Aerospace Actuators 2 E-Book

Jean-Charles Mare

• Herausgeber: John Wiley & Sons
• Kategorie: Fachliteratur
• Sprache: Englisch

Seitenzahl: 362

Veröffentlichungsjahr: 2017

Beschreibung

This book is the second in a series of volumes which cover the topic of aerospace actuators following a systems-based approach. This second volume brings an original, functional and architectural vision to more electric aerospace actuators. The aspects of signal (Signal-by-Wire) and power (Power-by-Wire) are treated from the point of view of needs, their evolution throughout history, and operational solutions that are in service or in development. This volume is based on an extensive bibliography, numerous supporting examples and orders of magnitude which refer to flight controls and landing gear for various aircraft (fixed or rotorwing, launchers) in commercial, private and military applications. The topics covered in this set of books constitute a significant source of information for individuals and engineers from a variety of disciplines, seeking to learn more about aerospace actuation systems and components.

Leseprobe

Table of Contents

Cover

Title

Copyright

Introduction

I.1. Requirements in terms of the actuation for piloting an aircraft

I.2. Functions and architecting

1 Electrically Signaled Actuators (Signal-by-Wire)

1.1. Evolution towards SbW through the example of the flight controls

1.2. Incremental evolution from all mechanical to all electrical

1.3. Challenges associated with electrical signaling

1.4. The example of landing gears

2 Signal-by-Wire Architectures and Communication

2.1. Architectures

2.2. Data transmission

2.3. Evolutions in data transmission

3 Power-by-Wire

3.1. Disadvantages of hydraulic power transmission

3.2. Electrical power versus hydraulic power

3.3. Improving hydraulically supplied solutions

3.4. Concepts combining hydraulics and electrics

3.5. All electric actuation (hydraulic-less)

4 Electric Power Transmission and Control

4.1. Electric power transportation to PbW actuators

4.2. Electric motors

4.3. Power conversion, control and management

4.4. Induced, undergone or exploited effects

4.5. Integration

5 Electro-hydrostatic Actuators

5.1. Historical background and maturing of EHAs

5.2. EHA in service and feedback

5.3. EHA specificities

6 Electro-mechanical Actuators

6.1. Development and operation of electromechanical actuators

6.2. Specificities of EMAs

Bibliography

Notations and Acronyms

Index

End User License Agreement

List of Tables

Introduction

Table I.1.

Examples for specifying maximum control force

1 Electrically Signaled Actuators (Signal-by-Wire)

Table 1.1.

Evolution towards an all-electric information chain for the flight controls of military aircraft

Table 1.2.

Evolution towards an all-electric information chain for the flight controls of European commercial aircraft

Table 1.3.

History of the development of electric flight controls for helicopters and compound helicopters

Table 1.4.

Examples of FbW industrial helicopter and compound helicopter programs

2 Signal-by-Wire Architectures and Communication

Table 2.1.

Main digital data transmission standards used for actuation

Table 2.2.

Comparison between optical fiber and shielded twisted pair wiring

Table 2.3.

Some striking examples of SbL to the flight controls

3 Power-by-Wire

Table 3.1.

Comparison of hydraulic and electrical power technology

Table 3.2.

More or fully electrical actuation on the Airbus A380 and Boeing B787

4 Electric Power Transmission and Control

Table 4.1.

Basic relations for power transmission in alternating form

Table 4.2.

Example of electrical wire characteristics

Table 4.3.

Orders of magnitude of the dynamics present in the electromechanical power chain of a PbW actuator

Table 4.4.

Orders of magnitude of the characteristics of an IGBT (1,200 V, 75 A) with its free-wheel diode

Table 4.5.

Orders of magnitude of the failure rate of a three-phase PbW actuator, according to [CAO 12]

5 Electro-hydrostatic Actuators

Table 5.1.

Demonstrators of the IAP actuator concept

Table 5.2.

Data on the development of EHA-FP for flight controls in the United States

Table 5.3.

Data on the development of EHA-FP in Europe

Table 5.4.

Extension/retraction EHA-FD [TAK 08]

Table 5.5.

In-service aircraft using EHA or EBHA actuators for flight controls (the numbers indicate the type of HSA/EHA/EBHA actuator)

6 Electro-mechanical Actuators

Table 6.1.

Examples of EMA development for space applications in the United States

Table 6.2.

Characteristics of EMAs of the Vega European launcher

Table 6.3.

Characteristics of EMAs in EMAS and EPAD projects

Table 6.3.

