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FLY.BY-WIRE: FLIGHT CONTROL SYSTEMS

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8 SEPTEMBER 1968

This document has been approved for pub-L.lie release; its distributi6n is unlirited.

Prepared orJoint Meeting of Flight Mechanics and

Guidance and Control Panels of AGARDOslo, Norway C.) C .

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AD 679 i158FLY-BY-WIRE FLIGHT CONTROL SYSTEMSJ. P. Sutherland ,Air Force Flight Dynamics LaboratoryWright Patterson Air Force Base, Ohio3 September 1968

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FLY-BY-WIRE FLIGHT COINTROL SYSTEMSIntroduction

;>The purpose of this paper is to provide the reader with an intro-duction to fly-by-wire and an outline of state-of-the-art fly-by-wiretechniqui-s. An outline of the philosophy of fly-by-wire flight controlsystems is given, the evolution of fly-by-wire is discussed, the ad-vantages of fly-by-wire over mechanical systems are listed, currentfly-by-wire techniques are outlined, and a brief review of, he Airforce Flight Dynamics Laboratory proposedcin-house and contractedfly-by-wire development programs is given.'

The Philosophy of Fly-by-WireBefore discussing fly-by-wire, it is important to understand what

is meant by the term "fly-by-wire". Two other terms, "electricalprimary flight control system" and "pseudo fly-by-wire", are oftenused in discussions of fly-by-wire and therefore also require definition.The following definitions of these three terms apply throughout thispaper and have been generally accepted by the Air Force Flight DynamicsLaboratory.

Electrical Primary Flight Control System (EPFCS) - A flight controlsystem mechanization wherein the pilot's control commands are transmittedto the moment or force p/roducer only via electrical wires.

Fly-by-Wire - A fly-by-wire flight control system is an electricalprimary flight control system employing feedback such that vehiclemotion is the controlled parameter.

Pseudo Fly-by-Wire - A fly-by-wire flight control system with anormally disengaged mechanical backup.Fly-by-wire, that is, the complete replacement of the mechanical

linkages between the pilot's stick and the control surface actuatorsby electrical signal wires, offers a convenient and logical solutionto many of th e control system problems associated with modern highperformance aircraft and aerospace vehicles. However, there existsa strong reluctance on the part of both pilots and flight controlsystem designers to remove all flight control cables and mechanicallinkages and rely solely on electrical signals and electronic devices.Nor is this reluctance unreasonable. Since the Wright Brothers firstflew at Kittyhawk in 903, there has been some form of direct mechanicallinkage between the pilot and the control surfaces or control surfaceactuators. The succesaful use of such systems has resulted in hegrowth of a sense of security toward mechanical control linkages whichnow tends to inhibit fly-by-wire development. "Security is a mechanical Kflight control system" quips Snoopy as he pursues the Red Barron(Figure 1). Yet in today's high performance aircraft, security is


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definitely not a mechanical control system. Instead, security ishaving a reliable Stability Augmentation System (SAS). For withoutSAS many high performance aircraft are only marginally stable andmust, therefore, rely on electronic devices (black boxes) for thesuccessful completion of a mission. In uch cases, an effort hasbeen made to obtain some of the advantages of fly-by-wire withoutlosing the "security" of a mechanical system with th e result that '-many of the disadvantages of the mechanical systea axe retained. AThe state-of-the-art in electronic circuits and redundancy techniqueshas now antiquated this approach. It is now possible to talk realistic-ally about building a pure fly-by-wire flight control system that ismore reliable than its mechanical counterpart. Until it is actuallydone, however, and successfully demorstrated in flight tests, theMissourian in many of us will prevail and the security stigmaassociated with mechanical control systems will predominate. Ourfly-by-wire effort is orientated towards fulfilling this need.

