounding wires, two on VTP side and one on rudder side, were fastened with no 

indication of burning. There was no sign of any extreme side-to-side travel as with hinge arms 5 

and 6. There were impact marks on the rudder leading edge fairing caused by the hinge arms. 

This damage was restricted to the centre region, and the damage lines up with the hinge arms 

when the rudder is not deflected. 

At the hinge arm 2, 3, and 4 positions, the hinge arms are co-located with the hydraulic 

actuators. De-synchronization of the hydraulic actuators can result in force-fighting between 

them, which could lead to damage at their attachment points. Visual and NDI of the 

attachments did not find any sign of damage to the structure or to the mechanical fasteners. 

There was no indication of any structural damage that would degrade the stiffness of the 

actuator attachment. All the electrical grounding wires were fastened with no sign of burning. 

There was no indication of any extreme rudder travel as was the case at hinge arms 5 and 6. The 

z-strut located above hinge position 4 is designed to transfer vertical loads from the rudder into 

the VTP. The attachment fitting at the upper end of the z-strut showed no visual sign of 

damage, such as cracking of the paint or sealant, but there were paint chips on its top surface. 

This damage was probably caused by the upper end of the rudder separating and dropping 

vertically. 

At the hinge arm 5 position, the metal hinge arms were still securely fastened to the fitting on 

the aft face of the VTP rear spar, and the fitting showed no visible signs of damage, such as 

cracks in the paint or in the sealant. The forward ends of the hinge arms showed damage 

consistent with the hinge arms having moved upwards. The rudder-side hinge fitting was still 

attached along with a short section of the rudder spar, roughly 23 cm high by 26 cm wide. 

Ultrasonic NDI found no delamination around the mechanical fasteners that attach the CFRP 

fitting to the VTP rear spar, but the shim layer used to adjust the thickness of the CFRP fitting 

was mostly disbonded. 

All the electrical grounding wires were fastened with no sign of burning. The rudder-side 

bonding cable was badly frayed at the forward end, roughly at the position of the hinge bolt. 

The hinge arm was damaged by extreme side-to-side travel of the rudder hinge fitting, reaching 

nearly 90° deflection in each direction. 
FACTUAL INFORMATION

TRANSPORTATION SAFETY BOARD 25 

Manufacturerís drawings indicate that, at a rudder deflection of 45°, the rudder leading edge 

fairing cut-out strikes the hinge arm. At 60°, the rudder-side hinge fitting strikes the hinge arm. 

At 84°, the rudder side panel strikes the VTP trailing edge panel. There was no damage to the 

hinge arm where the leading edge fairing would have struck as the rudder passed through 45° 

of travel, and no damage to the VTP trailing edge where the rudder would have struck while 

passing through 84° of travel. The absence of such damage indicated that the damage from the 

extreme side-to side travel occurred after detachment of the rudder and would have started as 

the rudder passed through 60° of travel, progressing until the rudder reached about 90° of 

travel. 

At the hinge arm 6 position, the metal hinge arms were still securely fastened to the fitting on 

the aft face of the VTP. The fitting showed no visible signs of damage. The forward ends of the 

hinge arms did not show any sign of upwards travel. The rudder-side hinge fitting was still 

attached along with a short section of the rudder spar, roughly 15 cm high by 22 cm wide. 

Ultrasonic NDI found delamination around the mechanical fasteners that attach the CFRP 

fitting to the VTP rear spar. All the electrical grounding wires were fastened, with no sign of 

burning. The bolt that attaches the left side of the hinge arm to the VTP was nearly seized. 

The hinge arm had been damaged by extreme side-to-side travel of the rudder, where the 

rudder-side hinge fitting had struck the hinge arm, reaching nearly 90° deflection in each 

direction. The damage was less severe than that at hinge arm 5. Manufacturerís drawings 

indicated that, at a rudder deflection of 43°, the rudder leading edge fairing cut-out strikes the 

hinge arm. At 70°, the rudder-side hinge fitting strikes the hinge arm. At 84°, the rudder side 

panel strikes the VTP trailing edge panel. 

There was no damage to the hinge arm where the leading edge fairing would have struck as the 

rudder passed through 43° of travel, and no damage to the VTP trailing edge where the rudder 

would have struck while passing through 84° of travel. The absence of such damage indicated 

that the damage from the extreme side-to-side travel occurred after detachment of the rudder; 

the damage would have started as the rudder passed through 70° of travel and would have 

progressed until the rudder had reached about 90° of travel. 

At the hinge arm 7 position, the VTP-side CFRP attachment fitting had fractured and separated. 
FACTUAL INFORMATION

26 TRANSPORTATION SAFETY BOARD

1.11.3 Rudder Damage 

1.11.3.1 General 

As will be discussed later in this report, 

subsequent analysis found that the rudder loss 

during flight was progressive, and by the time 

the aircraft landed, most of the rudder had 

separated from the aircraft. Photo 3 shows the 

empennage after landing in Varadero. The 

separated pieces fell into the ocean and none 

were recovered. Rib 0 remained attached, as 

did the length of rudder spar up to hinge 

point 4. A small piece of rudder side panel 

from each side remained attached to the spar in 

the region between hinge points 2 to 4, and also 

in the corner where the spar meets rib 0. The 

leading edge fairings between hinge points 2 

and 4, and below hinge point 1, were still 

attached. The leading edge fairing between hinge points 1 and 2 had separated and some pieces 

were found jammed between the rudder and the VTP. At hinge points 5 and 6, small pieces of 

the rudder spar remained attached to the hinge arm. At hinge point 7, the VTP-side hinge 

bracket had fractured and separated, so none of the rudder remained. 

1.11.3.2 Detailed Description of Rudder Damage 

The front face of the rudder spar was cleaner at the hinge positions, consistent with what would 

be expected if the hinge areas had been cleaned for inspection. The aft face of the rudder spar 

was generally clean along its entire length, becoming slightly dirtier toward the bottom. There 

were drip stains oriented downwards originating at the lightening holes, consistent with 

normal in-service staining caused by dripping hydraulic fluid, corrosion inhibitor, or other 

fluids. There were dark stains observed on the interior of the side panels where the reinforcing 

bolts pass through the GFRP blocks at the hinge points. These stains originated at the bolts and 

progressed in a downward direction. The foils, which normally cover the lightening holes on 

the rudder spar, were missing, and the pattern of dirt around each lightening hole suggested 

that the foils had been removed for a considerable period before the occurrence. The top surface 

of rib 0 was visibly dirty. There were no stains on the interior bottom of the rudder to suggest 

that fluid had been pooling in the bottom of the rudder. The fluid drain holes and drain paths at 

the bottom of the rudder were not plugged. There were no stains in the honeycomb cells to 

suggest the presence of trapped fluids; however, very little of the honeycomb remained for 

examination. 

Cross-section examination of the rudder exterior skin revealed 10 layers of paint composed of 

primer, anti-static, filler, and topcoat. There was an accumulation of three paint re-sprays. It 

was calculated that the mass of the extra two re-sprays was approximately 19.3 kg. The total 

mass of a rudder with the nominal paint system is approximately 190 kg. 

Photo 3. Leftñside view of VTP and rudder 

residuals 
FACTUAL INFORMATION

TRANSPORTATION SAFETY BOARD 27 

The small amount of surviving rudder was examined for indications of contact with 

maintenance equipment, FOD, or damage by misuse. The only finding was a circular grinding 

mark on the exterior of the right rudder side panel. Cross-section examination determined that 

the grinding mark only extended down into the first few layers of paint, with no damage to the 

CFRP or discolouration due to heating. 

On each side of the rudder, there are three LPPs running chordwise. On the occurrence rudder, 

recent maintenance had involved the replacement of the lower right LPP in May 2004. A short 

length of this LPP remained and its fractured end was bent forward. A section of the side panel 

at this LPP was taken for subsequent laboratory analysis. 

The right side panelís outer face sheet exhibited many small multidirectional surface marks. 

There were similar marks on another aircraft (MSN 600), whose rudder was inspected and 

found undamaged. A section cut through these marks on the occurrence rudder revealed that 

they were cracks that had originated in the paint, caused by excessive paint thickness. It was 

further found that, when a paint crack was parallel to the direction of the CFRP fibres, the crack 

could extend down into the CFRP resin matrix. These cracks were limited to the matrix and did 

not damage the fibres. 

The CFRP face sheets had separated from the honeycomb core and the separation had a 

different appearance depending on whether it was an interior or exterior face sheet. The interior 

face sheets had generally separated from the honeycomb very cleanly near the bond line. 

However, the exterior face sheets had separated from the honeycomb in a very jagged manner 

with separations occurring at different depths in the honeycomb. 

Microscopic examination of the inner skin separation found that they were mostly cohesive 

failures within the bond line through the meniscus.10 Since the honeycomb had been so badly 

damaged during the occurrence, it was not possible to distinguish damage to honeycomb cells 

that may have been caused by freezing of trapped water. At the actuator locations, the interior 

CFRP face sheets had separated into four plies. There were no significant gaps in the coverage 

of the splice bond adhesive at the edges of the honeycomb sheets. In regions where a separation 

occurred near a honeycomb splice bond, the separation tended to occur in the weaker density 

honeycomb, and not in the bond line. Exposed regions of honeycomb that were no longer 

supported by the CFRP had tended to split into many small chordwise ìfingers,î each about 

25 to 50 mm wide. 

The rudder spar had fractured just above the hydraulic actuator attachments. Examination of 

fractured fibres indicated that the spar separated in an up and aft direction. The metal strip 

along the z-section at the front edge of each side panel had also fractured at this location, and 

examination revealed that it was a ductile overload failure. 

Photo 1 shows that there was no significant amount of rudder side panel still attached between 

hinge positions 1 and 2. In this region, more honeycomb remained on the right side, but more 

inner skin remained on the left side. The joint between the side panels and the spar, which uses 

blind mechanical fasteners, had not failed and the fasteners were intact. Examination of the 

 

10 The adhesive at the joint between the honeycomb and the face sheet forms a curved 

surface known as a meniscus. 
FACTUAL INFORMATION

28 TRANSPORTATION SAFETY BOARD

fractures at the joint between the front spar and the side panels revealed that the side panels or 

part of the side panels separated toward the outboard. Along the length of rudder spar between 

hinge points 1 and 2, the z-sections had fractured and separated along with the side panel on 

both sides. Since the leading edge fairing attaches to the z-sections, this explains why the 

leading edge fairing was missing in this region. The metal actuator attachment fittings at the 

hinge point 2, 3, and 4 positions did not show indications of damage, deformation, or looseness. 

The joint between the side panels and rib 0, which uses blind mechanical fasteners, had not 

failed and the fasteners were intact. There was a wipe mark across the top of rib 0, consistent 

with a fractured section of the left side panel moving towards the right and downwards. At the 

left side panel separation, more of the z-section had remained than on the right side. The 

fastener holes had been torn out towards the bottom, suggesting that the left side panel or part 

of the side panel separated from rib 0 in a downwards or outwards direction. The left side panel 

also had compression damage, suggesting the inboard skin moved downwards during 

separation. A failure in the z-section remains suggests that the outboard skin moved outboard 

during separation. 

At the right side panel separation, a length of the z-section had fractured and separated. Marks 

were observed on the remaining CFRP edge and their spacing corresponded to the spacing of 

the missing mechanical fasteners. These marks suggest that the right side panel or part of the 

side panel separated in an upwards direction. Examination of the right side panel showed that 

it separated from rib 0 in a tension flexion failure. The metal strips along the z-section had failed 

by overstress, a combination of tension and bending to the outside. There was a skin buckle on 

each side panel consistent with rib 0 moving upwards. There was a crack at the tip of rib 0 

whose orientation was consistent with rib 0 twisting to the right. 

Examination of other rudders as part of the fleet inspections following the occurrence found 

side panel damage at the hoisting points and the trailing edge fasteners. Since none of these 

areas of the occurrence rudder were recovered, it was not possible to examine them. 

Furthermore, since the entire upper end of the rudder was not recovered, the area around the 

1997 lightning strike damage could not be examined. 

Only a short section of the rudder spar at hinge point 5, roughly 23 cm high by 26 cm wide, 

remained attached. The lower surface of the spar section included the edge of a lightening hole. 

The rear reinforcement plate was still securely fastened to the spar and all its fasteners were still 

present and appeared undamaged. The separation between the honeycomb and the CFRP skin 

had generally occurred in the honeycomb, at varying depths and not along the 

honeycomb/CFRP bond line. On the front surface of the spar, the rudder-side hinge bracket 

had fractured. The fractured surface appeared typical of a tensile/bending overload failure with 

no indication of fatigue. Metallurgical analysis determined that the fittings were made of the 

correct aluminum alloy and temper. 

Only a short section of the rudder spar at hinge point 6, roughly 15 cm high by 22 cm wide, 

remained attached. The lower surface of the spar section included the edge of a lightening hole. 

The rear reinforcement plate was still securely fastened to the spar, and all its fasteners were 

still present and appeared undamaged. The separation between the honeycomb and the CFRP 

skin had generally occurred in the honeycomb, at varying depths, and not along the 

honeycomb/CFRP bond line. On the front surface of the spar, the rudder-side hinge bracket 
FACTUAL INFORMATION

TRANSPORTATION SAFETY BOARD 29 

had fractured. The fractured surface appeared typical of a tensile/bending overload failure with 

no indication of fatigue. Metallurgical analysis determined that the fittings were made of the 

correct aluminum alloy and temper. 