Characteristics of EMA for IBC “Project zero”, according to [GIA 14]

List of Illustrations

Introduction

Figure I.1.

Example of a block-diagram representation of an actuator according to a signal vision

Figure I.2.

Example of block-diagram representation of an electro-mechanical actuator according to a power vision. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure I.3.

The different feedback loops for an aircraft

Figure I.4.

Links and control functions to be performed electrically on a flight control actuator of a Northrop B2 Bomber

Figure I.5.

Electrical, mechanical and hydraulic interfaces of a rudder actuator for a Boeing B777 (according to [SHA 15]). For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure I.6.

Joint representation (signal/power) of hydro-mechanical actuation. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

1 Electrically Signaled Actuators (Signal-by-Wire)

Figure 1.1.

Major developments of flight control signaling

Figure 1.2.

Roll control for the McDonnell Douglas F-15 fighter (http://www.f15sim.com). For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 1.3.

Roll control for the Airbus A310 [VAN 02]. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 1.4.

Flight controls of the Sikorsky S-76 helicopter (©I. Sikorsky historical archives). For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 1.5.

Summation principles of mechanical transmission. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 1.6.

Flight control cables of the Super Guppy cargo plane

Figure 1.7.

Generic architecture, signal and power, for a conventional flight control axis. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 1.8.

Auxiliary actuator SCAS group of the Eurocopter Tiger helicopter

Figure 1.9.

Analog control diagram of the CAS servovalves for the pitch control of the McDonnel Douglas fighter F-15, (http://www.f15sim.com). For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 1.10.

Dual-parallel hydromechanical aileron actuator for the Falcon 900

Figure 1.11.

Generic architecture, signal and power, of an electrically signaled flight control axis. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 1.12.

Full FbW actuator for the rudder control of the Boeing B777

Figure 1.13.

Main rotor actuator of the NH90 helicopter, the pseudo FbW version used in the flight tests for development

Figure 1.14.

Dual input servovalve for the SCAS of the Eurocopter EC225 helicopter

Figure 1.15.

Purely hydromechanical feedback loop actuator for the FbW TVC actuator of the main engine in the NASA Space Shuttle (© Moog Inc.)

Figure 1.16.

Functioning principle of LVDTs and ratiometric demodulation

Figure 1.17.

Artificial feel actuator for the Concorde

Figure 1.18.

Side-stick Left: passive (A320, courtesy of SwissTeknik LLC); Right: active (KAI T-50)

Figure 1.19.

Signal architecture with ACE in the Boeing B777. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 1.20.

Evolution of yaw control from hydromechanical actuation (Airbus A340-200, left) to electrohydraulic actuation (Airbus A340–600, right). For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 1.21.

COM/MON architecture for controlling the actuators (according to [TRA 06])

Figure 1.22.

Evolution of braking. Left: “conventional” brake with antiskid and autobrake (from [CRA 01]); right: fully integrated electrohydraulic brake. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

2 Signal-by-Wire Architectures and Communication

Figure 2.1.

Data transmission

Figure 2.2.

Federated architecture

Figure 2.3.

Integrated modular architecture

Figure 2.4.

Structure of an IMA module

Figure 2.5.

Data transmission topologies

Figure 2.6.

Example topology allowed by the ARINC 429 standard

Figure 2.7.

Example of multi-point topology by the MIL-STD-1553B redundant bus. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 2.8.

Triplex architecture to the standard AS-5643/IEEE-1694b, representative of the implementation adopted by the F35 Lightening (from [BAI 07]). For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 2.9.

Topology and redundancy for the AFDX bus. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 2.10.

Multi-point or star topologies for standard Time Triggered Protocol. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

3 Power-by-Wire

Figure 3.1.

An example of conventional architecture for centralized 3H generation as per the Airbus A330. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 3.2.

Evolution of actuation (top to bottom, from all hydraulic to all electric). For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 3.3.

Type 2H–2E power architecture for actuation functions on the Airbus A380 [MAR 04]. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 3.4.

Power architecture for actuation functions on the Boeing B787. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 3.5.

From valve control to displacement control for slats actuation. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 3.6.

Local electro-hydraulic generation [DEL 04] (see also Figure 7.8 in Volume 1). For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 3.7.

Simplified architecture of an EHA-VD. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 3.8.

Prototype of a duplex EHA-VD aileron actuator for Lockheed C-141

Figure 3.9.

Simplified architecture of an EHA-FD. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 3.10.