The Evolution of Fly-by-WireThe concept of fly-by-wire is not something which sprung up overnight, but rather it evolved slowly through th e years as aircraftflight control system requirements changed. With progressive increases .in aircraft size and speed, power-boosted control quickly became arequirement in order to enable the pilot to utilize the full maneuver

capability of the aircraft. Hydraulic boost, wh'ere a hydraulic actu-ator is onnected in arallel to add to the pilot's force on thecontrol cables, is still used on many aircraft; for example, the B-47,T-33, 707 rudder, and 72 7 elevators and ailerons. Shortly afterWorld War II, ully powered controls came into being. Here the controlcables from the pilot's stick are attached direc!y to the spool ofthe servo valve on the actuator and are in no way physically connectedto the control surface Feel is introduced into the systen artificiallywith springs, dash pots, bob weights, and in ome :ases "q" ellows.This artificial feel, while not required in oving the control surfaces,is eeded to give the pilot th e proper handling quelities characteristicsfor control of the aircraft. Hence, although the pilot has no directphysical connection with the control surfaces, the artificial feel systemgives him the impression that he has. Examples of aircraft using fullypowered controls are th e F-86, F-4C, F-104, F-105, and 727 rudder. Oneof th e primary reasons for using fully powered cont-ol is ha t in hetransonic region the forces on the surfaces vary gi-atly and are highlynonlinear. The resulting stick forces with direct mechanical connectionto the control surfaces were unacceptable from a handling qualitiespoint of view. Fully powered controls are inherently irreversible andthus unaffected by nonlinearities in he transonic region, allowingthe artificial feel system to be designed to give smooth transition fromsubsonic to supersonic flight. K'i


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As aircraft continued to increase in size ant performance, itbecame necessary to add stability augmentation to assist the pilotin his control task. Stability augmentation systems (SAS), havingvery limited authority, were added in series with the normal flightcontrol system. For some aircraft in certain flight regimes, however,the proper functioning of the SAS was required for the very survivalof the aircraft. Th e success of SA S led to the introduction of CAS,Control Augmentation System. A control augmentation system has anelectrical system operating in parallel with the mechanical controlsystem. The electrical system predominates by virtue of its highgain and servo authority and performs essentially as a fly-by-wiresystem. Th e step from CAS to pseudo fly-by-wire is a small one andinvolves declutching the mechanical system when it is not in use.To get a fly-by-wire system from a pseudo fly-by-wire system, oneneeds only to remove the mechanical flight control system entirely.Fly-by-wire flight control systems are currently used in some spacevehicles. Figure 2 illustrates the SAS to CAS to FBW evolution.

Need for and Advantages of Fly-by-WireThe flight control systems of yesteryear, which consisted of

relatively simple direct mechanical linkages, cables, and feel springs,can no longer meet the demands of advanced aircraft control systemrequiretents. The flight control designer has been forced to replacethe simple manual control system witl: complex nonlinear linkages, mix-ing assemblies, power actuation devices, and active artificial feelsystems containing literally hundreds of different parts and inter-connections. In his struggle to meet rigid performance and environ-mental requirements (such as immunity to aircraft structural changesdue to flexing and thermal expansion) the designer has been confinedby the requirements for low weight and high reliability. Hence, acompromise is orced and the full potential of many aircraft is everrealized because of th e resulting control system limitations. Thedegree of complexity to which flight control designers have had to goin their effort to solve these proolems is est i'lustrated by anexamination of Figure 3 which depicts a portion of the flight control*system of a typical high performance tactical fighter aircraft. Youwill note that the system is ade up of a great number of relativelyheavy push rods, bell cranks, and other linkages with a total of onehundred and fourteen bearing points. Each bearing point represents asource of fri tion and a possible failure point. Nor is he complexityof this example flight control system illustrated in igure 3 unique.The B-70 flight control system is even more complex but would requiresuch a large foldout to display that one might say it is beyond thescope of this paper. Helicopter flight control systems are alsoenormously complex and their problems were multiplied several foldwith the introduction of V/STOL aircraft. Figure 4 illustrates whata simple nonredundant fly-by-wire system might look like, This fly-by-wire system would do the same task as that complex mechanicalsystem shown in igure 3 and do it etter. We would not suggest,4