1.11.4 Chemical Attack and Contamination 

The rudder residuals were examined to study the possibility that they had been contaminated 

and degraded by exposure to chemicals. The manufacturer provides a list of approved 

consumables, as well as procedures to follow for the approval of materials not on that list. No 

indication was found that unapproved consumables were being used by the operator. There is 

no in-service experience to suggest that there was a systemic problem with chemical attack by 

approved consumables. During the material qualification process at certification, extensive 

testing was conducted to understand the interaction between the materials and possible 

contaminants, including hydraulic fluid. However, the bond between the honeycomb and the 

CFRP face sheets was not included in these tests because it was in the interior structure and 

considered to be sealed from such exposure. 

A water and hydraulic fluid mixture may react under certain concentrations to form phosphoric 

acid, which can attack epoxy resin creating irreversible damage to the core/face sheet interface. 

Microscopic examination of the rudder of aircraft MSN 361, which was known to be 

contaminated by hydraulic fluid, revealed that hydraulic fluid had attacked the matrix of the 

GFRP layer adjacent to the honeycomb, weakening the bond, but not leading to a disbond. 

Based on service experience with the aircraft MSN 361 and MSN 545 rudders, the access path 

for hydraulic fluid into the sandwich structure is around the blind fasteners at the front and 

bottom edges of the side panels. Three methods were used to search for the presence of 

hydraulic fluid contamination: energy dispersion X-ray spectroscopy (EDX), X-ray 

photoelectron spectroscopy (XPS), and infrared (IR) spectroscopy. The rudder of aircraft 

MSN 361, which was known to be contaminated by hydraulic fluid, was used to calibrate these 

three analysis methods. EDX testing of regions that were visibly stained by hydraulic fluid 

found roughly 2 per cent phosphorus content, and XPS testing found roughly 0.8 per cent 

phosphoric acidñester content. 

The top surface of rib 0 of the occurrence aircraft was visibly dirty, and an area at the front near 

the spar was analyzed. EDX results indicated 0.4 per cent11 phosphorus, considerably lower 

than the 2 per cent associated with the visibly contaminated region of the rudder of aircraft 

MSN 361. An area of inner skin (non-honeycomb side) from the left side panel front bottom 

corner was analyzed. EDX results indicated less than 0.1 per cent phosphorus, considerably 

lower than values from the aircraft MSN 361 rudder. An area of the inner skin (honeycomb 

side) from the left side panel front bottom corner was analyzed. EDX results indicated 0.3 per 

cent phosphorus, and XPS results indicated 0.18 per cent phosphoric acidñester, both 

considerably lower than values from the rudder of aircraft MSN 361. An area of inner skin 

 

11 EDX and XPS percentages are measured using different scales (that is, 1 per cent 

phosphorus measured using EDX does not equal 1 per cent phosphoric acidñester 

measured using XPS). 
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30 TRANSPORTATION SAFETY BOARD

(honeycomb side) from the right side panel front bottom corner was analyzed. EDX results 

showed less than 0.1 per cent phosphorus and XPS results indicated 0.07 per cent phosphoric 

acidñester, both considerably lower than the values from the aircraft MSN 361 rudder. 

Since the suspected hydraulic fluid ingress path was around the blind fasteners, specimens 

were taken around blind fasteners on both side panels at the front spar and at rib 0. 

Measurements were taken inside the sandwich structure at the inner face of the skin. EDX 

results for phosphorus on interior surfaces were all below the 0.1 per cent detection limit. EDX 

results on the external surfaces at rivet positions showed readings as high as 3.0 per cent. In 

addition, cross-section microscopic examination of the bond area did not reveal visual 

indication of chemical attack. Therefore, these results show the presence of hydraulic fluid 

contamination on exterior surfaces, but no indication of seepage into the structure. 

1.11.5 System Inspection and Testing 

The inspection of the rudder system on the occurrence aircraft showed that the rudder control, 

in cruising flight at 270 knots, would not have exceeded 7° of travel per side; the RTL control 

would have prevented it (RTL systems do not only limit the pedal inputs, but also limit the sum 

of inputs from trim, pedals or APYA, and YD). The autopilot was active at the time of the 

occurrence. The YD was also active (YD is active in manual flight also). The YD was restricted to 

moving the rudder to no further than 3.7° either side to compensate for the natural Dutch roll 

tendency of the aircraft. 

The last rudder servo-control synchronization check, performed by Air Transat maintenance on 

01 March 2005, revealed that no anomalies and no adjustments were required. In Varadero, the 

synchronization check showed no movement between the neutral position of the three 

servo-controls; there was no force-fighting between servo-controls. Therefore, the 

synchronization between the servo-controls was within the Airbus aircraft maintenance manual 

(AMM) parameters before and after the event. 

The inspection and investigation of the aircraft flight control system and related subsystem 

components was performed by the investigation team in Varadero after the occurrence and 

revealed no anomalies. The rudder control system was checked and tested for proper operation 

in Varadero with no anomalies found. The rudder servo-controls and actuating spring rods 

were then removed, inspected, and laboratory tested, and no anomalies that would have 

affected the normal operation of the rudder system were found. In addition, all safety features 

that are part of the servo-controls and spring rods to ensure safe operations in case of 

servo-control malfunction were operational. 

The free play was measured from hinge point 1 to hinge point 6; one free-play measurement 

was out of tolerance on the hinge line bearing at hinge point 2. In addition, 3 out of 10 VTPñside 

hinge arm bearings were partly seized but could still be rotated. Airbus specifies that the hinge 

arm bearing free play has no impact on the structural integrity of the rudder. Free play at the 

hinge arm bearings would result in detectable rudder vibration that will trigger a specific 

troubleshooting inspection process. No in-flight rudder vibrations had been reported. Operators 

are provided with some troubleshooting guidelines that list the most probable causes when 

vibrations are felt. However, the main cause of rudder vibration is play at the servo-control 

bearings, rather than at the hinge arms. The Airbus troubleshooting philosophy is such that, if 
FACTUAL INFORMATION

TRANSPORTATION SAFETY BOARD 31 

there are no findings of free play at the servo-control bearings, or if they are replaced and the 

situation is not improved, the operator will contact Airbus for investigation. Airbus in-service 

experience has confirmed the relevancy of this approach. 

1.11.6 High-Intensity Radiated Fields Investigation 

The possibility that high-intensity radiated fields (HIRF) interference could affect the normal 

in-flight operation of the rudder system was investigated. Any oscillation from the YD system 

at a frequency of 20 Hz, representing the difference of frequency between the radar and rudder 

synchronization frequencies, will be attenuated by the YD actuator and the three servo-controls 

that would be acting as filters. A review of the theoretical rudder deflection when the YD 

system is subjected to HIRF, assuming susceptibility at 20 Hz, led to the conclusion that the 

maximum rudder deflection would be less than 0.1°. Therefore, the investigation determined 

that the effect of HIRF would have a negligible impact on the rudder surface control. 

1.11.7 Examination of Pre-Occurrence Photos 

Photographs of the aircraft taken before the 

occurrence showed curious visual features on the 

rudder. Photo 4 shows an example of one of 

these features. It was taken 11 days before the 

occurrence and shows light-coloured vertical 

lines on the left side of the rudder below the 

hydraulic actuators. There were also earlier 

photographs that showed arc-shaped lines on the 

left side panel just aft of the hydraulic actuators, 

and white spots on the trailing edge. These 

features were not present on the most recent 

photographs. 

There was insufficient resolution in the 

photographs to conduct a photogrammetric 

analysis that would determine whether these 

vertical lines represented an out-of-plane 

deformation such as a disbond bubble. Since the 

vertical line features were observed on 

photographs taken on different days under 

different lighting conditions by different 

photographers, they were actual physical features on the rudder, and not simply reflections or 

dirt on the camera lens. 

The vertical line features first appeared in photographs starting in early 2003, and the aircraft 

was subject to all its regular inspections in the intervening time. Examination of other aircraft 

found that staining of the rudder near the hydraulic actuators was not unusual. Subsequent 

testing found that hydraulic fluid could dissolve the Air Transat tail decal material and analysis 

of a vertical streak on sister aircraft MSN 600 found the streak to be composed of a mixture of 

hydraulic fluid and dissolved decal material. 

Photo 4. Pre-occurrence photograph of aircraft 
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32 TRANSPORTATION SAFETY BOARD

1.12 Tests and Research 

1.12.1 General 

A series of laboratory examinations and tests were conducted on the residuals of the occurrence 

rudder, on other rudders, and on test specimens. In addition, analyses of flight dynamics, 

flutter, and radar data were conducted. This work was performed by the manufacturer at its 

facilities, with the participation of the involved national investigation authorities and specialist 

advisors from France, Germany, United States, and Canada. Test progress was monitored and 

validated by the investigation team. 

A number of the tests and analyses helped to eliminate avenues of inquiry and allowed the 

investigation to concentrate on others that proved germane. For the sake of clarity, the report 

focuses on this second group of activities. 

A small number of test programs developed as a consequence of this occurrence are being 

continued under the auspices of other entities. In each case, the objectives of the program have 

shifted away from this investigation to broader issues. 

1.12.2 Examination of Two Sister Rudders 

The rudders from aircraft MSN 592 and MSN 614 were removed from service and inspected at 

the manufacturerís facilities. These two rudders were selected because their side panels were 

produced by Soko as part of the same production batch as the occurrence rudder. Both of these 

rudders were subjected to visual inspection and elasticity laminate checker (ELCH) testing 

under TSB supervision. No damage was found. 

1.12.3 Elasticity Laminate Checker Test 

1.12.3.1 Elasticity Laminate Checker Test Description 

The ELCH test is an NDI method developed by Airbus to detect disbonded face sheets on 

honeycomb-core panels. The machine applies a vacuum to a small area of the outer surface of a 

panel and measures deflection. If there is a disbond of either the inner or outer face sheet, a 

greater-than-normal deflection will be measured. The advantage of this test method is that it 

can be used to find disbonds on the inner face sheets, which are not easily accessible. Airbus 

Report TN-EV37-579/90 (18 December 1990) describes the qualification results for the ELCH 

inspection procedure. 

1.12.3.2 ELCH 1óComplete Grid Examination of 24 Rudders of Similar Design 

The exteriors of the rudder side panels were marked with a 75 mm grid, and an ELCH reading 

was taken at each grid point. There were approximately 2000 test points per side on each 

rudder. This would permit finding defects down to a size of 120 mm in diameter. A total of 

24 rudders were tested, which included the flight cycle and flight hour fleet leaders. No 

disbonds were found. 
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TRANSPORTATION SAFETY BOARD 33 

1.12.3.3 ELCH 2óFocused Examination of Rudders 

Thirteen rudders close in serial number to rudder serial number HF1090 (aircraft MSN 361) 

were selected for further ELCH examination. The test area was a 50 mm grid along the complete 

z-section, around the hoist points, and along the trailing edge fasteners. No large disbonds were 

found. 

1.12.4 Examination of Other Rudders 

During the course of the investigation, some rudders of other operators that had been damaged 

were inspected in greater detail. 

The rudder from aircraft MSN 251 was accidentally damaged at its lower end by contact with 

maintenance equipment during normal scheduled maintenance activities. A repair was carried 

out, which included the replacement of the lower LPP using heat-assisted cure. A tap test was 

conducted following the repair, and no defects were found. An ELCH test was then conducted 

as a precaution. It found that the heat applied during the LPP replacement had not caused a 

disbond of the inner skin in the area of the LPP replacement. However, the area aft of the 

trailing edge fasteners around the repair to the contact damage failed a tap test, and subsequent 

laboratory examination found that the inner skin was almost completely disbonded within the 

repair area. 

The rudder from aircraft MSN 361 was 

damaged at the trailing edge by 

unintentional contact with a maintenance 

dock during maintenance. Subsequent 

inspection found a disbond on the left 

inner-face sheet near the front bottom 

corner, approximately 830 mm long by 

350 mm high. Stress analysis determined 

that the disbond was not caused by the 

impact with the maintenance dock. This 

internal disbond had not been detected by 

the AOT-1 (all operators telex ñ see 

Section 1.14.1) external tap test. Further 

examination revealed that hydraulic fluid 

had seeped into the left side panel around 

the blind fasteners at the front spar as shown 

in Photo 5, and that this fluid had weakened 

the bond between the honeycomb and the face sheet. In addition to the hydraulic fluid 

contamination, this rudder exhibited signs of water ingress around some trailing edge fasteners 

and at the leading edge immediately aft of the z-section. 

Approximately one litre of water was drained from the area around the right lower aft hoisting 

point of the rudder from aircraft MSN 378 during inspection before re-paint. Subsequent 

inspection revealed that excessive grinding of the old paint had resulted in exposed cells in the 

upper GFRP area. The manufacturer then subjected the entire area of both side panels to X-ray 

Photo 5. MSN 361 honeycomb showing stain 

caused by hydraulic fluid ingress 
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34 TRANSPORTATION SAFETY BOARD

examination and found additional water ingress at some of the trailing edge screws. Endoscopic 

examination around the hoist points revealed that honeycomb cell walls had been damaged or 

torn in some affected areas. 

During the AOT-1 inspection of the rudder from aircraft MSN 530, two indications of possible 

irregularities were found on the exterior face of the right side panel. These were 80 cm2 and 

670 cm2 in size. There was no visual indication of impact damage at these locations. Destructive 

laboratory examination determined that it was a core crush, not a disbond. It was further 

determined that the core crush had occurred during cure at original manufacture. A review of 

manufacturing records did not indicate any concession related to this damage. There was no 

indication that this manufacturing deviation had grown in service. Stress analysis determined 

that this deviation had a negligible effect on the rudder structural strength. This rudder was 

pre-modification 8408 and had lower density honeycomb, which was more susceptible to core 

crush. The occurrence aircraft was post-modification 8408, with higher density honeycomb. 

