Appendices. Introduction to Appendices

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1 Appendices Introduction to Appendices To assist the reader in understanding how some of the analytical tools such as dependency diagrams, fault tree analysis (FTA) and Markov analysis may be applied to typical systems, four Appendices are included. These appendices address the following systems: Appendix A. Safety Analysis Flight System Appendix B. Safety Analysis Electronic Flight Instrument System Appendix C. Safety Analysis Electrical System Appendix D. Safety Analysis Engine System The analyses in the Appendices are presented in a simple mathematical fashion to provide the reader with purely advisory and illustrative material. The failure rate probabilities offered are for the purposes of illustrating the analysis methods and to draw simple conclusions; they are not representative of any real equipment or technologies. The analyses should not be considered as definitive of the standard that would be demanded during formal aircraft system design. Nevertheless, it is hoped that they will aid the reader in appreciating some of the design issues that need to be considered early on in the design process. Refer to Chapter 4 for further relevant sections. During formal design, engineers utilise professional, qualified design tools to undertake the appropriate analysis in a rigorous fashion. At the same time, these tools provide the required documentation to the standard necessary to convince the certification authorities that the design is safe. Civil Avionics Systems, Second Edition. Ian Moir, Allan Seabridge and Malcolm Jukes John Wiley & Sons, Ltd. Published 2013 by John Wiley & Sons, Ltd.

2 Appendix A Safety Analysis Flight System This example evaluates the catastrophic failure case of total loss of the pitch axis of a fly-bywire flight control system. It is a much simplified analysis, but serves to illustrate the principles and differences between the dependency diagram approach and the fault tree approach. A.1 Flight System Architecture The flight control system architecture is shown in Figure A.1. The primary means of controlling the aircraft in the pitch axis is by means of two elevators attached to either side of the rear of the tailplane horizontal stabiliser (THS). Each elevator section is operated by two hydraulic actuators powered from one of the three aircraft centralised hydraulic systems: yellow (Y), green (G) and blue (B), as indicated. Normally one actuator on each side is the controlling actuator, while the other is in damping mode, but for the purposes of this analysis we shall assume they are all the same, and that the aircraft is still controllable (albeit with some performance limitations) provided just one actuator is operational. We shall assume the actuators fail into a neutral position. Elevator actuator demand is digitally signalled from four flight control computers. In normal mode the two elevator/aileron computers (ELACs) implement the control functions for the primary flight control surfaces. Should these computers fail, then reduced functionality is available from the spoiler/elevator computers (SECs) which normally control the secondary flight control surfaces. Each of these computers utilises a dual command:monitor architecture, but for the purposes of this analysis we shall assume that fly-by-wire demands in the pitch axis are available provided one of the four computers is operational. In manual flight, the captain or first officer input their pitch commands through a side-stick controller, one per crew member. In autoflight, pitch commands are sourced from the autopilot. Envelope protection requires the provision of aircraft attitude and air data. For the purposes of this simplified analysis we shall assume the aircraft is being flown manually within its permitted flight envelope. Civil Avionics Systems, Second Edition. Ian Moir, Allan Seabridge and Malcolm Jukes John Wiley & Sons, Ltd. Published 2013 by John Wiley & Sons, Ltd.

3 Appendix A: Safety Analysis Flight System 535 Trim Wheel Mechanical Linkage THS Motors G Y THS B Y Elevators s Autopilot IRS ADC Primary Flight Computers ELAC Autotrim Normal Alternate Normal G B SEC SEC Autotrim Alternate Secondary Flight Computers Figure A.1 Fly-by-wire flight control system pitch axis The aircraft is trimmed in the pitch axis by slow long-term movements of the tailplane horizontal stabiliser (THS). In normal operation this function is commanded automatically by the flight control computers via two hydraulic motors and a ball screw drive. Total loss of these computers will leave the aircraft in a trim condition. Emergency backup in the pitch axis is provided by a mechanical linkage from the trim wheel located in the centre console between the two crew members. A.2 Dependency Diagram Figure A.2 provides the simplified dependency diagram for the architecture described. Considering first the primary flight control system: Pitch axis control is lost if both left-hand right-hand elevators fail indicated by the parallel nature of the branches. Each elevator is lost if both of its actuator channels are unavailable. An actuator channel is lost if either the actuator fails the hydraulic system that powers it fails the demand signal is lost indicated by the series nature of the each branch. The actuator arrangement as shown is quadruplex, but there is a problem in that the aircraft has only three hydraulic systems. The blue hydraulic system is used on both elevators. Loss of the blue hydraulics system is a common mode failure. It is not easy to model this in a