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IFhoaever, that a simple nonredundant fly-by-wire system should everbe used in such an aircraft since its reliability would not be highenough for "the mission requircmwint. A two-fail/operate or quadruplyredundant fly-by-wire system, as shcmi in Figure 5. could meet allhe requirements of a typical high pe:rformance aircraft includingreliability. This additional redundacy can be added to the simplefly- 3 , ire system with little complexity or weight penalty and witha significant increase in reliability. Figure 6 illustrates, on ahypothetical tactical fighter aircraft, the relative simplicity ofsuc'h a system especially when used with integrated servo actuatorpackages.

Many of the advantages of a fly-by-wire flight control systemover the conventional mechanical flight control system are self-evident. Some, however, tend to become obscured by misinformation,skepticism, inertia, prejudice, or just plain ignorance. Below areisted some of the advantages of fly-by-wire with supporting factsand figures where available and applicable.a. Design and Installation Savings - The design and installationmanhour savings that can be realized by using fly-by-wire are fairlyelf-evident. Cable tension, routing, and maintenance accessibilityare only a few of the many problems which are virtually eliminated byly-by-wire design. North American Rockwell Corporation estimatedhat, based on large production quantities, approximately 5000 manhoursper aircraft could be saved on the design and installation time of theflight control system for large, high performance, strategic bombertype aircraft.

b. Weight Savings - The weight saving that can be realized byusing fly-by-wire is very significant when considered as a percentagef the flight control system weight. For example, Sperry Pheonix haveestimated that as much as a 58% reduction in the flight control systemeight, or approximately 277 lbs could be realized by using fly-by-wire on tactical fighter aircraft. General Dynamics Corporationeatimated that an 84% reduction (535 lbs) in control system weightcould be realized by using fly-by-wire on large, high performance,strategic bomber type aircraft. Lockheed Aircraft Corporation haveestimated a savings of as much as 700 lbs by using fly-by-wire onlarge transport aircraft. A very significant savings in weight canlso be achieved by using fly-by-wire on helicopters. Vertol estimatesa savings of up to 86% (718 lbt) could be realized by using fly-by-wireon helicopters similar to those in current use.c. Volume Savings - Th e volume savings realized by using fly-by-wire is particularly significant on high performance aircraft. SperryPhoenix estimated that a reduction in volume of 1469 cubic inchescould be effected by using fly-by-wire on tactical fighter aircraft.For strategic bomber type vehicles the volume reduction is even more

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EQUIVALENT FLY-BY-WIRE SYSTEMQUADRUPLY REDUNDANTKAIRCRAFTMOTION SENSORS

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Experience has shown that pilots do not like the handling qualitiesof a flight control system which has a step response to a sudden input.Consequently, an electronic filter or model is introduced prior to thesummer which gives the desired handling qualities characteristics.Normally, this model is in the form of a lag circuit.

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pThen as before for H 1 and KG 1.C WZ~1pI+ Qi

which says that the aircraft flies like the model and this, of course,is the desired goal and we now have what is referred to as a Model-Following Control System. The desirable gust response cha--acteristicshave not been altered since C/D 0 as before. For those who 2aysuspect that the above is sone form of nathe=tical wizardry which inpractice probably would not work, let ne assure you that such systesdo work as advertised. The control augmentation systems (see Figure 2)which are daily logging flight tine on the F-!!! and A-7A aircraftattest to this fact.

Typical Fly-by-Wire IplementationA sixwlified block diagra for a single aAs fly-by-wire controlsystem as shown in Figure 7. A acre conplete representation of aquadruply redundant three-axis sysrem is depicted in Figure 11 . No

effort has been nade here to show the type or location of monitors andcoaparators nor has any atteupt been m.ade to outline the contents ofeach block in the diagram. These details are cocpletely dependentupon the system a lication; and if the reader desires such inforation,it may be obtained from the referenced literature.