Trapped fluid was found in the lower nose area of 

the rudder of aircraft MSN 701 during the AOT-2 

inspection (see Photo 6). The fluid was not 

formally identified, but was reported to be mostly 

water with some dirt, possibly containing 

hydraulic fluid and sealant chips. The fluid level 

was reported to be up to the lower inspection 

hole. The rudder box aft of the spar was dry. The 

two drain openings in the nose section were found 

to be clogged. The tap test did not show any sign 

of disbond. An X-ray and thermography 

inspection of side panels adjacent to the fluid did 

not reveal any fluidity entrapped inside the 

sandwich. 

1.12.5 Fluid Contamination Program 

As a result of the findings on aircraft MSN 361, a separate investigation was launched by the 

National Transportation Safety Board (NTSB) with support from Airbus to address the issue of 

hydraulic fluid contamination and its effect on structural strength. This program is ongoing, but 

preliminary conclusions include the following: 

• Hydraulic fluid contamination does not have an immediate effect on mechanical 

strength; the effect takes time to develop. 

• The effect of a hydraulic fluid/water mixture is more severe than that of hydraulic 

fluid alone. 

• The effect of hydraulic fluid/water or hydraulic fluid is not reversible whereas the 

effect of water alone is reversible. 

Photo 6. Fluid accumulation in the nose of the 

rudder of aircraft MSN 701 
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TRANSPORTATION SAFETY BOARD 35 

1.12.6 Impact Damage Tests 

A series of impact tests was conducted to determine whether it was possible to cause a disbond 

between the honeycomb and the CFRP face sheets. Drop tower tests were conducted with 

spherical impactors of nose radii of 12.5 to 100 mm. The test panels had honeycomb cores 

40 mm thick with 32 kg/m3 density, and face sheets with one CFRP ply and one GFRP ply. 

Advanced drop tower tests were conducted with a spherical impactor of 2000 mm nose radius 

and a cylindrical impactor of 100 mm radius. These tests resulted in crushed core or face sheet 

perforation depending on the energy level, but no disbonds. 

Chapters 12-21-11 (Cleaning), 12-31-11 (Anti Icing), and 12-31-12 (Ice & Snow Removal) of the 

Airbus AMM warn that the impact pressure from fluid jets must not be more than 0.35 bar. 

High-pressure jets can potentially damage honeycomb structure. A survey of maintenance and 

de-icing facilities found that nozzle pressures in the range of 3.4 to 4.1 bars were common, with 

maximum up to 10.2 bars. However, the impact force that results from a spray jet depends on 

many factors, including the width of the spray and the distance of the nozzle from the aircraft. 

Airbus reports that it has no knowledge of damage occurring with pressures as high as 1 bar at 

the impact point. No fluid impact damage tests were conducted in the course of this 

investigation. However, it is considered that the misuse of a high-pressure spray would most 

likely result in damages similar to those caused by blunt impact, and would include core crush 

but no disbonds. It was reported that the occurrence aircraft was last de-iced on 19 February 

2005, approximately two weeks before the occurrence. 

1.12.7 Laboratory Tests of Rudder Residuals 

1.12.7.1 General 

Only a small amount of the occurrence rudder was recovered. The remaining honeycomb was 

examined and its density was consistent with production drawings. 

Light microscopy examination of specimens from the occurrence rudder showed the 

characteristic kidney-shaped cross-section of Toray T300 carbon fibres, consistent with the 

design specification. 

Specimens were taken from the CFRP face sheets of both side panels, rib 0, the front spar, and 

the hinge point 7 fin-side bracket. Only one deviation from drawings was found; an additional 

reinforcement layer on the front spar around one of the access holes was in the wrong 

orientation. A review of manufacturing records did not indicate any concession related to this 

deviation. The rupture of the front spar did not pass through this deviated lay-up. Stress 

analysis determined that this deviation had a negligible effect on strength and stiffness. Apart 

from this exception, all the lay-ups corresponded to manufacturerís drawings. 

1.12.7.2 Bond Between Honeycomb and Face Sheets 

The quality of the bond between the honeycomb and the face sheets is normally evaluated using 

a climbing drum peel test. Insufficient undamaged residuals of the occurrence rudder remained 

to conduct drum peel tests. The sole alternative approach to evaluating this bond was by 
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36 TRANSPORTATION SAFETY BOARD

examining the shape of the meniscus at the interface between the honeycomb and the face 

sheets. A number of cross-sectional specimens were taken to examine the meniscus. Meniscus 

formation was generally found to be similar to the baseline comparison panels. 

1.12.7.3 Bond at z-Section 

The front and bottom edge of each rudder 

side panel is trimmed with a z-section. 

Specimens of the inner skin bond taken 

near the z-section along the front edge of 

the left side panel of the occurrence rudder 

had a meniscus whose appearance was 

consistent with insufficient bonding 

pressure within a width of 20 mm. The 

investigation revealed that this condition 

could be caused by insufficient caul plate 

pressure during cure, either resulting from 

mispositioning of the z-section or 

accumulation of tolerances of the 

components. In Figure 6, the upper sketch 

shows correct orientation during cure. The 

lower sketch shows how dimensional 

tolerances can result in low bonding pressure next to the z-section. 

1.12.7.4 Inter-laminar Bond Between Face Sheet Layers 

The quality of the inter-laminar bond between the individual face sheet layers was evaluated by 

examining sections under microscope. No unusual features were observed that suggested a 

poor laminate quality. 

1.12.7.5 Splice Bond 

A compound was used to bond around the perimeters of the honeycomb blocks. Small gaps in 

the bonding were observed at some locations, but these had a minimal effect on the strength of 

the joint. IR spectroscopy tests confirmed that the correct splice bonding material had been 

used. 

1.12.7.6 Type of Resins 

Resins in the residuals were identified using a combination of IR spectroscopy tests and visual 

microscopic examination. Results found that the resins in the occurrence rudder were consistent 

with Hexcel F550 for the CFRP and EHG250 or EP112 for the GFRP. It was not possible to 

distinguish between EHG250 and EP112 since they are identical in formulation and 

manufacturing process. It was confirmed that approved resin types had been used. 

Figure 6. Low bonding pressure next to the z-section 
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TRANSPORTATION SAFETY BOARD 37 

1.12.7.7 Cure of Resins 

Differential scanning calorimetry tests were conducted to determine the adequacy of curing. All 

cures exceeded 98 per cent. For these resin systems, any value greater than 95 per cent is 

acceptable. 

1.12.7.8 Lightning Protection Plate Replacement 

The region of the recently repaired lower right LPP was examined. IR spectroscopy found that 

the adhesive used to fasten the new LPP was Hysol EA934 NA, as specified in the SRM. A 

cross-section specimen of the CFRP layer beneath the LPP was examined under microscope, 

and the only unusual feature was matrix cracking in the CFRP. To provide a baseline for 

comparison, an LPP was peeled from a comparison rudder side panel. A tap test of that 

comparison specimen found that the peeling had not resulted in a disbond, and a microscopic 

examination of the CFRP cross-section found matrix cracking similar to that in the occurrence 

rudder. 

1.12.7.9 z-Strut Paint Chips 

The area around the paint chips on the z-strut was examined by EDX. This analysis found no 

traces of titanium or steel residues in the paint chipped areas that would have been caused by 

contact with mechanical fasteners from the upper part of the rudder separating in a downward 

direction. 

1.12.7.10 Explosion Damage 

The rudder residuals and the VTP were visually examined by explosives specialists from the 

German police. There was no indication of damage or residue radiating outwards from an 

origin point, as would be the case if there had been an explosion. 

1.12.8 Double Cantilever Beam Tests 

The American Society for Testing and Materials (ASTM) Specification D5528 DCB test was 

adapted to determine the fracture toughness of the interface between the core and the face 

sheet. The standard test coupon is 50 mm wide, 220 mm long, and has an initial crack length of 

35 mm. For this particular investigation, the test coupon was modified from the standard to 

include 0.8 mm and 1.6 mm-thick aluminum doublers over the face sheets to achieve peel 

angles that were more representative of disbond growth. Static and fatigue tests of double 

cantilever beam (DCB) specimens are ongoing. 

1.12.9 Disbond Growth 

1.12.9.1 History of Earlier Design Aramid FibreñReinforced Plastic Rudders 

In the earlier design of the rudder, GFRP was used as the bridging layer between honeycomb 

and CFRP only in the reinforced region around the hydraulic actuators. Elsewhere, AFRP was 

used. The AFRP system resulted in poor bond strength of the interior skin. Those rudders 

experienced large in-service disbonds of the inner skins. Early AFRP rudders were of the design 
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38 TRANSPORTATION SAFETY BOARD

type that had two-part side panelsóupper and lower. In the rudder of aircraft MSN 237, 

growth of the disbond was stopped by this joint and did not progress beyond it. Disbonds were 

discovered in service during routine inspection and had not led to rudder separation or adverse 

aircraft performance. AFRP rudders are no longer in service. 

1.12.9.2 Temperature Effects 

Tests were conducted to investigate the effect of temperature on the out-of-plane strength in 

terms of climbing drum peel strength and fracture toughness (G1c) measured by a modified 

DCB test. Climbing drum peel tests were conducted at room temperature and at -55oC. At room 

temperature, the failures tended to be in the honeycomb core, whereas at -55oC, they tended to 

be in the interface bond line between the honeycomb and the face sheet. It was found that the 

peel strength at cold temperature was significantly reduced to about 45 to 67 per cent of its 

value at room temperature. The steep angle involved in this method of test is not representative 

of the shallow angle at which a disbond in the rudder would propagate; therefore, DCB tests 

were also conducted. 

DCB tests were conducted at room temperature and at -55oC. In all the tests, the failure tended 

to be in the honeycomb core, regardless of temperature. However, the cold temperature tests 

had fracture toughness values about 20 per cent lower. In addition, disbond growth at room 

temperature was steady and continuous, whereas at cold temperature, it was unstable. 

1.12.9.3 Vacuum Cycling Tests 

Test panels with various natural and artificial damage were placed in a vacuum chamber and 

cycled to a differential pressure of -0.7 bar to simulate ground-to-air cycles. The tests were 

accelerated, with each cycle lasting 90 seconds from ground to altitude and back to ground. In 

real time, a flight cycle lasts 540 minutes. One aircraft lifetime is 48 000 cycles. To provide 

conservative results, the test panels were fully sealed at the edges to impede breathing. 

The following results were obtained: 

• A specimen from the damaged left side panel of the rudder of aircraft MSN 361, 

which included the disbonded area (approximately 2256 cm2), was subjected to 

vacuum cycling at room temperature. The area of the damage almost doubled 

instantly at a pressure of 0.44 bar (absolute). The rapid propagation event was 

reported to be explosively loud and violent, resulting in some damage to the interior 

of the test chamber. During the test, a surface crack developed in the CFRP face sheet 

at the panel edge (Photo 7). This fracture halted any further damage growth by 

removing the pressure differential. 
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TRANSPORTATION SAFETY BOARD 39 

Photo 7. Damage growth after one vacuum cycle 

(rudder of aircraft MSN 361) 

• A specimen from the right side panel of aircraft MSN 361 rudder was also tested. 

Since it was undamaged, an artificial disbond (approximately 338 cm2) was 

introduced. It was subjected to vacuum cycling with a change of pressure (ΔP) of 

-0.7 bar at room temperature. The specimen completed the test program without 

failure. 

! Eight test panels were constructed, which included 24 and 32 kg/m3 core densities 

and 30 and 40 mm thicknesses. Damage was introduced by sweeping out a disbond 

damage with a knife. The artificial damages had diameters of 100 to 250 mm. When 

exposed to cyclic vacuum loading, four of these panels demonstrated slow steady 

growth of the disbond and eventually failed in rapid propagation. 

• A test specimen from the blunt impact test program, which had experienced crushed 

core but no disbond, was cycled for 10 000 flight cycles and showed no propagation 

of the damage. 

1.12.10 Computer Simulation of Disbond Growth 

Computer simulations were conducted using the LS-DYNA software to study the growth of the 

MSN 361 rudder disbond. The following conclusions were drawn: 

• The MSN 361 rudder side panel with the disbond damage as found at the time of 

detection was sufficiently large to achieve unstable growth at reduced pressure 

altitude. Since it did not propagate in flight, it is most probable that it experienced 

leakage during the period before detection. 

• A disbond of a size corresponding to the initial contaminated region of the MSN 361 

rudder did not grow under reduced pressure altitude. 

• When a disbond size corresponding to the initial contaminated region of the MSN 361 

rudder was combined with a disbond at the z-sections at the lower front corner of the 

side panel, unstable disbond growth occurred at reduced pressure altitude. 
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40 TRANSPORTATION SAFETY BOARD

1.12.11 Effects of Temperature and Moisture 

Liquid water trapped in the honeycomb will expand as it freezes, and the repeated freeze-thaw 

cycle can gradually damage the cell walls. In other rudders where water ingress into the 

honeycomb has been documented, this infiltration has tended to occur around the hoisting 

points or near the trailing edge fasteners. These areas are the most likely entry paths for water 

since these design features involve a break in the continuity of the CFRP face sheets. On the 

occurrence aircraft, these areas of the rudder were not recovered, and it was not possible to 

determine whether the occurrence rudder had experienced liquid water ingress into the 

honeycomb. 