EHA-FD actuator for the Airbus A400M

Figure 3.11.

Simplified power architecture of an EBHA. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 3.12.

EBHA spoiler for the Airbus A380, according to [BIE 04]

Figure 3.13.

Gear drive EMA for the P80 first-stage nozzle orientation of the VEGA launcher [DÉE 07]

4 Electric Power Transmission and Control

Figure 4.1.

Notations and conventions for three-phase electrical circuits

Figure 4.2.

Corona effect in an electric motor winding [COU 16]. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 4.3.

Elementary electric machine (left) and the principle of current commutator (right). For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 4.4.

Various electric motor concepts, according to [CAO 12]

Figure 4.5.

Various power electronic converters. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 4.6.

Principle of the triangular carrier PWM

Figure 4.7.

Characteristics of a static switch. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 4.8.

Groups of static switches

Figure 4.9.

IGBT symbol and example of IGBT modules in the power drive electronics of a large civil aircraft EBHA. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 4.10.

Example of six-step control of three-phase BLDC motor (concentrated winding, six teeth and two pairs of poles) by a three-phase inverter based on three Hall effect sensors spaced at 120° electrical degrees, adapted from [JIA 14]

Figure 4.11.

Principle of vector control for PMSM. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 4.12.

Architecture of position control with field-oriented control of the motor

Figure 4.13.

Architecture of a PbW actuator. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 4.14.

Example of power characteristics of a motor/control electronics set (automotive application for 5 kW peak), according to [SON 16]

Figure 4.15.

Winding diagrams [ELR 10] and photographs [DES 12]. On the left: distributed winding, on the right: concentrated winding with single layer. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 4.16.

Redundant architectures for the MPE/motor, according to [CAO 12]. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 4.17.

Example of HUM functions integration, according to [TOD 14a]

Figure 4.18.

Power architecture and integration of an actuation system

Figure 4.19.

Example of integration of the motor control electronics: a: EBHA Gulfstream G650, b: EBHA spoiler Airbus A350, c: EMA spoiler Boeing B787, d: thrust vector control of the first stage of the VEGA launcher

Figure 4.20.

Confinement of the aileron EHA of Airbus A380 (image according to [TOD 07])

Figure 4.21.

Underwing scoops for actuator cooling (EHA of Airbus A350 aileron)

Figure 4.22.

Examples of mechanical integration of the motor control/power electronics, according to [TOD 14b] and [TOD 16]. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 4.23.

Wiring of control/power electronics. On the left: conventional wiring; On the right: example of bus-bar use in rudder EBHA, according to [TOD 16]. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

5 Electro-hydrostatic Actuators

Figure 5.1.

The first EH-VD actuators: example of elevons of the Avro Vulcan, on the left: Vulcan B.mk.1 (EIS 1957), on the right Vulcan B.mk.2 (EIS 1960), (images by courtesy of aviationancestry.co.uk)

Figure 5.2.

Survivable Stabilator Actuator Package, according to [HOO 71]. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 5.3.

Power architecture of an IAP actuator. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 5.4.

Images of EHA-FP for flight controls developed in the United States

Figure 5.5.

Images of EHA-FP developed in Europe

Figure 5.6.

EAHA actuator images based on [DOR 07]. Upper image: diagram, middle image: power capacity on opposing load, lower image: control strategy. For a color version of this figure, see www.iste.co.uk/mare/aerospace2

Figure 5.7.

LGERS system with EHA-FD within the POA program [GRE 04]. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 5.8.

Landing gear steering by EHA as part of the MELANY project, extracted from [LIE 12]

Figure 5.9.

LGERS system with EHA-FD within THERMAE II program. Upper image: power architecture [ELL 16], lower image: overall view

Figure 5.10.

Dual-tandem flaperon EHA actuator on Lockheed F-35

Figure 5.11.

Simulated temperature field and thermography image of a prototype EHA operating at null speed and 140 bar, according to [TAK 04]. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 5.12.

Ground-measured temperatures for spoiler EBHA of the Airbus A380, according to [BIE 04]. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 5.13.

Fluid circulation in a wet rotor/dry stator EHA-FD actuator in the phase of maintaining force on a load at null speed (drawing on the left extracted from [DOR 07]). For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 5.14.

Principles adopted for EHAs with dissymmetric actuator, according to [VAL 14]. Upper image: addition of a piloted check valve, lower image: axial piston pump with 3-port valve plate

6 Electro-mechanical Actuators

Figure 6.1.

Examples of EMAs for space applications

Figure 6.2.