Sensors, Transducers and ElectronicsOne of the things which makes a two-faii/operate or quadruply

redundant scheme feasible is the current state-of-the-art of sensors,transducers, and electronics. Small, reliable, lightweight, andrelatively low cost sensors and transducers are available as off-the-shelf hardware. The penalty which rust be paid for using quadrupleredundancy is thus minimized. Sensors of the future currently underdevelopment at the Air Force Fiight _ynamicsaboratory promise toreduce this penalty even further. An example of this is the DARTsensor (Figure 12) which uses rotating mercury to sense rate of rotationand linear acceleration in two axes. A quadruply redundant sensor packagemeasuring pitch rate and normal acceleration could be contained in apackage 6" x 3" x 3" eighing less than three pounds. The electronicsfield has miniaturized even more so. Microelectronic circuits enablemixing, blending, voting, and, in general, response shaping to be dneon a relatively small number of easily replaced cards. Metal oxidesemiconductor (MOS) techniques decrease the size of these components byseveral orders of magnitude. Using these techniques it is ow possibleto redundacize electronics at a functional module-level vith a resultingdecrease in size, weight, and cost and a net increase in system reliability.

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QUARUPLY RDJNAN.FLYMBY-WIRE. FLIGHT CONTROLSYSTEMSTRUCTURAL

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IVElectrical Wiring and Connectors

The advent of fly-by-wire puts a copietely new importance andp-'ority on aircraft electrical wiring and connectors. The techniquespresently being used in ost aircraft would not be suitable for useon a system which will be a safety-of-flight item. The primary flightcontrol wiring installation will require the same care an d prioritywhich is ow given to control cables, fuel lines, an d hydraulic lines.Isolation of the flight control wiring from the rest of the aircraft'swiring will be a must and special consideration will have to be givento dispersion, protective conduit an d channeling to avoid maintenancedamage, end-to-end hard wiring (i.e., no press-fit comiections), pro-tection against heat and fire damage, and cable jacket monitoringtechniques which permit damage detection before an actual failureoccurs. Where connectors ar e necessary, they must be designed to be"idiot proof" so that it is virtually impossible to force them togetherwrongly, even with the help of a large hammer. They must also bepositively sealed to exclude moisture or other contaminants. A specialeffort should be made to eliminate electrical connectors wherever possible.

Redundant Fly-by-Wire ActuatorsUntil recently very little research work had been done in hearea of redundant servo actuators at the two-fail/operate level. Thisdiscussion will be confined to hydraulic servo actuators since theyappear to be most suitable for aircraft requirements of the immediatefuture. That is not to say that pneumatic, electrical, or other typesof servo actuators will not be required or used on future aircraft.The redundant servo actuators currently under development fall con-veniently into the following three categories: (a) lectronic logicand switching; a two-fail/operate hydraulic servo actuator using electroniclogic and switching technique has been designed by the General E] ectric

Company (Johnson City) under the sponsorship of the Air Force FlightDynamics Laboratory an d is escribed in technical report AFFDL-TR-67-17(see Figure 13), (b) fail passive with electronic logic; the Sperry PhoenixCompany has designed a two-fail/operate, fail passive redundant servoactuator under their fly-by-wire contract with the Air Force FlightDynamics Laboratory, and it is escribed in echnical report AFFDL-TR-67-53(see Figure 14), (c) ydraulic logic and switching; Hydraulic Research andManufacturing Company has designed and built a single-fail/operate redun-dant actuator using hydraulic logic and switching which was installed andflight tested in n F-4C aircraft at Edwards Air Force Base this year (seeFigure 15). A similar two-fail/operate redundant actuator has been designedand is resently being built by HRM for flight tests during Phase III ofou r in-house fly-by-wire program to be explained later in he paper. Eachof these techniques has specific advantages and disadvantages in omparisonwith one another. The first, electronic logic and switching, can make fulluse of th e size and weight advantages to be gained by using MOS techniques.The use of well proven electronic logic techniques and the ease of faultdetection and correction through the use of modular packaging also makesthis method attractive. It oes, however, require transformation fromone power media to another; i.e., electrical to hydraulic, in order first