The epoxy matrix in the face sheets and the NomexÆ core material can absorb moisture from 

humid air or condensed water. Moisture absorption degrades the mechanical properties, 

especially of face sheets that are matrix or matrix-interface dominated, particularly at elevated 

temperatures. At certification, static strength of the composite structure was substantiated 

assuming a worst case of maximum service moisture content and by conducting tests at 

elevated temperature. 

The temperature of an aircraft can reach extreme levels while parked on the ground due to the 

ambient temperature, solar radiation, and the colour of the aircraft. The rudder of the 

occurrence aircraft was painted dark blue, and its previous paint scheme had a black stripe 

across the rudder. United States Department of Transport paper DOT/FAA/AR-04/30 

describes tests involving CFRP laminate panels of different colours, whose temperature was 

measured during exposure to sunlight. The highest measured value was 82oC at an ambient 

temperature of 33oC. No thermal analysis was conducted for the A310 rudder, but a thermal 

analysis for the VTP fin box of another model of transport aircraft found maximum 

temperatures of 91oC in still air on the ground, and 76oC when moving. 

High temperature can have an adverse effect on the mechanical properties of composite 

materials. If the materialís glass transition temperature is exceeded while under load, the 

structure can experience plastic deformation. The onset glass transition temperature for 

F550/EHG250 resin is 102oC dry or 75oC wet. Given the occurrence rudderís age, its onset glass 

transition temperature had most probably reached its saturated or equilibrium state, and was 

closer to the 75oC value. 

1.12.12 Age-Related Structures and Materials Degradation 

The possibility was studied that there might be some unknown phenomenon at work that could 

cause a reduction in structural stiffness with age. Such a reduction in stiffness could result in a 

reduced flutter speed and lead to flutter. In 2004, Airbus conducted GVT in support of its MRTT 

program. The testing was conducted on an aged A310 aircraft (MSN 523) that had accumulated 

over 28 000 flight hours. This test aircraft had the same design of VTP and rudder as the 

occurrence aircraft. GVT results found that fin bending and rudder rotation frequencies of the 

MRTT test aircraft were consistent with those obtained during the original A310-300 

certification. No indication was found to suggest that stiffness had reduced with age. 
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TRANSPORTATION SAFETY BOARD 41 

Airbus Report TN-ESWCG-1181/02 documents an earlier investigation of material properties 

changes with age. The glass transition temperature (Tg) was measured for specimens of CFRP 

materials Hexcel F913 and F550, and GFRP material EHG250. These results were compared with 

those from specimens analyzed 14 years earlier during material qualification, and there was no 

indication of age-related Tg degradation. 

Climbing drum peel tests were conducted with specimens from the two unused baseline 

reference panels, which were roughly the same age as the occurrence rudder. Tests results were 

in the range expected, and there was no indication of age-related degradation in peel strength. 

1.12.13 Flight Dynamics Analysis 

1.12.13.1 Background 

To study the aircraft response during the occurrence, the Airbus A310 flight model was 

configured to simulate the aircraftís behaviour following a loss of rudder surface area. It was 

first necessary to determine the movement of the flight controls (ailerons, rudder, elevator, and 

spoilers). This was done using the merged DFDR and DAR data. It was also necessary to derive 

parameters not available directly from the DFDR, such as sideslip, and to modify existing 

parameters to take into account data latencies. The mass properties of the aircraft (that is, 

weight, inertia, CG) at the time of the incident were derived using the load and trim sheet for 

TSC961. Modifying the aircraft configuration to simulate rudder loss required modification of 

lateral-directional aerodynamic coefficients to take into account the loss of surface area. This 

modification was done by applying an adjustable ratio to the appropriate aerodynamic 

coefficients, allowing investigation of differing amounts of rudder loss. 

1.12.13.2 Initial Simulations 

Seven initial simulations were performed and the outputs were compared to the existing DFDR 

and DAR data to determine their applicability to the occurrence scenario. Of these seven 

simulations, four were performed with different amounts of rudder loss and with different 

aircraft mass properties. Three simulations were performed to evaluate additional aircraft 

behaviour related to the incident. The timeframe of interest was from 0701:57 (approximately 

4 seconds before autopilot 2 was disconnected) to 0702:14 (approximately 13 seconds after 

autopilot 2 was disconnected). The actual event was not simulated, since there were significant 

structural dynamics issues with which the flight dynamics simulator could not deal. 

The results of the simulations were presented graphically showing different lateral-directional 

flight parameters as determined by the simulation compared with flight data derived from the 

DFDR and DAR. In all cases, the properties of interest for comparison with the flight data were 

the magnitude and the frequency. Matching these two properties of interest for all aircraft 

parameters would indicate a simulation configuration similar to the aircraft at the time of the 

occurrence. 
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42 TRANSPORTATION SAFETY BOARD

The only simulation that reasonably matched the frequency and amplitude of the DFDR and 

DAR data was one in which there was a reduction of the useable rudder by 84 per cent and a 

reduction in the yaw moment of inertia by 10 per cent. The moment of inertia calculations for 

the aircraft at the time of the incident were only accurate to 10 per cent, which means that the 

mass properties for this simulation were within the limits of the calculations for the aircraft. 

1.12.13.3 Simulations to Determine the Lateral Force at Tail During the Event 

The initial event was characterized by significant lateral acceleration excursions, suggesting that 

a lateral force was applied to the right-hand side of the aircraft during the event. The Airbus 

flight dynamics simulator was unable to investigate structural dynamic aspects of the initial 

event but could provide an indication of the magnitude of the force required to initiate the 

Dutch roll motion. 

The only way the simulation could introduce this force into the aircraft was through rudder 

movement. The rudder input was constrained by the limitation of the rudder control jacks of 60°

per second. The simulated lateral acceleration resulting from the rudder input was then 

compared to the DFDR and DAR data. The rudder motion required to obtain the lateral 

acceleration was significant when compared to the DFDR and DAR data. A movement of this 

magnitude and duration would have been recorded on the DFDR and DAR; therefore, the 

rudder was not the source of the force. The force generated by this rudder motion had a peak 

value of 108 000 newtons (N) applied at 0701:54. Although the source of the force is not known, 

this does give an indication of the magnitude of the lateral force involved in the initial event. 

1.12.13.4 Simulations to Determine the Longitudinal/Vertical Force at Tail During the Event 

In addition to the lateral acceleration, increases in pitch angle, angle of attack, and vertical 

acceleration were observed in the DFDR and DAR data during the initial event, indicating a 

nose-up motion, whereas the stabilizer position was commanding a nose-down motion. The 

Airbus flight dynamics simulator was used to investigate the magnitude of longitudinal and 

vertical forces required to produce this motion. 

The forces were simulated through changes to the pitching moment coefficient, lift coefficient, 

and drag coefficient. The combination of these changes that best matched the DFDR and DAR 

data indicated that a downward vertical force of approximately 36 000 N at 28 m aft of the CG 

and a rearward horizontal force of approximately 35 000 N at approximately 9.5 m above the 

CG were applied to the aircraft approximately one second after the application of the lateral 

force described above or at 0701:55. 

1.12.13.5 Additional Simulations 

Additional simulations were conducted to better understand the results of the initial 

simulations. These additional simulations covered the time period from 0701:50 to 0702:15 and 

included the time of the excessive rudder deflection. Four simulations were performed with 

different rudder breakup scenarios. The simulation that best matched the DFDR and DAR data 

was 76 per cent rudder loss at 0701:28, increasing to 80 per cent rudder loss at 0701:32.5 to 
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TRANSPORTATION SAFETY BOARD 43 

0701:33, and increasing again to 84 per cent rudder loss at 0701:34.5 to 0701:35. An important 

result of these simulations was that the rudder movement recorded on the DFDR and DAR was 

made by a part of the rudder that was not aerodynamically effective. 

Combining the results from these simulations with the investigation of the lateral, longitudinal, 

and vertical forces suggests the following scenario: 

• At 0701:54, a lateral force of approximately 108 000 N was applied to the aircraft. 

• One second later, at 0701:55, a horizontal force of 35 000 N was applied 9.5 m above 

the CG, and a vertical force of 36 000 N was applied 28 m behind the CG. 

• The application of the horizontal and vertical force coincided with an initial rudder 

loss of 76 per cent. 

• Over a period of approximately seven seconds after the initial rudder loss, the rudder 

continued to break up, with 80 per cent rudder loss occurring between 0701:59.5 and 

0702:00, and 84 per cent rudder loss occurring at 0702:02. 

• The results also suggest that the excessive rudder deflection recorded on the DFDR 

and DAR at 0702:00 was made by a part of the rudder that was not aerodynamically 

effective. 

1.12.13.6 Yaw Damper Modelling 

After the initial event, the DFDR and DAR data indicated that the rudder was moving in a 

sinusoidal pattern, which frequency analysis showed to be similar to the frequency of the lateral 

acceleration. This would indicate that the movement of the rudder was connected with the 

aircraft movement. One theory proposed for the rudder movement was the operation of the YD. 

To confirm this theory, the YD system was modelled after information on YD mechanization 

provided by Airbus. 

When the YD model output was compared with DFDR and DAR rudder movement data, 

corrected to remove a rudder motion sensor bias and a time lag introduced by the system data 

acquisition concentrator, it showed that the rudder movement closely matched the predicted 

YD output. The only significant deviation occurred approximately five seconds after the initial 

event, where the DFDR and DAR data showed a large rudder motion that exceeded the 

mechanical limits of the rudder. This motion was attributed to the loss of rigidity in the rudder 

due to the breakup. 

1.12.14 Effect of Disbond Bubble on Static Aerodynamic Loads 

A rudder side panel disbond or in-plane core fracture, under the influence of aerodynamic 

loads, could result in the affected area bulging outwards with a bubble-like appearance. An 

analysis was conducted to determine the effect of such a bubble on static aerodynamic loads. 

The analysis was conducted using computational fluid dynamics based on geometrical 

information from the finite element analysis model. The analysis was repeated for rudder 

deflection angles of 1° and 6°, under the flight conditions that existed at the time of the 
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44 TRANSPORTATION SAFETY BOARD

occurrence. At 6° deflection, the disbond bubble caused a 1 per cent increase in rudder force and 

a 2.4 per cent increase in total VTP force. At 0° rudder deflection, which was the situation at the 

time of the occurrence, the disbond bubble had a negligible effect on rudder and VTP loads. 

Therefore, the presence of a disbond bubble would have a comparatively small effect on static 

aerodynamic loads. 

1.12.15 Flutter Analysis 

1.12.15.1 Analysis of a Dynamic Event 

The lateral load traces obtained from the DFDR and DAR were coincident except for a 

two-second period at the beginning of the occurrence. This suggested the possibility of a 

dynamic event. Since the sampling rate was only 4 Hz, it was not possible to determine the 

amplitude and frequency of the dynamic signal. Therefore, the signal was analyzed using a 

simplistic manual experimental curve-fitting approach. Using this method of analysis, it was 

not possible to find a unique solution. 

However, it was possible to fit the DFDR and DAR data by assuming both a high- and 

low-frequency component, a loss in damping, and by assuming that the first part of the signal 

was divergent and the second part was stable. Since there was a dynamic event, and flutter is a 

dynamic phenomenon, a flutter analysis was conducted to determine whether it was possible to 

find a scenario involving flutter before or after the rudder rupture, consistent with the available 

data. 

1.12.15.2 Method of Analysis 

A theoretical flutter analysis was conducted to study the effects of various failure scenarios on 

the aircraftís flutter characteristics. The analysis was done using a complete aircraft model, 

flutter computations with 70 modes, and 1 per cent structural damping. This same method of 

flutter analysis was used for the original certification, and at the time, was substantiated both by 

GVT and flight testing. The doublet lattice analysis method that was used does have limitations 

that may affect the accuracy of predictions in cases such as a disbonded skin panel on one side 

panel only. A preliminary investigation of advanced techniques was conducted, but the state of 

development of these approaches was not sufficiently advanced to be useful in the analysis. 

1.12.15.3 Baseline Analysis 

The baseline case represented an aircraft with no structural damage. The flutter analysis found 

that the basic coupling mechanism involved a 6.76 Hz fin bending mode and a 13.18 Hz rudder 

rotation mode. There were sufficient damping margins at both the occurrence speed of 

270 KCAS and the design dive speed of 406.5 KCAS. 
FACTUAL INFORMATION

TRANSPORTATION SAFETY BOARD 45 

1.12.15.4 Failure Scenarios 

Flutter analyses were conducted to study the effect of the following structural failure scenarios: 

• DisbondóThe 

scenarios involved 

varying degrees of 

disbond between the 

honeycomb core and 

the inner face sheet, 

as shown in Figure 7. 

Single-sided 

disbonds were 

studied on the left 

panel, the right 

panel, and both 

panels. Simulating a 

disbond between the core and the face sheet was accomplished in the analysis by 

reducing the stiffness of the face sheet to 0 or 20 per cent of its original value in the 

affected area. The analysis found that there was an increase in flutter tendency with 

greater disbonds, and that double-sided disbonds were more critical than 

single-sided ones. There was no significant difference between left-side and right-side 

disbonds. No flutter was observed when stiffness was reduced to 20 per cent. When it 

was reduced to 0 per cent, the double-sided 2.9 m2 disbond scenario and the 

single-sided 5.6 m2 disbond scenario resulted in violent flutter below 150 KCAS. A 

more detailed examination of the single-sided 5.6 m2 disbond scenario, within the 

range between 0 and 20 per cent stiffness, found that the flutter speed decreased as 

the stiffness was reduced, and that this decrease was steady and continuous down to 

about 5 per cent stiffness where flutter started. Below 5 per cent stiffness, the 

frequency and damping decreased rapidly leading to flutter conditions. 