Mechanical architectures of EMAs, Upper image: ELAC EMA, Sabca (according to [MON 96], Lower image: EPAD EMA [KOP 01]). For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 6.3.

Examples of EMA demonstrators for primary flight controls of aircraft

Figure 6.4.

Examples of EMA for aircraft secondary flight controls

Figure 6.5.

In service EMA for PbW brake

Figure 6.6.

Simplified architecture of the braking system of the Boeing B787

Figure 6.7.

Examples of EMA extension/retraction demonstrators

Figure 6.8.

Examples of EMA demonstrators for landing gear extension/retraction

Figure 6.9.

EMA for helicopter or convertiplane flight control. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 6.10.

Simplified power architecture of ETRAS of the Airbus A380. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 6.11.

Redundant actuation topologies. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 6.12.

Direct drive EMA and gear drive EMA. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 6.13.

Overall mechanical stiffness of an EMA prototype (50 kN/100 mm/s), according to [KAR 09]. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 6.14.

Motor/nut–screw integration in an EMA. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 6.15.

Nut–screw layout in an EMA, according to [KAR 06]. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 6.16.

Realization of the anti-rotation function in an EMA

Figure 6.17.

Possibilities for using the plane epicyclical gear. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 6.18.

Mechanical end-stop by dog-teeth device. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 6.19.

Typical characteristic of a short-circuited permanent magnet motor, according to [ROT 14]. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 6.20.

Overall efficiency of a roller screw EMA (diameter 50 mm, lead 3 mm), according to [KAR 09]

Figure 6.21.

Jamming tests of a roller screw [TOD 12b]

Figure 6.22.

Pyrotechnic declutch [NAU 13]

Figure 6.23.

Declutching by axial thrust bearing release, according to [MAR 11]

Figure 6.24.

Declutching by separation of the translational element of the nut–screw and the load, upper image: [JIM 12, JIM 16], middle image: [NAU 16], lower image: [BIL 14]. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 6.25.

Jam-tolerant nut–screw, according to [CHE 10]. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 6.26.

Breaking-resistant EMA. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 6.27.

Thermography image of an EMA [GRA 04]. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

Figure 6.28.

Dynamic force feedback, from [DÉE 07]. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

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Series EditorJean-Paul Bourrières

Aerospace Actuators 2

Signal-by-Wire and Power-by-Wire

Jean-Charles Maré

First published 2017 in Great Britain and the United States by ISTE Ltd and John Wiley & Sons, Inc.

Apart from any fair dealing for the purposes of research or private study, or criticism or review, as permitted under the Copyright, Designs and Patents Act 1988, this publication may only be reproduced, stored or transmitted, in any form or by any means, with the prior permission in writing of the publishers, or in the case of reprographic reproduction in accordance with the terms and licenses issued by the CLA. Enquiries concerning reproduction outside these terms should be sent to the publishers at the undermentioned address:

ISTE Ltd27-37 St George’s RoadLondon SW19 4EUUK

www.iste.co.uk

John Wiley & Sons, Inc.111 River StreetHoboken, NJ 07030USA

www.wiley.com

© ISTE Ltd 2017

The rights of Jean-Charles Maré to be identified as the author of this work have been asserted by him in accordance with the Copyright, Designs and Patents Act 1988.

Library of Congress Control Number: 2017930553

British Library Cataloguing-in-Publication Data

A CIP record for this book is available from the British Library

ISBN 978-1-84821-942-7

Introduction

This book is the second volume in a series dedicated to aircraft actuators. The first volume, Aerospace Actuators 1, focuses on the actuation needs in the aerospace industry, specifically on the reliability and on the hydraulically supplied actuators. This second book, Aerospace Actuators 2, is the logical continuation of this. It is, in effect, about the evolution of aircraft towards more (or total) electric systems, and involves “signal” as well as “power”. The third volume of the series, Aerospace Actuators 3, is dedicated to the detailed analysis of recent achievements. It builds on the concepts and generic solutions that are presented in the first two volumes.

The first two chapters of this book relate to the processing and transmission of signals in electrical form, which will be designated by the general name Signal-by-Wire1: the actuator receives and transmits signals electrically. Chapters 3 and 4 relate to the electrical component of the actuators driven by electric power, and categorized by the designation Power-by-Wire. Chapter 5 is dedicated to Electro Hydrostatic Actuators (EHA) and Chapter 6 to Electro-Mechanical Actuators (EMA).