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to detect the fault and second to correct it. The result is an increasein switching time and a decrease in reliability. A typical solenoid valveoperates in 20 to 25 milliseconds. To keep the switching time down to areasonable value (say 50 milliseconds), it would be necessary for thedetection and electronic logic to take place in 25 milliseconds or less.Such a reaction time, although demanding, is within the current state-of-the-art. The second technique, fail passive with electronic logic, has theadvantage of no switching transients since all servo valves are operatingand when one or more fail passively, the remaining servo valves continueto drive the secondary actuator with negligible system performance degrad-ation. This system has a further ad.antage of being able to supply two-fail/operate redundancy with three servo valves. Th e main disadvantagewith this technique is that it is extremely difficult to design a 100%pure fail passive system. Consequently, it is necessary to include anelectronic model as protection against a hardover failure even thoughthe possibility of one occurring is remote. The third technique,hydraulic logic and switching, eliminates the power interface problemby performing all detection, logic, and switching functions in thehydraulic medium. The result is a decrease in switching time and anincrease in reliability. Typical detection logic and switching timesfor such actuators are less than 10 milliseconds. These systems aresensitive to contamination and silting in the hydraulic fluid, butstate-of-the-art filtering techniques can minirize this problem.

Integrated Hydraulic Servo Actuator Package ConceptThe use of a hydraulic servo actuator package consisting of anelectrically driven motor, hydraulic pump, accumulator, reservoir,servo valve and hydraulic power ram, all contained within the same

unit, dates back to World War II when the Germans employed them intheir Vl and V2 rockets and several of their fighter-bomber aircraft.Integrated hydraulic servo actuator packages, which were electricallypowered and controlled, were used as the rudder parallel actuator forautopilot directional control on the HE-11, JU-38, ME-110, and DO-17Luftwaffe aircraft. The success of these servo packages led to thetesting of an all-attitude autopilot employing such packages in allthree axes. Although these tests were successful and showed greatpotential because of the reduction in vulnerability they provided,lack of funds prevented further development or production by theGermans. A Siemans unit of this type was tested at Wright Air Develop-ment Center as long ago as 1950.

This technique has been successfully applied fcr thrust vectorcontrol on missiles. Figure 16 shows a typical missile servo actuatorpackage which consists of an electrically driven motor, hydraulic pump,accumulator, reservoir, servo valve and hydraulic power ram, all con-tained within the same unit and weighing about 22 pounds. It is asimple, non-redundant, short life system which would be unsuitable fordirect application to aircraft use. It does, however, demonstrate thatthis technique is within the present state-of-the-art. The aileron,elevator and rudder actuators on the VC-10 are an example of the use

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of this technique on current aircraft. Eleven similar units are usedon split control surfaces--two split ailerons (4), three rudders (3),and two split elevator surfaces (4). Redundancy is thus achieved rightout to the control surface, although each individual package is itselfnon-redundant. An example of the use of a redundant integrated packageis on the spoilers of the Belfast where a duplex integrated hydraulicservo actuator package is employed. Since these electrically poweredunits are completely self-contained and relatively easy to remove andreplace, the aircraft down time due to maintenance and repair is signi-ficantly shortened. Programs which are now underway at the Air ForceFlight Dynamics Laboratory will develop and flight test demonstrateredundant integrated hydraulic servo actuator packages for use in highperformance tactical fighter-bomber aircraft (See Figure 6).