• Free RudderóThis scenario simulated a chordwise fracture of the rudder just above 

the hydraulic actuators, resulting in a free upper rudder. This model demonstrated 

flutter of several modes below 270 KCAS, the speed at the time of the occurrence. 

• Trailing Edge Screws RemovedóThis scenario simulated the removal of all the 

trailing edge screws. The analysis found that the flutter behaviour of this model was 

nearly unchanged relative to the baseline scenario. The reason for this was that the 

trailing edge connection was still intact so the torsion cell was still closed. 

• Trailing Edge Connection RemovedóThis scenario simulated the removal, to 

varying degrees, of the trailing edge connection. The analysis found that the tendency 

to flutter increased with the increasing size of the damage, and that flutter within the 

flight envelope was possible with the larger damage areas. 

Figure 7. Disbond scenarios studied in flutter analysis 
FACTUAL INFORMATION

46 TRANSPORTATION SAFETY BOARD

• Fractured Front SparóThis scenario assumed that the rudder front spar was 

fractured above the hydraulic actuators. The analysis found that there was only a 

slight deterioration in the rudder-rotation/fin-bending flutter coupling relative to the 

baseline scenario. 

• Loss of Rudder Below Hinge 2óThis scenario assumed that the section of rudder 

below hinge point 2 was lost. The analysis found that the rudder rotation frequency 

increased significantly due to the missing rudder mass, but there was no coupling. 

• Upper-End DisbondóThis scenario assumed a large one-sided disbond at the upper 

end of the rudder. Twenty percent stiffness was assumed in the disbonded area. The 

analysis found that there was only a small change in damping relative to the baseline 

scenario. 

• Hinge FailuresóFive scenarios were evaluated involving the following failures: 

failure of hinge 1; failure of hinge 7; failure of hinges 1 and 7; failure of hinges 6 and 

7; and failure of hinges 5, 6, and 7. The analysis found that scenarios for failures of 

hinge 7, and failures of hinges 1 and 7 did lead to flutter, but this occurred beyond 

400 KCAS, outside the design envelope. The remainder of the scenarios did not show 

any flutter-critical couplings. 

• Hinge StiffnessóThree scenarios were evaluated involving the following failures: 

75 per cent of nominal stiffness, all 7 hinges; 50 per cent of nominal stiffness, all 

7 hinges; and 25 per cent of nominal stiffness, all 7 hinges. The analysis results 

showed that none of these scenarios presented any significant deterioration in 

coupling of the fin bending and rudder rotation modes compared to the baseline 

scenario. 

• Hydraulic Actuator StiffnessóThree scenarios were evaluated involving the 

following failures: 75 per cent of nominal stiffness, all 3 hydraulic actuators; 50 per 

cent of nominal stiffness, all 3 hydraulic actuators; and 25 per cent of nominal 

stiffness, all 3 hydraulic actuators. The analysis results showed that decreasing the 

hydraulic actuator stiffness increased the tendency to flutter. At 25 per cent stiffness, 

a flutter speed of 268 KCAS was calculated. 

• Single Hydraulic Actuator DisconnectionóThree scenarios were evaluated 

involving the following failures: disconnection of actuator at hinge position 2; 

disconnection of actuator at hinge position 3; and disconnection of actuator at hinge 

position 4. The analysis found no flutter with the disconnection of one actuator. 

Further analysis found that, when simulating a double hydraulic failure, the 

reduction of rudder rotation mode frequency was only 0.33 Hz and did not lead to 

flutter. 

• Actuator Stiffness versus Actuator DisconnectionóA comparative analysis of the 

two previous scenarios found that disconnecting one hydraulic actuator has the same 

effect as all three actuators having 66 per cent of their nominal stiffness. 
FACTUAL INFORMATION

TRANSPORTATION SAFETY BOARD 47 

• VTP Attachment FailureóA flutter analysis was conducted to determine the effect of 

failed VTP main attachment fittings on flutter. The analysis found that the 

disconnection of one or both of the VTP rear attachments did not provide a fluttercritical coupling. 

• Extra Rudder MassóThis design of rudder is not mass-balanced. Any condition that 

adds mass to the rudder and moves its CG aft, further away from the hinge line, has 

an adverse effect on the flutter margin. Excessive paint layers or fluid ingress are 

examples of conditions that could cause such an effect. 

- Extra Mass at Trailing EdgeóThe analysis simulated extra rudder mass by 

increasing the mass of the trailing edge screws. This analysis found that 142 kg 

distributed at the trailing edge was necessary to produce flutter at the occurrence 

speed of 270 knots. The analysis also found that, with 71 kg of extra mass, there 

was adequate damping at the occurrence conditions. 

- Extra PaintóThe analysis simulated extra mass distributed evenly over the 

surface of the rudder. This analysis found that, at the occurrence speed of 

270 knots, the addition of mass did not significantly reduce the damping. At high 

speeds, the addition of the initial 19.3 kg of extra paint significantly reduced the 

damping. 

- Pooled Fluid in RudderóA flutter analysis was conducted to determine the 

effect of pooled fluid in the rudder leading edge as was found in the rudder of 

aircraft MSN 701. It was found that, since this extra mass is concentrated so close 

to the hinge line, it had a negligible effect on flutter. 

1.12.15.5 Summary of Flutter Analyses 

The analyses found that flutter could occur within the certification envelope for three damage 

scenarios: 

• free rudder or free rudder section; 

• significant decrease in hydraulic actuator stiffness; and 

• significant reduction in rudder torsion stiffness by extensive disbond or failure of 

trailing edge connection. 

In all these scenarios, the necessary damage was significant and exceeded certification 

requirements. 
FACTUAL INFORMATION

48 TRANSPORTATION SAFETY BOARD

1.12.16 Time-Domain Flutter Analysis 

1.12.16.1 General 

The flutter analysis results described above were presented graphically by frequency and 

damping curves versus speed. In order to correlate this flutter analysis with the results of the 

flight dynamics analysis and the recorder data, which were presented as curves versus time, the 

following time-domain flutter analysis was conducted to observe the flutter amplitude versus 

time. 

1.12.16.2 Method of Analysis 

A time-domain flutter analysis was conducted for two of the failure scenarios: single-sided 

5.6 m2 disbond with less than 5 per cent stiffness of the original panel and free upper rudder. 

The aim was to produce time histories and correlate the resulting loads. The analysis was 

conducted in time steps of 0.001 second. The structure was excited with a lateral force of 

1 kilonewton (kN) at the VTP tip acting as a step pulse starting at time step 100 with a 

0.1-second duration. 

The following variables were examined: 

• rudder deflection (top, middle, bottom); 

• loads at all seven hinges; 

• forces at all three actuators; 

• loads at VTP attachments (front, middle, rear); and 

• lateral accelerations (cockpit, CG/DFDR, VTP attachment, VTP tip, rudder top and 

bottom) 

1.12.16.3 Results 

Upon excitation, the amplitudes of all 

the responses started to grow. 

Figure 8 shows a typical load 

response at the left rear VTP 

attachment. At design ultimate load, 

the maximum force acting at the aft 

VTP attachment is approximately 

700 kN, and based on the attachment 

fittings damage, it is known that 

loads during the occurrence did 

exceed ultimate strength. Therefore, 

the time step where the load peaked 

above 700 kN was selected as a 

reference point for each scenario. All 

the remaining monitored variables 

were measured within an envelope 

around the reference time step, since all the maxima did not occur at the same time. Their peak 

value within this envelope was recorded. Both investigated scenarios showed no contradictions 

Figure 8. Typical load response from time-domain flutter 

analysis 
FACTUAL INFORMATION

TRANSPORTATION SAFETY BOARD 49 

with respect to restraints like VTP attachment rupture load and maximum rudder rotation. The 

free upper rudder scenario did not exhibit any significant hinge loads in the fore/aft direction, 

and the VTP attachments would have failed before the hinges. Since this contradicts the 

observed damage, this scenario is excluded as a realistic damage scenario. 

The single-sided 5.6 m2 disbond with less than 5 per cent stiffness of the original panel scenario 

did exhibit significant hinge loads in the fore/aft direction. The time-domain flutter analysis 

determined that hinge 5 would be the first to fail and that it would fail before the VTP main 

attachment fittings were damaged. A second time-domain flutter analysis was conducted to 

study this scenario with a disconnected hinge 5. That analysis found that, after the failure of 

hinge 5, the loads of hinge 6 exceed allowable loads while VTP rear attachments reached a level 

above ultimate but below rupture. 

1.12.17 Summary of Flutter and Time-Domain Flutter Analyses 

Flutter and time-domain flutter analyses were conducted for a number of failure scenarios. One 

scenario, the single-sided 5.6 m2 disbond with less than 5 per cent stiffness of the original panel, 

provided a credible response and showed good correlation with the observed damage, the 

recorder data, and the flight dynamics analysis. 

1.12.18 National Transportation Safety Board Radar Data Analysis Study 

The NTSB conducted a trajectory analysis of the Florida Keys (KEY), Melbourne (MEL), 

Tamiami (TMA), and Cudjoe Key (CUD) air route surveillance radar data to identify the 

numerous tracks in the vicinity of the track of TSC961 at the approximate time the rudder 

separated from the aircraft. The aim was to distinguish larger pieces of debris from smaller 

shards and examine relative timing of events. The caveat was that radar has many variables that 

cannot always be reverse-engineered. No discernable track of an initial separation or a 

sustained track that would resemble a large piece of aeroplane part was evident. The tracks of 

the returns resembled numerous small pieces floating with the prevailing winds. 

1.12.19 Other Aircraft in Vicinity 

ATC records indicated that there were no other aircraft with an instrument flight rules discrete 

transponder code at a similar altitude within 60 nautical miles (nm) of the occurrence aircraft at 

the time of the event. The nearest aircraft was at FL 320 feet approximately 70 nm to the east. All 

other aircraft were considerably lower and further away. 

1.12.20 Space Objects 

Military authorities reported that, at the time and place of the occurrence, records did not 

indicate any man-made objects re-entering the atmosphere. Records were unavailable for 

natural objects re-entering the atmosphere. 
FACTUAL INFORMATION

50 TRANSPORTATION SAFETY BOARD

1.13 Organizational and Management Information 

1.13.1 Company Operations 

Air Transat is authorized by TC to provide the types of services specified in its air operating 

certificate (AOC). Part I of the AOC, in part, authorizes non-scheduled and scheduled 

international operations between Canada and points abroad and between points abroad using 

its A310 aircraft. 

1.13.2 Maintenance Organization 

Since February 2002, the occurrence aircraft was maintained by Air Transat, which operates a 

fleet of 14 wide-body aircraft (A310 and A330) under a quality safety management system 

(QSMS). Air Transatís QSMS is a component of a management control system that deals with 

quality and safety. Air Transatís QSMS includes an accountable executive, a safety management 

plan, safety oversight, training, quality assurance, documentation, and an emergency response 

plan. 

The Air Transat maintenance program in place at the time of the occurrence was approved by 

TC on 10 December 2004 under approval number Q-0188. Air Transat holds a TCñapproved 

maintenance organization (AMO) certificate, as per Part 5, Sub-Section 73, of the CARs, under 

AMO 32-87. The company has approximately 285 employees in the maintenance department. 

The company is capable of doing line and heavy maintenance, minor and major repairs, and 

modifications for Lockheed 1011, Airbus 310 series, Airbus 330 series, and Boeing 757 series 

aircraft. 

Air Transat also holds the privileges of specialized maintenance rating on sheet metal structure, 

composite structure, arc welding, avionic systems, and components, as per Section 573.02 of the 

CARs. The review of the Air Transat AMO showed that the organization has up-to-date 

Maintenance Control and Maintenance Policy manuals that outline the requirements for the 

technical operation. 

Air Transat was last audited by TC from 06 to 17 May 2002. Air Transat was not approved to 

perform C-checks on its Airbus aircraft. The last C-check on C-GPAT was completed by 

TAP Portugal in May 2004. Air Canada Technics has been used for C-checks on the occurrence 

aircraft as well as on other Air Transat aircraft. 
FACTUAL INFORMATION

TRANSPORTATION SAFETY BOARD 51 

1.14 Additional Information 

1.14.1 All Operators TelexóFleet Inspection of Rudder 

1.14.1.1 AOT-1óFleet Inspection of Rudder Exteriors 

As a result of this investigation, the 

aircraft manufacturer issued AOTs 

A310-55A2035, A300-55A6035, 

A330-55A3035, and A340-55A4030 on 

17 March 2005 for the inspection of all 

aircraft with part number A55471500 

series rudders (Figure 9). These AOTs 

were subsequently made mandatory by an 

Airworthiness Directive issued by the 

Direction GÈnÈrale de líAviation Civile 

(DGAC) of France, the State of 

manufacture. This inspection included 

222 A310s, 146 A300-600s, 6 A330s, and 

34 A340s. The aim was to verify the 

structural integrity of the rudder and its 

attachment by means of one-time visual 

and tap-test inspections. This included a 

GVI of the VTP rear spar aft face, DVI of the hinge arms and actuator support fittings, DVI of 

the rudder hinge fittings, and a tap test of the rudder side panels. The tap test was conducted 

around the exterior perimeter of the rudder side panels as per the normal five-year inspection, 

as well as the inspection of additional bands through the centre as shown in Figure 9. It is noted 

that an exterior tap test is unable to detect disbonds on the interior face sheets. Airbus obtained 

results from operators for over 80 per cent of the affected aircraft, with the following results: 

• Side panel disbonds and delaminations were found on a small number of aircraft but 

were all within SRM repair limits. 