Like the first, this second volume focuses on the needs, the architectures (functional, conceptual and technological), the benefits and the limitations of technological solutions and the orders of magnitude. It uses mathematical models and goes into detail on technological implementation only when it is necessary for the understanding of those principles. This approach allows us to concentrate on the analysis, or synthesis, of technological concepts and their implementation. That being said, it is evident that the capacity for mathematical modeling and a detailed knowledge behind the technology, and the inevitable imperfections of said technology, are major factors in the success of developing and operating industrial products such as actuators. As always, the architectural choices and their implementation are strongly impacted by the capability to make realistic models and by the constraints induced by technology, thereby producing a looped feedback effect on the design process itself.

Throughout this book we will note that the complexity for engineers is greatly increased in the transition to more or total electric actuation, because this feedback loop effect acts in addition to the numerous and new coupled effects that occur between generic topics: mechanical (vibration, tribology, thermal), signal processing (networks, interfaces, control), electric (power electronics, electromagnetic) and other (dependability, human–machine interface), etc. This is why it is particularly important to effectively combine a top-down approach (from the needs to architectures, and then to technological solutions) and a bottom-up approach (from mature technology to architectures). This leads to a hybrid approach, or middle-out. The vision of “requirements, architectures and concepts”, is therefore complementary to the vision of the expert to build on a cross-functional approach, with a real system vision in mind. As it turns out, the vision of the expert is well documented in the scientific literature, in contrast to the visions of the designer or system supplier. This book seeks to capitalize and document this system vision applied to more or total electric actuation in aerospace. As such, focused experts will not find a high level of detail in their specialization (signal transmission, power electronics, electric machines). On the contrary, a particular effort will be to focus on popularization, to help the reader to become accustomed to conventional solutions and realize the principles and characteristics of more or total electric actuation. As with Volume 1, the following literature is recommended to accompany this comprehension:

– for more electric actuation in aerospace [RAY 93, SCH 98];

– for electrical systems of aircraft and avionics [COL 11, CRA 08, DAN 15, DUB 13, MOI 08, MOI 13, SPI 14, USF 12, WIL 08, WIL 09];

– for power electronics and electrical machines [DED 11, GIE 10, GRE 97, LAC 99, RAS 11].

NOTES:

– If we adopt a needs/solutions vision, it is interesting that more or full electrics is often presented as an objective

2

(a need) when in fact we should see it as a means (an evolution) to increase performance, reduce constraints and create entirely new services. In commercial aerospace, the final needs of the passenger or even the airline, are neatly summarized by four qualifiers: cheaper, safer, greener and faster. In recent times, this last qualifier tends to be forgotten, because in light of the present technology, an increase in speed impacts negatively on the other qualifiers.

– Although it seems inconsistent, with respect to the advancement towards more or total electric actuation, there is a great need for more research to be done in all fields of mechanical engineering (e.g. solid mechanics, materials, vibrations, tribology, thermal). And this will be a recurrent theme throughout this book.

I.1. Requirements in terms of the actuation for piloting an aircraft

The piloting of an aircraft, initially purely manual, developed actuation functions at both power and signal3 levels. This was due to some very diverse needs, listed in the following sections, and which are closely related to the aircraft type, mission and the inherent nature of the actuation itself (to do with the flight control, landing gear or engine).

I.1.1. Reducing the control forces needed for piloting

It is necessary to limit the forces to be exerted by the pilot, and make them compatible with human capabilities with a view to reduce physical fatigue. This is to ensure that the pilot can apply the necessary level of effort required in the worst piloting situations (during transient) or for longer periods (steady state conditions). Table I.1 gives some examples for the maximum levels of control forces conceivable, as defined by European standards. For a large aircraft, note the factor 7.5 to 10 between the short-term and long-term control forces. The specification of the maximum control forces needed to pilot military aircraft [MIL 80, MIL 97] is much more complex because they depend on many factors (phase of the mission, flight quality, load factor, etc.). Reducing control force mainly concerns the power aspect of actuation, which will be discussed in Chapter 3.