4 536 Civil Avionics Systems Side Stick Captain Side Stick F/Officer 4.0x x10-16 ELAC 1.0x10-4 ELAC 2 SEC1 SEC2 ELAC = Elevator/Aileron Computer SEC = Spoiler/Elevator Computer 10.x10-8 Blue Green Yellow Blue 9.1x x10-4 Note: Ignore Blue actuator for Elevator common mode faults Left Hand Elevator Right Hand Elevator Primary Pitch All failure rate probabilities are per flight hour 9.5x x10-15 Total Pitch TrimWheel Back-Up Motor 3.0x10-4 Motor Ball Screw 1.5x10-5 THS 1.2x x10-8 Figure A.2 Fly-by-wire flight control system dependency diagram (simplified) dependency diagram, so the simple approximation is to ignore the second blue channel and the treat the actuator channels as triplex. This will yield a pessimistic, but safe result. Demands to the actuators are computed by the four flight control computers. Crew input commands are sourced from two side-stick controllers (captain and first officer). For the primary flight control function, if we assume the probability of failures per hour are as follows (note: these figures are for illustrative purposes only and do not represent the failure probabilities of actual equipment), then: Side-stick controller: per flight hour (MTBF 500,000 hrs): duplex = per flight hour. Flight control computer: per flight hour (MTBF 10,000 hrs): quadruplex = per flight hour. Elevator actuator: signal transmission: per flight hour plus hydraulic system: per flight hour (MTBF 5000 hrs) plus hydraulic actuator: per flight hour (MTBF 4000 hrs) So single actuator lane = per flight hour: triplex = per flight hour. Thus the probability of loss of the primary flight control system is per flight hour.

5 Appendix A: Safety Analysis Flight System 537 For the trim function, if we assume the probability of failures per hour are as follows: Trim wheel: per flight hour (MTBF 500,000 hrs). motor: per flight hour [ /hr] (MTBF 3333 hrs): duplex = per flight hour. Ball screw: per flight hour (MTBF 100,000 hrs). Thus the probability of loss of the mechanical backup trim function is per flight hour. Although we could compute the catastrophic loss of flight control in the pitch axis as per flight hour, it is somewhat unrealistic as the aircraft would never be flown by the emergency backup system for anything other than a dire emergency, and then only to make an emergency landing at the nearest available airfield. The duty cycle of the emergency system is somewhat uncertain and a judgement would need to be made as to what figure to use. A.3 Fault Tree Analysis Figure A.3 provides the simplified fault tree for the architecture described. The fault tree approach facilitates a more complete model of the system architecture in respect of the 2.5x x10-8 LHI G Databus 1 2.5x10-4 LHO RHO 6.3x10-8 B Databus 2 Databus 3 1.0x x10-16 Elevator 4.1x10-11 All failure rate probabilities are per flight hour 1.0x10-4 ELAC 1 ELAC 2 SEC 1 SEC 2 1.0x10-16 RHI Y Databus 4 Sidestick Captain Sidestick F/ Officer 4.0x10-12 Primary Pitch 4.5x10-11 Total Loss of Pitch Trim Wheel 3.0x10-4 Hyd Motor G Hyd Motor Y 1.0x x10-8 Ball Screw 1.2x10-5 Mechanical Back-Up 5.4x10-16 Figure A.3 Fly-by-wire flight control system fault free diagram (simplified)

6 538 Civil Avionics Systems actuator power sources and the dual use of the blue hydraulic system by actuators driving both the left-hand and the right-hand elevators. In Figure A.3 the elevators are modelled as: Two inner sections (left hand and right hand), either of which is lost if either the associated actuator fails the associated hydraulic system fails the associated data bus signal fails: = per flight hour each; and Two outer sections, both of which are lost if both actuators fail both data bus signals fail the associated hydraulic system fails: = per flight hour. Using the same probabilities as before: The probability of loss of all four elevator sections (left-hand inner right-hand inner both outer sections) becomes per flight hour (c.f for the dependency diagram method); The probability of loss of the primary flight controls system (both controllers all four flight control computers the elevators) becomes per flight hour (c.f for the dependency diagram method). The difference between the two methods reflects the improved fidelity of the fault tree analysis, but it is not that significant in the overall scheme of things. It is the order of magnitude that is important when assessing the architecture against the safety objective, in this case to be less than per flight hour for a catastrophic event. The probability of loss of the primary flight control system for a three-hour flight time is obtained by factoring the probability of failure of each component by the time at risk. Thus, for example, the probability of failure of all four flight control computers during a three-hour flight time is ( per flight hour) 4 = per flight. The probability of loss of the primary flight control system for a three-hour time of risk becomes (fault tree method) which indicates that a more robust architecture may be needed for long haul and ETOPS flight operations, as is indeed the case in the flight control system architectures for the Airbus A340, the A380 and the Boeing 787.

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