• Damage at hoisting points was found but the damage was below the size requiring 

immediate repair and was within SRM repair limits. 

• Some corrosion was reported at hinge fittings but was assessed as having no impact 

on structural integrity. 

• Some free play was reported at hinge bearings but was assessed as having no impact 

on structural integrity. 

• No trends were observed in the damage data, which would suggest batch tendencies. 

• None of these findings approached a level that could result in failure in flight. 

Figure 9. Schematic of AOT-1 areas of inspection 
FACTUAL INFORMATION

52 TRANSPORTATION SAFETY BOARD

1.14.1.2 AOT-2óFleet Inspection of Rudder Interiors 

On 02 March 2006, the aircraft 

manufacturer issued a second series of 

AOTs (AOT A310-A552043, 

AOT A300-A556042, AOT A330-A553036, 

and AOT A340-A554031) for the inspection 

of all aircraft with rudder part number 

A55471500 series rudders. These AOTs 

were subsequently made mandatory by 

two Airworthiness Directives issued by the 

European Aviation Safety Agency (EASA), 

representing the State of manufacture. The 

aim was to verify the structural integrity of 

the rudder by means of a one-time tap-test 

inspection to the interior face sheets of the 

rudder side panels, as well as checking the 

drainage holes at the bottom of the rudder 

and cleaning hydraulic fluid from the 

external surfaces. Access to the interior of 

the rudder was through the inspection 

holes in the front spar of the rudder. It is noted that the access to inboard surfaces from the 

inspection holes is limited. Figure 10 shows the areas of inspection. Airbus obtained results 

from operators for over 90 per cent of the affected aircraft and no disbonds were found. 

Figure 10. Schematic of AOT-2 areas of inspection 
ANALYSIS

TRANSPORTATION SAFETY BOARD 53 

2.0 Analysis 

2.1 General Information 

TSC961 departed Varadero on a scheduled flight and was being flown by qualified crew 

members in accordance with applicable regulations and procedures. Documentation indicates 

that the aircraft was equipped and operated in accordance with applicable regulations and 

procedures. Weather and navigation aids were not considered as factors in this occurrence.

2.2 Flight Control System 

2.2.1 General 

The investigation of the aircraft flight control system and related subsystem components 

revealed that there were no control system anomalies or conditions that could have led to the 

breakup of the rudder. 

2.2.2 Ability to Diagnose the Source of Flight Control Difficulty 

Throughout the flight, the nature of the structural damage could not be precisely identified. 

Only when rudder inputs were made in the final stages of approach and landing did it become 

apparent that rudder response was abnormal and inadequate. 

The problem the crew was facing could be described as ìflight control difficulties of unknown 

origin.î There is no established procedure for this problem. The ambiguous nature of the 

symptoms made it difficult for the crew to assess the situation and to form a clear diagnosis of 

what had caused the control problems that they had experienced. 

2.2.3 Dutch Roll Recovery 

During their initial training, the pilots were shown a Dutch roll in the simulator; this was their 

only encounter with this situation. The integrity of any abnormal procedure checklist relies on 

the premise that only the correct procedure should be used for a given situation, and that the 

procedure should be completed in its entirety. On the A310, most procedures are displayed on 

the ECAM, and as each procedure item is properly completed, the item is removed from the 

display screen. Completing the associated procedure ensures best safety of the flight. There is 

no procedure in the QRH or on the ECAM that deals with Dutch roll. Only the Airplane Upset 

Recovery Training Aid gives general guidelines, and it does not deal specifically with systems 

such as the autopilot. There was insufficient guidance regarding Dutch roll recovery technique. 

Additional information may have prevented the crew from worsening the flight characteristics 

as was the case when the autopilot was re-engaged. 

2.2.4 Decision to Return to Varadero 

Shortly after the rudder loss, the crew took steps to descend and divert to a nearby airport. As 

the flight progressed, the Dutch roll decreased and then ceased as they descended. By the time 

the crew was in a position to complete an approach to either Miami or Fort Lauderdale, there 
ANALYSIS

54 TRANSPORTATION SAFETY BOARD

were indications that the aircraft would continue to fly normally. The determination of where to 

land was influenced by the following: there were no symptoms remaining related to the noise, 

vibration, or Dutch roll; there were no ECAM messages, warning lights or cockpit indications 

related to the control problems; the flight could continue at low altitude; and the company was 

better equipped to deal with the passengers and aircraft at Varadero. 

The investigation determined that the aircraft was not in danger of losing the VTP during the 

flight, either through loss of static strength or loss of stiffness. 

2.2.5 Decision not to Declare an Emergency 

The timely declaration of an emergency allows the aircrews to be helped to the greatest possible 

extent when dealing with abnormal or emergency situations. 

Declaring an emergency and clearly communicating the nature of the problem allows ATC to 

more easily coordinate between units and anticipate the needs of the crew in planning traffic 

management. It also serves to ensure that the flight will get immediate attention from 

controllers should the situation change. If ATC is aware that an aircraft is having control 

difficulties, it can incorporate this knowledge into its planning, and provide, among other 

things, more manoeuvring space and a longer final approach. Without this information, there 

could be unexpected and undesirable consequence such as the aircraft being unable to comply 

with ATC requests resulting in the requirement to execute a missed approach or other 

manoeuvre that would delay landing. 

Air Transatís operations manual recognizes that there are a wide variety of possible emergency 

situations and leaves it to the discretion of the crew as to when an emergency should be 

declared. In this occurrence, by the time the crew was in a position to communicate to ATC, 

they had full control of the aircraft and did not feel it necessary to declare an emergency. 

2.2.6 Crew Communication and Decision Making 

In abnormal and emergency situations, effective crew communication enhances effective pilot 

decision making in that decisions will be based on all available information. In this occurrence, 

information that might have been relevant, specifically the magnitude of the forces incurred in 

the aft galley and the subsequent adverse effects or a description of the noise (volume and type 

of sound) heard by the FAs, was not communicated to the flight crew. 

This information was not communicated because, given the intensity of the noise and vibration, 

all members of the cabin crew assumed that their experience was representative of the 

experience throughout the cabin. In addition, the captain assumed that, if potentially valuable 

information existed within the cabin, the FD would have such information and would provide 

it without being asked. 

The assumptions persisted despite training given to all crew members that emphasized the 

importance of sharing all information and not assuming other crew members have a complete 

understanding of events. Air Transatís written procedures provide a structured format for the 

flight crew to provide a briefing to cabin crew but there is no requirement for the FD to seek 

information from the rest of the crew. 
ANALYSIS

TRANSPORTATION SAFETY BOARD 55 

In this case, the provision of information related to the severity of events experienced in the aft 

cabin may not have had a significant impact on the decisions taken by the crew or the outcome 

of the occurrence. However, in other circumstances, the lack of information could have severe 

consequences. Procedures and practices that do not facilitate information sharing between crew 

members increase the likelihood that decisions will be based on incomplete or inaccurate 

information, placing passengers and crew at risk. 

2.3 Maintenance 

2.3.1 Maintenance Program 

No shortcomings were found in the Air Transat maintenance organization, facilities, 

procedures, the control of maintenance activities, or personnel qualifications. It was concluded 

that the aircraft was maintained in accordance with the approved maintenance program. 

2.3.2 Maintenance Records 

The investigation examined the maintenance records of Air Transat as well as those of the 

aircraftís previous owner. There were no remarkable issues with the rudder, apart from the 

repair of some minor lightning strike damage at the tip of the rudder approximately eight years 

prior. 

2.3.3 Hinge Bearing Condition 

Although the occurrence aircraft had one rudder hinge line bearing at position 2 that exceeded 

the AMM tolerances and 3 of 10 VTP side hinge arm bearings that were partly seized, there had 

been no reports of rudder vibration on the occurrence aircraft, no significant free play was 

measured on the residuals, and partially seized bearings could still be rotated. It is therefore 

concluded that the general condition of the rudder hinges indicates that they were not a factor 

in the occurrence. 

2.3.4 Adequacy of Rudder Inspection Program 

Daily and transit checks of the rudder consist of a GVI conducted from the ground. These 

inspections can only detect significant external damage because the rudder is as high as 15 m in 

the air, and the view of the rudder is partially blocked by the horizontal stabilizer. Therefore, 

the daily and transit checks are limited in their effectiveness to detect rudder damage. 

Air Transat conducts the 2-C check of the rudder every 30 months and it consists of a GVI 

conducted at armís length. This inspection can only detect externally visible damage. It does not 

assess the inside condition of the rudder, nor detect such anomalies as inner skin disbonds or 

fluid ingress. Therefore, the 2-C check is limited in its effectiveness. 

A tap test of the rudder side panels is conducted every five years. This tap test is limited to a 

40 mm-wide strip along the front edge of the rudder side panels, and a similar narrow strip 

along the lower part of the trailing edge. This inspection cannot detect damage in the side 

panels outside these limited areas, and the tap test will only detect large inner-skin disbonds. 

Therefore, the five-year inspection is limited in its effectiveness. 
ANALYSIS

56 TRANSPORTATION SAFETY BOARD

During the fleet-wide inspection that followed the occurrence, other NDI techniques were used 

such as ELCH, X-ray, ultrasonic, and thermography, and they demonstrated their effectiveness 

in finding damage not detected by tap test or visual inspection. For example, these techniques 

were responsible for finding cases of water ingress and inner-skin disbond. Although these 

alternate NDI techniques are available, and even used on other parts of the aircraft (that is, 

thermography is used to inspect the elevators), they are not part of the scheduled maintenance 

program for the rudder. Therefore, there are more effective NDI techniques available than those 

used by the current maintenance program. 

An effective inspection program must offer an acceptable probability of detecting damage 

before it can grow to critical size. The occurrence rudder did receive its five-year tap-test 

inspection in May 2001, and during the intervening period, it was visually inspected in 

accordance with the maintenance program with no finding. Nevertheless, the limitation of the 

inspection techniques does not guarantee that there was no damage present that could grow to 

critical size without detection. Therefore, the current inspection program is not adequate to 

detect damage to the rudder assembly in a timely and consistent manner. 

2.4 Recorders 

2.4.1 Cockpit Voice Recorder Duration 

The lack of information from the 30-minute CVR regarding the rudder-loss event, including the 

noises heard by the cockpit and cabin crew and the associated vibrations, hindered the 

investigation. A two-hour CVR would have captured the sounds of the vibrations on the 

cockpit area microphone, providing important information on the vibration frequencies. The 

lack of adequate data increased the workload of investigators and hampered their ability to 

obtain a timely, complete, and accurate understanding of the event. 

2.4.2 Digital Flight Data Recorder Data Sampling 

A two-second highly dynamic event was identified when the DFDR and the DAR data were 

merged and lateral accelerations were compared. The determination of the frequencies involved 

was not possible due to the low sampling rates of the recorded accelerations. Although the 

sampling rates meet current performance standards required by regulation, they were not 

adequate for capturing the highly dynamic conditions that may exist during an accident. 

2.4.3 Digital Flight Data Recorder Filtered Data 

To investigate the rudder failure and resulting aircraft response, a performance analysis was 

undertaken that required accurate control surface position data. The DFDRñrecorded control 

surface position data are not the raw sensor data. The raw sensor data are filtered by the system 

data analog converter before being recorded. The probable rudder position history was 

calculated using filtered information. The analysis suggested that the filtering produced 

0.4-second data latency and reduced the amplitude by up to 1° during the initial high-frequency 

oscillations. The rudder position filtering and the necessity for additional analysis adversely 

affected the accuracy and effectiveness of the investigation efforts. 
ANALYSIS

TRANSPORTATION SAFETY BOARD 57 

2.4.4 Preservation of Recorder Information 

In this occurrence, disabling the CVR at engine shutdown would not have prevented the loss of 

information recorded at the time of the rudder failure, because the CVR was a 30-minute device, 

and the event occurred more than one hour before landing in Varadero. Nevertheless, there are 

situations for which securing the recorders after landing will preserve valuable evidence, as 

past investigations have shown (TSB reports A00A0185, A00P0040, and A01W0117). No 

procedure for disabling recorders after landing was available to the crew. 

2.5 Analysis of Rudder Failure Mechanism 

2.5.1 General 

The investigation studied the possibility of the rudder failing either as a result of a static loading 

phenomenon or of a dynamic loading phenomenon. 

2.5.2 Static Loading Phenomenon 

2.5.2.1 Large Rudder Deflection 

The rudder control system was operating correctly and no indications were found that the 

rudder made a deflection beyond authorized deflection limits. It is concluded that the failure 

was not caused by a large rudder deflection. 

2.5.2.2 High Static Load 

The investigation reviewed the static load tests, sub-component tests, and damage tolerance 

tests that had been conducted during initial certification, and concluded that the rudder was 

designed with adequate strength to react to static loads encountered within the structural 

design envelope. A review of recorder information revealed that the aircraft was operating 

within its design envelope and that it did not experience a high load event either on the 

occurrence flight or on an earlier flight. It is concluded that the failure was not caused by a high 

static load. 

2.5.3 Dynamic Loading Phenomenon 

The analysis of the lateral load signal from the DFDR and DAR found an indication that the 

occurrence was associated with a dynamic event. The following were examined as possible 

causes of this dynamic event. 