Table I.1.Examples for specifying maximum control force

Standard

Force applied to the control wheel or pedals (N)

CS-25.143d Large aeroplanes [EUR 15]

Pitch/Roll

Yaw

Transient, one handed

222/111

667

Transient, both hands

334/222

667

Long duration

44.5/22

89

CS-29.397 Large rotorcraft [EUR 12]

Longitudinal/Lateral

Yaw

445/298

578

I.1.2. Reducing the intellectual burden on the pilot

In order for the pilot to concentrate on the mission at hand, it is important to reduce the intellectual exertion associated with the conduct of the flight and the aircraft. This requirement therefore concerns the development of control commands:

– to stabilize the aircraft (Control Augmentation System or CAS) by rejecting the effects of various disturbances (e.g. gusts or crosswinds);

– able to

decouple, synchronize or coordinate

the different commands to act only on the desired degrees of freedom (e.g. remove the yaw which is induced by the aileron deflection; or further increase the engine throttle of a helicopter when there is an increase in the cyclic pitch);

– to

compensate

the controls to ensure the balance of the aircraft (e.g. action on the angle of attack of the horizontal stabilizer to ensure the longitudinal balance through the pitch trim);

– to

remain within the flight envelope

by monitoring margins compared to permissible limits (e.g. in terms of the load or the never-exceed speed).

I.1.3. Allowing independent and automatic piloting

Autopilot (AP) or an Automatic Flight Control System (AFCS) eliminates the need for a human pilot onboard service in conjunction with an autopilot4. This occurs:

– either to unburden long and repetitive tasks (e.g. the automatic pilot of an airliner);

– or simply to allow unmanned flight. This therefore applies to both space launch vehicles and missiles, Unmanned Aerial Vehicle (UAV), Optionally Piloted Vehicle (OPV) or Remotely Piloted Aircraft (RPA).

I.1.4. Increasing the performance of the aircraft

With human piloting, the issuing of actuator control orders is limited by the accuracy and subjectivity of a pilot’s senses, his capacity for processing information in real time (especially under heavy loads) and finally by his ability to react.

This latter limitation can be illustrated by studying the transfer functions of the pilot, that is to say, the mathematical model representing the transfer between perception and action [MIL 97, MC 74, ROS 03]. However, as the pilot “adapts” to the nature (to the transfer function) of the system under control, it is impossible to uniquely define his transfer function. The pilot can for example “correct” the dynamics of the controlled system by introducing a proportional, proportional-integral or lead-lag action. In any case, it is noted that the command applied by the pilot is tainted with a pure delay varying from 0.1 to 0.4 s typically. It is thus clear5 that it is impossible, for example, to dose a braking action without skid, hence the required bandwidth is within the range of a few Hz.

By pushing the limits of human control, the development of computer-controlled commands can significantly increase the performance of an aircraft, and at different levels:

– for the

flight envelope

the reduction of margins is permitted by the speed of monitoring and limiting reaction, timing and decoupling. This allows, for example, Limiting structural loads due to maneuvering and gusts (Manoeuver Load Alleviation or MLA and Gust Load Alleviation or GLA) or to increase passenger comfort by reducing the impact of accelerations;

– for the

stability

of the aircraft (Stability Augmentation System, or SAS) through the introduction of correctors in the flight control laws that prevent, for example, aerodynamic coupling, like the Dutch roll;

– for the

flight qualities

in order to improve aerodynamic efficiency, for example, by lowering the ailerons (aileron droop) when the flaps are deployed in the landing phase;

– for the

dynamics

regarding the generation of flight control setpoints, for example, hyper-maneuverability, provided that the actuators have sufficient bandwidth.

I.1.5. Facilitating development and integration

The use of computers to develop or process orders issued from the human pilot provides flexibility in the development of an aircraft. It is, for instance, possible to change very quickly the control laws of the actuators in response to the orders of the pilot in order to assess them, or even to emulate the performance of another plane.

Note also that the transmission of information in electrical form releases many geometric integration constraints in terms of the routing for control signals within the airframe.

The list of needs is of course huge as it depends on the application. Using military aircraft as the example, we could cite many improvements on survivability6 and the simplification of operational support.

I.2. Functions and architecting

Actuation can be perceived in different ways depending on the perspective we adopt. Concerning the architecture, concepts and sizing, we draw largely on a graphic representation that allows us to highlight the processes and the quantities by which they operate. We may, for example, adopt the generic form shown in Figure 1.5 of Volume 1, or choose other more appropriate graphics to represent either the vision for signal or the vision for power. An intermediate representation that describes both aspects of signal and power is often adopted as a compromise.

I.2.1. Signal vision

If we take the point of view of the control, it more often than not shows the actuation in the form of block diagram, like the one in Figure I.1 below.

Figure I.1.Example of a block-diagram representation of an actuator according to a signal vision

The values circulating on the pathways connecting the blocks are considered as pure signals. The controlled variable is generally a position (e.g. flight controls) or a force or pressure (e.g. braking). The inputs are either functional stimuli (e.g. the order y* issued by a pilot) or disturbances d. These disturbances can be external (e.g. the measurement disturbance dM, the variation in power supply dD (or power disturbance) and the quantization noise of control signals dC). They can also result from coupling produced by the power transmission between the “blocks” (e.g. opposing force of a control surface on its actuator or flow consumed by the set of brake pistons). Thanks to the measures ym, the control develops the orders u to ensure that the output variable y is consistent with the orders y* (tracking function) and is insensitive to various disturbances d (rejection function). In general, the power source and the power transfer do not appear explicitly on this representation.

It is important to note that the boundary of an actuator is not universally defined. In terms of aircraft, the actuator is most often associated with a physical unit that the aircraft manufacturer integrates between the airframe and the driven load. It will be seen later that the development of an actuation modifies the boundary of the actuator as defined by its physical interfaces. This is illustrated in Figure I.1 by the dotted lines. For example, for hydromechanical actuators that receive a mechanical position setpoint, the “control” block is purely hydro-mechanical. This is opposite to conventional hydraulic servo-actuator for which the Electro-Hydraulic Servo Valve (EHSV) current, that is to say the control signal u, is calculated in computers7 that are often still located in the cockpit. In the actuators which are electrically signaled only (full SbW), the control and measurement signals are sent in electrical form.

I.2.2. Power vision

If we instead seek an energy vision, the graphical representation emphasizes the energy transfers between the various components (blocks in the block diagram) of the actuator, since the energy flows from sources (inputs) to the users (outputs). This is interesting as it explicitly distinguishes the visions for signal and power. For example, arrows in bold lines can be used to illustrate power flows. In Figure I.2, this distinction is reinforced using a half arrow in reference to the Bond graph [KAR 00]. We further augment readability by using different colors depending on the physical field.

Figure I.2.Example of block-diagram representation of an electro-mechanical actuator according to a power vision. For a color version of this figure, see www.iste.co.uk/mare/aerospace2.zip

The power bonds carry both power variables (see Table 1.3 of Volume 1), for example voltage U and current I, pressure P and volume flow rate Q or force F (or T) and velocity v (or ω). We may also consider these with respect to their time integral (e.g. the position in the mechanical domain). From a functional point of view, it is often necessary to define causalities on these links, that is to say, what is the variable imposed on the block involved (the cause) and what is the variable that responds to the link (the result). This distinction may be apparent, for example, the position control where the actuator’s function imposes the load’s position, and in turn produces an opposing force. However, the choice of causality to adopt the representation is not always that simple. It depends in particular on the level of detail: does an electric motor impose torque in response to the applied current or speed in response to the applied voltage? It further depends, for example, whether or not we consider the inertia of the rotor or the current feedback loop.

As the focus is on the exchange of power, it is useful in some cases to represent the heat flux t to the atmosphere. This provides a balanced power representation when performing energy audits. Similarly, it would be necessary to represent the mechanical power links associated with force flow from the structure and from the load [DAU 14], which would provide a mechanically balanced representation (these connections are not shown in Figure I.2 in the interests of clarity). The Power-by-Wire (PbW) actuators are powered by a source of electrical power. For these actuators, the need to use balanced representations in force and in power is much stronger than for the hydraulically powered actuators. Indeed, in these PbW actuators, we will see that the rejection of the heat generated by power losses, reaction forces and inertial effects all impact heavily on the architecture and overall sizing. In practice, it is unfortunately difficult to produce diagrams which completely reflect these balances (in the sense of effort, energy, etc.) while maintaining readability.

It is equally interesting to explicitly highlight the signals that act on/control the power transfer (e.g. the input current of a servovalve or the command in the form of a pulse width modulation, which is applied to an inverter).

In practice, the distinction between signal and power, however, is not unique. It depends on the system under consideration, the engineering task and the level of detail of the study. For example, the electrical control of a flight control actuator can be viewed as a signal, without power transfer. Indeed, the power required to control a servovalve (some 10 mW) is negligible when compared to the power transmitted by the actuator on a moveable surface (a few kW to tens of kW). Similarly, at the level of the flight control system, the dynamics of the servovalve (a few tens of Hz) can be neglected when compared to the bandwidth of the position control (a few Hz). By contrast, from the perspective of the computer that issues the orders to the actuator, it may be necessary to take into account the input electric power and the dynamics of the servovalve coils.