2.5.3.1 High-Frequency Control Input 

The investigation of the aircraft systems did not reveal any conditions that would have resulted 

in a control-induced dynamic event. DFDR and DAR recordings did not show any indications 

of high-frequency control movement in the period leading up to the occurrence. It is concluded 

that the dynamic event was not caused by a high-frequency input from the control system. 
ANALYSIS

58 TRANSPORTATION SAFETY BOARD

2.5.3.2 Flutter 

The lateral load signals recorded, the damage to the VTP main attachment fittings, the damage 

to the rudder hinge arms at positions 5 and 6, as well as the noise and vibrations felt during the 

event are consistent with flutter. 

2.5.4 Possible Causes of Flutter 

2.5.4.1 Flutter without Prior Structural Deviation 

Flutter analysis confirmed that a rudder with no structural deviations will not flutter within the 

design envelope. The investigation showed that the rudder was operated within the design 

envelope; therefore, the rudder did not experience flutter without a prior structural deviation. 

2.5.4.2 Flutter Following Structural Deviation 

The investigation revealed that rudder imbalance and hinge free play would not have led to 

flutter. It was determined that the most probable cause of flutter was a large disbond-type 

damage. The presence of additional minor factors such as possible water trapped in the 

honeycomb and excess paint would marginally reduce the size of the disbond necessary to 

cause the flutter. 

2.5.5 Growth of Rudder Damage 

Vacuum cycling tests conducted resulted in damage growth. Therefore, the pressure differential 

between the air inside the honeycomb and the reduced external air pressure at cruise altitude 

might have acted as the driving force for the growth of core/face sheet separations or in-plane 

core fractures. 

This particular rudder design does not include any damage growth arrest features in the side 

panels such as a mechanical barrier. Once damage starts to grow, it can continue to grow until it 

reaches critical size. Such a feature was not specifically demanded for certification. 

2.5.6 Possible Causes of Rudder Damage 

2.5.6.1 Manufacturing Process 

Examination of the rudder residuals determined that correct resin systems had been used and 

that the degree of cure was adequate. Although this analysis was based on the examination of 

only the small amount of the rudder that remainedósince each side panel, spar, or rib is cured 

as a unitóthe state of the small residuals is representative of the overall components. Therefore, 

the entire rudder was most probably constructed using the correct resin system and was 

adequately cured. 

Some non-conformities in the occurrence rudder were found by quality assurance at 

manufacture, and corrected. Since most of the rudder was missing, it was not possible to 

examine each of these locations on the residuals and formally exclude them as a cause of the 
ANALYSIS

TRANSPORTATION SAFETY BOARD 59 

occurrence. However, a review of the repair schemes, repair procedures, and quality assurance 

used to correct these non-conformities was conducted, and did not reveal any inadequacies. It is 

considered improbable that these particular repairs led to the occurrence. 

The residuals showed indications of possible insufficient bonding pressure during cure at the 

bond between the honeycomb and the inner skin along the edges of the z-section of the left side 

panel within a width of 20 mm. Subsequent investigation revealed that low bonding just aft of 

the z-section could be caused by insufficient caul plate pressure during cure as a result of 

mispositioning of the z-section, or of adverse accumulation of tolerances. This deviation would 

not necessarily be open to the outside air and could grow by vacuum cycling loads into a 

disbond. Further computer analysis determined that it was possible for such a disbond to grow 

under the influence of vacuum cycling. This deviation would have been present since 

manufacture and is a possible cause of the initial damage to the rudder. 

2.5.6.2 Material Degradation 

It is improbable that the occurrence rudder was damaged due to degradation by fatigue, aging, 

chemical contamination, or exposure to high temperature. 

2.5.6.3 Mechanical Damage 

It is improbable that the initial damage to the rudder was caused as a result of grinding 

damage, water ingress in the honeycombs, a seized hinge point, or the LPP repair. 

Impact damage tests conducted during the course of this investigation demonstrated that, 

although blunt impact could cause core crush, it could not cause disbonds. Subsequent vacuum 

cycling tests demonstrated that core crush damage did not grow. Therefore, it is improbable 

that the initial damage was caused by a blunt impact. Since the misuse of high-pressure spray 

jets would likely result in damage similar to blunt impacts, it is also improbable that the initial 

damage to the rudder was caused by the misuse of high-pressure spray jets. 

There was no evidence of impact damage on the residuals, but only a small amount of rudder 

survived for examination. The investigation was not able to discount the possibility that the 

rudder may have experienced a discrete event12 that resulted in significant damages either on 

the ground or in flight. 

The lightning strike to the tip of the rudder approximately eight years prior was a discrete event 

that could not be discounted as a possible cause of the initial damage since the entire upper end 

of the rudder was missing and could not be examined. 

 

12 The term discrete event is used to refer to an occurrence that could have resulted in 

significant damages to the rudder either on the ground or in flight. Possibilities include, 

but are not limited to, impact by foreign objects or a lightning strike. Such damage cannot 

be discounted as a possible cause because only a small amount of the rudder survived for 

examination, and because visual inspection of the rudder from the ground is limited. 
ANALYSIS

60 TRANSPORTATION SAFETY BOARD

2.5.7 Failure Scenarios 

2.5.7.1 Summary of Important Points 

• The dynamic event was most probably caused by rudder flutter. Flutter analysis 

showed that the rudder would only flutter under the occurrence conditions if it were 

damaged. Therefore, the rudder was most probably damaged. Flutter analysis 

determined that the amount of damage necessary to cause flutter was significant. The 

occurrence conditions had been experienced by the aircraft many times previous, but 

the rudder had not experienced vibration, flutter, or failure. Therefore, the original 

damage was most likely small and grew over time. 

• Although the mechanical cycling tests that were conducted during the original 

certification showed no damage growth, vacuum cycling tests conducted during this 

investigation demonstrated that it is possible for an initial damage to grow due to 

pressure differential associated with altitude. 

• Results of investigative activity did not support the likelihood of a blunt impact 

scenario. A discrete event could not be discounted as a possibility since most of the 

rudder was not recovered for examination. Positive indications were found 

suggesting the possibility of a weak z-section bond at the interior lower front of the 

left side panel, and analyses showed that such damage could grow under vacuum 

loads. 

• The interior skin is not easily accessible for inspection, and at the time of the 

occurrence, there was no inspection program for the inner skin. If the damage had 

been growing on the interior, it would not have been found by the existing 

inspections. A weak z-section bond would manifest itself in the bond of the inner 

skin. 

• In the time leading up to, and at the time of the occurrence, the aircraft was neither 

manoeuvring nor experiencing turbulence. Therefore, the most significant load on the 

rudder would have been pressure differential between the core interior and the 

ambient air at altitude. This suggests that differential pressure may have driven the 

event. 

• The investigation revealed that the first event in the occurrence sequence was a loud 

bang. Vacuum cycling damage growth tests found that, when the damage reached 

critical size, it grew explosively with a sudden violent release of energy, which 

caused a loud noise. 

• The vacuum cycling tests found that explosive damage growth was so violent that it 

damaged the interior of the test chamber. In flight, such a violent event could 

possibly damage the opposite side panel. The flight dynamics analysis found that 

there was a large lateral force at the rudder, possibly the explosive damage growth in 

one side panel striking the other side panel. 
ANALYSIS

TRANSPORTATION SAFETY BOARD 61 

• The sudden explosive growth of damage in one side panel and possible collateral 

damage to the adjacent panel would have resulted in a sudden reduction of rudder 

stiffness. The flutter analysis indicated that such a loss in stiffness could lead to flutter 

under the occurrence conditions. 

• The time-domain flutter analysis for the large disbond case found that, shortly after 

the initiation of flutter, a large aft force at hinge 5 would exceed the failure load. This 

is consistent with the flight dynamics analysis that found that, shortly after the initial 

event, there was a large aft and downward tug on the tail. This is also consistent with 

the damage observed at hinge 5. 

• The time-domain flutter analysis for the large disbond case also found that the next 

hinge to fail would be hinge 6, and that the loads on the rear VTP main attachment 

fittings would exceed ultimate strength. This is consistent with the damage observed 

at hinge 6 and at the VTP attachments. 

2.5.7.2 Most Likely Failure Scenario 

Some time before the occurrence flight, a disbond or in-plane core fracture occurred. The cause 

of this initial damage may have been a discrete event or a weak bond at the z-section. An 

indication of weak bonding was found at the z-section along the interior lower front of the left 

side panel. This damage then grew, possibly due to reduced pressure cycling loads associated 

with normal flight, without detection until it reached a critical size. 

During the occurrence flight, having reached the critical size, the damage rapidly propagated, 

resulting in a loud and sudden explosion of the skin. This separation could have damaged the 

opposite side panel and created a large sideways force on the empennage. The resulting sudden 

reduction in torsional stiffness led to the onset of rudder flutter. About one second later, there 

was a large aft and downward force associated with failure of the upper hinge points, as the 

rudder separated. The rudder-separation event lasted about seven seconds, after which only 

16 per cent of rudder effectiveness remained. During the remainder of the flight, more rudder 

pieces separated, and the aircraft landed with no aerodynamically effective rudder remaining. 

CONCLUSIONS

TRANSPORTATION SAFETY BOARD 63 

3.0 Conclusions 

3.1 Findings as to Causes and Contributing Factors 

1. The aircraft took off from Varadero with a pre-existing disbond or an in-plane core 

fracture damage to the rudder, caused by either a discrete event, but not a blunt 

impact, or a weak bond at the z-section of the left side panel. This damage 

deteriorated in flight, ultimately resulting in the loss of the rudder. 

2. The manufacturerís recommended inspection program for the aircraft was not 

adequate to detect all rudder defects; the damage may have been present for many 

flights before the occurrence flight. 

3. This model of rudder does not include any design features in the sandwich panels to 

mechanically arrest the growth of disbond damage or in-plane core failure before the 

damaged area reaches critical size (such a feature was not specifically demanded for 

certification). 

3.2 Findings as to Risk 

1. A cockpit voice recorder with a 30-minute recording capacity was installed on the 

aircraft, and its length was insufficient to capture the rudder-loss event, resulting in 

critical information concerning the rudder failure not being available to investigators. 

2. There was no published procedure for disabling the recorders once the aircraft was 

on the ground; valuable investigation information can be lost if the data are not 

preserved. 

3. The sampling intervals for lateral and longitudinal acceleration captured by the 

digital flight data recorder were insufficient to record the highly dynamic conditions 

present at the time of the occurrence. This resulted in incomplete information being 

recorded. 

4. The rudder position filtering and the necessity for additional analysis adversely 

affected the accuracy and effectiveness of the investigation efforts. 

5. There are insufficient published procedures available to flight crew members to assist 

in recovering from a Dutch roll. 

6. Declaring an emergency and clearly communicating the nature of the problem allows 

air traffic control to more easily coordinate between units and anticipate the needs of 

the crew in planning traffic management. 

7. Procedures and practices that do not facilitate information sharing between crew 

members increase the likelihood that decisions will be based on incomplete or 

inaccurate information, potentially placing passengers and crew at risk. 
CONCLUSIONS

64 TRANSPORTATION SAFETY BOARD

3.3 Other Findings 

1. Throughout the event, the crew received no electronic centralized aircraft monitor 

message relating to the control problem that the aircraft had experienced, and there 

were no other warning lights or cockpit indications of an aircraft malfunction. 

2. After the rudder-separation event, the aircraft was not in danger of losing the vertical 

tail plane during the flight, either through loss of static strength or loss of stiffness. 
SAFETY ACTION

TRANSPORTATION SAFETY BOARD 65 

4.0 Safety Action 

4.1 Action Taken 

4.1.1 Transportation Safety Board of Canada 

4.1.1.1 TSB RecommendationsóAirbus Composite Rudder Inspection Program 

The separation of the rudder from Air Transat Flight 961 and the damage found during the 

post-occurrence fleet inspections suggest that the current inspection program for Airbus 

composite rudders may not be adequate to provide for the timely detection of defects. In 

addition, preliminary tests demonstrating that disbonds can grow due to altitude-related 

pressure differential suggest that increased attention is warranted to mitigate the risk of 

additional rudder structural failures. The consequences of a rudder separation include reduced 

directional control and possible separation of the vertical tail plane (VTP). 

Therefore, on 27 March 2006, the Board recommended that: 

The Department of Transport, in coordination with other involved 

regulatory authorities and industry, urgently develop and implement an 

inspection program that will allow early and consistent detection of 

damage to the rudder assembly of aircraft equipped with part number 

A55471500 series rudders. (A06-05, issued March 2006) 

On 14 June 2006, Transport Canada (TC) responded to Board Recommendation A06-05. TC 

concurs with the TSB suggestion that the current A310-300 inspection program may not be 

adequate to provide timely detection of defects to the rudder assembly. 

Specifically, TC has indicated that the following corrective actions will be taken: 

• A letter will be sent to Airbus and the Direction GÈnÈrale de líAviation Civile 

(DGAC) of France detailing the results of additional inspection on a Canadianregistered A310-300 series aircraft. 

• TC will recommend that a detailed inspection of the drainage path of the rudder for 

blockage be added to the current inspection program to ensure that there is adequate 

drainage. 

• TC will request that Airbus review the current inspection program for the vertical 

stabilizer and rudder assembly for the A300/A310 aircraft series. 

• A tap test is potentially not effective in determining small areas of delamination or 

disbond of composite materials; therefore, TC is working with the National Research 

Council of Canada to identify more suitable inspection techniques to detect failures in 

composite materials. 
SAFETY ACTION

66 TRANSPORTATION SAFETY BOARD

• To better identify failures in composite material, TC will coordinate with the 

International Maintenance Review Board to review the logic used in developing 

maintenance programs. 

The TSB has reviewed TCís response and assessed it as Satisfactory Intent. 

Further, on 27 March 2006, the Board recommended that: 

The European Aviation Safety Agency, in coordination with other involved 

regulatory authorities and industry, urgently develop and implement an 

inspection program that will allow early and consistent detection of 

damage to the rudder assembly of aircraft equipped with part number 

A55471500 series rudders. (A06-06, issued March 2006) 

On 22 November 2006, the European Aviation Safety Agency (EASA) stated that it agreed with 

Board Recommendation A06-06 and that Airworthiness Directive 2006-0066 issued on 24 March 

2006 requiring a mandatory one-time inspection satisfactorily addressed the Board 

recommendation. 

Although the EASA agreed with the Board recommendation, Airworthiness Directive 2006-0066 

referenced in its 22 November 2006 response does not provide for a repetitive inspection cycle 

that will allow early and consistent detection of damage, as is implied in the core of 

Recommendation A06-06. Nevertheless, the TSB assessed that the EASA is well positioned to 

take a leadership role within the industry in advocating for the development and integration of 

an inspection program dealing with composite materials. On that basis, a conference call was 

initiated on 20 December 2006. 

Following the conference call, the EASA released a further response dated 17 January 2007. This 

response stated that all elements that may have potentially caused the damage growth were still 

being investigated. Furthermore, the EASA stated that, within the Continued Airworthiness 

process and in cooperation with Airbus, it continues its efforts to determine the most 

appropriate corrective actions. Subsequently, the EASA will consider mandating those actions, 

including amending the maintenance program to require repetitive inspections. 

The 17 January 2007 response reflects EASAís commitment to continue to develop corrective 

actions that may include amending the maintenance program to require repetitive checks. 

Because EASAís most recent response contains a proposed action that, if implemented, will 

reduce or eliminate the risks associated with this deficiency, the response to 

Recommendation A06-06 is assessed as Satisfactory Intent.

4.1.1.2 TSB Safety AdvisoryóCockpit Voice Recorder Duration 

The cockpit voice recorder (CVR) installed on Air Transat Flight 961 employed a 

continuous-loop magnetic tape of 30-minute duration. The event of the rudder separation on 

the Air Transat Flight 961 CVR was recorded approximately 60 minutes before landing. Crew 

conversations and cockpit sounds before the event of the CVR recording may have provided 

substantial insight into any initiating or precursor events that led to the accident. Given the 
SAFETY ACTION

TRANSPORTATION SAFETY BOARD 67 

need for longer periods of recorded sound to capture the initiating events of aviation accidents 

and the availability of two-hour CVRs, the Board believes that such recorders should be 

mandated by regulatory authorities worldwide. 

Consequently, the TSB issued, on 03 March 2006, a Safety Advisory to TC re-addressing its 

concern that, in 2005, there are still commercial aircraft not equipped with a CVR with at least 

two-hour recording capacity. 

4.1.1.3 TSB Safety AdvisoryóDigital Flight Data Recorder Recording of Filtered Data 

With filtering, the ability to differentiate between a rudder excursion and a data filtering 

artefact is limited. The filtering of raw sensor data necessitated additional analysis to estimate 

the probable rudder position history, ultimately affecting the accuracy and timeliness of the 

investigation efforts. The Canadian Aviation Regulations (CARs) do not address the requirement 

to test parameter accuracy under both static and dynamic conditions as does 14 CFR (Code of 

Federal Regulations) of the United States. The CARs continue to refer to the previous minimum 

operational performance specifications (MOPS) for flight recorders (ED55), rather than the 

current ED112, which offers guidelines on data filtering. The current Federal Aviation 

Administration (FAA) Notice of Proposed Rule Making (issued 28 February 2005) regarding 

revision of digital flight data recorder (DFDR) regulations does not address the recurring 

problem of filtered data. 

Consequently, the TSB issued, on 03 March 2006, a Safety Advisory to TC addressing its 

concern that data filtering may prevent investigators from determining accurate control surface 

positions from recorded data, particularly under dynamic conditions. 

4.1.1.4 TSB Safety AdvisoryóDigital Flight Data Recorder Low Recording Rates 

The Air Transat occurrence demonstrated that further improvements to DFDRs are needed to 

more effectively determine the sequence of events in an accurate and timely manner. 

Specifically, due to the low recording rates for acceleration data, the existence of aeroelastic 

effects as a possible failure mode could not be positively identified. The limited lateral 

acceleration data also prevented the characterization of the initiating event. 

Consequently, the TSB issued, on 08 March 2006, a Safety Advisory to TC addressing the 

possible conduct of a review of recording rates of DFDR data to ensure that adequate 

information is made available to analyze dynamic flight events. 

4.1.1.5 TSB Safety AdvisoryóDutch Roll Recovery Procedure 

For this loss of rudder occurrence, the absence of sufficient guidance in Dutch roll recovery 

resulted in a situation wherein the crew engaged the autopilot, which led to a worsening of the 

flight characteristics. Although the engagement of the autopilot did not increase the severity of 

consequences for this occurrence, under other circumstances, such action might have led to an 

aircraft upset. 
SAFETY ACTION

68 TRANSPORTATION SAFETY BOARD

Consequently, the TSB issued, on 08 March 2006, a Safety Advisory to TC suggesting that TC, in 

concert with industry, FAA, DGAC, and EASA, may wish to conduct a review of the adequacy 

of published procedures to ensure that pilots have the required knowledge to safely recover 

from a Dutch roll situation. 

4.1.2 National Transportation Safety Board 

As a result of its investigation into an Airbus A300-600 aircraft operated by FedEx Express that 

was damaged during routine maintenance on 27 November 2005, the National Transportation 

Safety Board (NTSB) recommended on 24 March 2006 that the FAA 

Require that all operators of Airbus A-300 series airplanes immediately 

[possibly before further flight] comply with four Airbus All Operators 

Telexes (AOT) A300-55A6042, A310-55A2043, A330-55A3036, and 

A340-55A403 dated March 2, 2006. Any disbonding to the rudder skins that 

occurs in the presence of hydraulic fluid contamination should be repaired 

or the rudder should be replaced as soon as possible, well before the 

2,500 flights specified in the AOTs. (A-06-27, issued March 2006) 

The NTSB further recommended that the FAA 

Establish a repetitive inspection interval for Airbus premodification 8827 

rudders until a terminating action is developed. The interval should be well 

below 2,500 flights. (A-06-28, issued March 2006) 

4.1.3 Bureau díEnquÍtes et díAnalyses pour la SÈcuritÈ de líAviation Civile 

On 10 March 2006, the Bureau díEnquÍtes et díAnalyses pour la SÈcuritÈ de líAviation Civile 

recommended that [translation] the EASA impose as soon as possible an appropriate inspection 

program for the concerned rudders (part number A55471500). (000153/BEA/D, issued 

March 2006) 

4.1.4 Airbus 

4.1.4.1 All Operators Telex (AOT-1) 

Based on the initial information uncovered during this TSB investigation, Airbus, on 17 March 

2005, issued an AOT for the inspection of all aircraft equipped with part number A55471500 

series rudders. This one-time visual and tap-test inspection involved 222 Airbus A310s, 

146 Airbus A300-600s, 6 Airbus A330s, and 34 Airbus A340s, for a total of 408 aircraft. In 

addition, a more detailed inspection of rudder side panels on over 20 aircraft was conducted 

using the elasticity laminate checker (ELCH) test method. Finally, the attention drawn to 

rudders by the occurrence resulted in operators examining their rudders more closely during 

maintenance. These various inspections found examples of disbonds, damage around hoisting 

points and trailing edge fasteners of the rudder, corrosion and abrasion at hinges, seized hinges, 

hinges with excessive free play, and water ingress. 
SAFETY ACTION

TRANSPORTATION SAFETY BOARD 69 

4.1.4.2 All Operators Telex (AOT-2) 

On 02 March 2006, the aircraft manufacturer issued a second series of AOTs for the inspection 

of all aircraft with rudder part number A55471500 series rudders. These AOTs were 

subsequently made mandatory by an Airworthiness Directive issued by the EASA, representing 

the State of manufacture. The aim was to verify the structural integrity of the rudder by means 

of one-time tap-test inspection to the interior face sheets of the rudder side panels, as well as 

checking the drainage holes at the bottom of the rudder and cleaning hydraulic fluid from the 

external surfaces. Access to the interior of the rudder was through the inspection holes in the 

front spar of the rudder. It is noted that the access to inboard surfaces from the inspection holes 

is limited. No disbonds have been found. 

4.1.5 Air Transat 

4.1.5.1 Abnormal Situation 

Based on the initial information uncovered during this investigation, Air Transat issued, on 

10 November 2006, new procedures for situations that are not typical. The following text was 

added to the Cabin Attendant Operation Manual: 

[...] 

When experiencing an abnormal situation, you must report to the Flight 

Director as soon as possible and describe the event such as, but not limited 

to: 

• Unusual odour 

• Unusual noise 

• Presence of an unusual object 

• Instability of objects, passenger or crew 

• Inappropriate passenger behavious 

• Medical situation 

Never assume that others experienced the same effects that you have or are 

fully aware of the events taking place. Communication is the key for a 

proper evaluation of the situation and appropriate corrective measures. If 

the Flight Director can not be reached and time is critical, Cabin Attendants 

must contact the Flight Deck Crew without delay. 

4.1.5.2 Preservation of Recorder Information 

Air Transat issued, on 16 May 2006, a new Accident/Incident Response Checklist as part of the 

Flight Crew Operating Manual for the A-310. A procedure to disable the appropriate circuit 

breakers to preserve the recorded data for both the CVR and the DFDR is described. This new 

standard operating procedure also highlights that preserving the recorded information is 

critical to the investigative process after an occurrence. 
SAFETY ACTION

70 TRANSPORTATION SAFETY BOARD

This report concludes the Transportation Safety Boardís investigation into this occurrence. Consequently, 

the Board authorized the release of this report on 21 June 2007. 

Visit the Transportation Safety Boardís Web site (www.tsb.gc.ca) for information about the 

Transportation Safety Board and its products and services. There you will also find links to other safety 

organizations and related sites. 
APPENDICES

TRANSPORTATION SAFETY BOARD 71 

Appendix A ñ Direct Access Recorder/Digital Flight Data 

Recorder Data Comparison 

APPENDICES

TRANSPORTATION SAFETY BOARD 73 

Appendix B ñ Glossary 

AFD assistant flight director 

AFRP aramid fibreñreinforced plastic 

agl above ground level 

AIM Aeronautical Information Manual

AMM aircraft maintenance manual 

AMO approved maintenance organization 

AOC air operating certificate 

AOT all operators telex 

APYA autopilot yaw actuator 

ARINC aeronautical radio incorporated 

asl above sea level 

ASTM American Society for Testing and Materials 

ATC air traffic control 

ATPL airline transport pilot licence 

CARs Canadian Aviation Regulations

CFRP carbon fibreñreinforced plastic 

CG centre of gravity 

cm centimetre(s) 

cm2 square centimetre(s) 

CVR cockpit voice recorder 

daN decanewton(s) 

DAR direct access recorder 

DCB double cantilever beam 

DFDR digital flight data recorder 

DGAC Direction GÈnÈrale de líAviation Civile (France) 

DVI detailed visual inspection 

EASA European Aviation Safety Agency 

ECAM electronic centralized aircraft monitor 

EDX energy dispersion X-ray spectroscopy 

ELCH elasticity laminate checker 

EUROCAE European Organisation for Civil Aviation Equipment 

FA flight attendant 

FAA Federal Aviation Administration 

FCC flight control computer 

FD flight director 

FL flight level 

FOD foreign object damage 

g load factor 

GFRP glass fibreñreinforced plastic 

GVI general visual inspection 

GVI (G) general visual inspection from the ground of empennage 

GVI (A) general visual inspection at armís length of empennage 

GVT ground vibration test 

HIRF high-intensity radiated fields 

Hz hertz 

IR infrared 
APPENDICES

 74 TRANSPORTATION SAFETY BOARD 

KCAS knot(s) calibrated airspeed 

KFLL Fort Lauderdale/Hollywood International Airport 

kg kilogram(s) 

KIAS knot(s) indicated airspeed 

KMIA Miami International Airport 

kN kilonewton(s) 

LPP lightning protection plate 

m metre(s) 

m2 square metre(s) 

m3 cubic metre(s) 

mm millimetre(s) 

mm2 square millimetre(s) 

MOPS minimum operational performance specifications 

MRTT multi-role tanker transport 

MSN manufacturerís serial number 

MUVR Varadero/Juan Gualberto GÛmez International Airport 

N newton(s) 

NDI non-destructive inspection 

nm nautical mile(s) 

NTSB National Transportation Safety Board 

PPC pilot proficiency check 

QRH quick reference handbook 

QSMS quality safety management system 

RTL rudder travel limiter 

SB Service Bulletin 

SRM structural repair manual 

TC Transport Canada 

Tg glass transition temperature 

TSB Transportation Safety Board of Canada 

TSC921 Air Transat Flight 921 

UFDR universal flight data recorder 

UTC Coordinated Universal Time 

VTP vertical tail plane 

XPS X-ray photoelectron spectroscopy 

YD yaw damper 

° degree(s) 

°C degree(s) Celsius