I.2.3. Basic functions in actuation

Understanding the needs and the evolution of actuation functions is facilitated by firstly considering human piloting in view of flying rules. It is quite possible to represent the general architecture in the form of a block diagram, as shown in Figure I.1. It is considered that the output is the actual attitude and trajectory of the aircraft. The functional input is the y* desired by the pilot. The disturbance input d alters the actual output (e.g. wind gusts affecting the flight controls, imperfections of the runway on the landing gear, etc.). The pilot therefore aims to achieve a tracking function (the actual output of the aircraft follows the desired input by the pilot) and a rejection function (the actual output of the aircraft is immune to disturbances acting on the aircraft).

In this representation, it is therefore possible to distinguish:

– a comparator function and generation of the command u. The pilot compares the desired attitude and trajectory

y*

of the aircraft with his perception

y

m

of the actual output of the aircraft. He creates the command

u

to apply the inceptors (control stick, pedals, etc.) in order to reduce the error between the desired change and the actual output;

– a measurement function on the actual response of the aircraft. The pilot perceives the aircraft’s attitude and trajectory through the sensors which act as extensions of his senses (sight, the feeling of acceleration, noise, etc.);

– an actuation function in the strictest sense of the meaning. In manual piloting, that is to say without any amplification of force, the muscles of the pilot carry out the function of the actuator. They develop the power required to move parts of the aircraft via the inceptors. This power is then transmitted mechanically (rods, levers, pulleys, cables, etc.) or hydrostatically for nonassisted braking (master cylinder, tubes and cylinders receptors).

Figure I.3.The different feedback loops for an aircraft

This breakdown generates several observations:

a) when talking about a closed-loop control system, there is implicitly a control function and a power amplification function. The power is provided by a source outside the considered system, although it does not explicitly appear on the graphic in

Figure I.1

.

b) in practice, there are usually several nested feedback loops, as shown in

Figure I.3

: the aircraft in air traffic control , the trajectory and attitude of the aircraft , the actuators of the aircraft and finally, the internal feedback loops to the actuator (e.g. internally from the feedback of a servovalve spring or the feedback from the motor power drive of a brushless electric motor). The more internal feedback loops, the higher the dynamics (e.g. 1,000 Hz for a current loop, for a servovalve 70 Hz, 3 Hz for an actuator, etc.).

c) the signals which are propagated between the blocks of the block diagram in

Figure I.1

are multidimensional. For instance, regarding the control interfaces, we can indeed mention the stick, rudder pedals and the elevator trim wheel. To activate movable surfaces, we can distinguish the actuators for roll, pitch, yaw, the horizontal stabilizer, etc. For sensors, there are, for example, the gyrometers, the airspeed sensors, the position sensors on movable surfaces, etc.

d) some of these “blocks” generate strong couplings between the various input and output signals. These couplings may be undesired, e.g. a yaw induced by aileron deflection on an airplane; or by increasing the collective pitch of the main rotor of a helicopter. These couplings can also be functional, e.g. at the level of a mixer for controlling the main rotor actuators of a helicopter; or a command law for the uncoupling integrated into the flight control computer.

e) for the flight controls, it will later be seen that the CAS and SAS produce other links in the internal feedback loops. The CAS augments the pilot’s commands as issued to the actuators so as to improve the maneuverability of the aircraft, for example a turn without slipping or skidding. The SAS rejects the effect of rapid aerodynamic disturbances and improves the stability of the aircraft. The AP substitutes for the pilot to generate the flight control commands to the actuators, as per the setpoint of the aircraft loop, according to the desired flight path as provided by the Flight Director (FD).

In the end, it is clear that the success of a flight mission depends on several generic actuation functions. To facilitate reading and reduce complexity, the rest of the book will distinguish these functions as either an aspect of “signal” or an aspect of “power”. However, it will also repeatedly show how these two aspects can be coupled, specifically for all the intermediate power levels between the computers and the load, e.g. concerning the power command of a Direct Drive Valve (DDV); or the supply of a solenoid valve. These aspects are illustrated by the following two practical examples.

I.2.3.1. Example 1: Flight control actuator of the Northrop B2 Bomber

Figure I.4 shows a representation, as taken from a control point of view, of one of the two actuators associated with the active–active mode for operating a flight control surface of a Northrop B2 Bomber [SCH 93].

Figure I.4.Links and control functions to be performed electrically on a flight control actuator of a Northrop B2 Bomber

In this example, it is possible to identify several functions related to the active mode of the actuator8