Current AFFDL Fly-by-Wire ProgramsTh e fly-by-wire effort of rhe Control Elements Branch, FlightControl Division, Air Force Flight Dynamics Laboratory is divided

into two parts: an in-house effort being conducted in the ControlTechniques Laboratory in Building 195 by Hydraulic Research andManufacturing Company personnel working under contract; and twocontracted efforts (facetiously referred to as our out-house efforts)with Douglas-Long Beach and Sperry Pheonix Company.In-House Programs

The in-house program consists of the design, manufacture, assembly,and flight test of a single axis (pitch) fly-by-wire flight controlsystem for a B-47 aircraft. This program is being accomplished in thefollowing three phases:

Phase I - The existing B-47 control stick and feel system was usedin conjunction with a simple nonredundant fly-by-wire system. Lineardisplacement transducers (LVDT's) connected to the pilot's control stickoperated a servo actuator (modified 5-47 actuator) in parallel with theexisting aircraft pitch actuator. During tests of the fly-by-wire system,the normal aircraft pitch actuator was bypassed. Over 40 flight hourswere flown without a failure or malfunction in the fly-by-wire system.The test pilots observed a- appreciable improvement in aircraft responseto rapid inputs at high "4" flight conditions. The lag in the normalsystem could be attributed to aircraft cable stretch which, of course,was eliminated when using the fly-by-wire system.

Phase I! - A side-stick controller was installed in the pilot'scockpit and a C* feedback system was installed to provide the necessaryfeel/response. The same nonredundant servo actuator as was used inPhase I was used here. Flight tests are currently being conducted onthis system.

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Phase III- The side-stick controller and C* feedback system fromPhase II will be used but with a quadruply redundant actuator installedin place of the nonredundant servo actuator from Phase I. This actuator .will use hydraulic logic and will be powered by four 3,000 psi hydraulicpower supplies installed in he tail section of the aircraft. jThroughout this in-house effort, extensive use is being made of theunique facilities of th e AFFDL Fly-by-Wire Development Laboratory. Bydesigning, installing an d laboratory testing the fly-by-wire systems onthe B-47 tail section shown in Figure 17 prior to aircraft installation,much aircraft down time is being avoided. Figure 18 shows the testaircraft during one of its fly-by-wire test flights.The Douglas Effort

This consists of the completion of an extended program in whichmuch effort was expended in an attempt to design and build a pureelectrical (nonelectronic) a.c. primary flight control system. Thistask, as originally directed by AFFDL, proved impractical and theprogram was then redirected to permit the use of electronics and d.c.in an effort to obtain more positive results from the remainingresources. A breadboard model of a triply redundant fly-by-wiresystem was designed and built.The Sperry Phoenix Effort

This consists of : (1)'a fly-by-wire study and research contract,the results of which are included in Technical Report AFFDL-TR-67-53,"Fly-by-Wire Techniques", prepared by Mr. F. L. Miller and Mr. J. E.Emfinger of the Sperry Phoenix Company under the direction ofMr. V. R. Schmitt and F/L J. P. Sutherland, Project Engineers, FDCL,AFFDL: and (2) a recently completed contract to design and build athree-axes, quadruply redundant experimental laboratory model of afly-by-wire system for a B-47 aircraft. This system, which is shownin igure 19, is urrently being used in he AFFDL Fly-by-Wire Develop-ment Laboratory as a design tool for future fly-by-wire flight controlsystem development.Future Efforts

AFFDL's future programs in fly-by-wire include a program for theflight test demonstration of a complete three-axis, two-fail/operate(ouadruply redundant) fly-by-wire flight control system employingredundant integrated hydraulic servo actuator packages, on a tacticalfighter aircraft. This two year program is scheduled to begin in arly1969.

Summary and ConclusionsSimple direct mechanical linkages, cables, anti feel springs for

manual control can no longer cope with many of the control systemproblems associated with modern high performance aircraft and aerospace31

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vehicles. In an effort to meet the greater de=ands of these advancedaircraft control system requirements, the flight control designer hasbeen forced to increase the complexity of the mechanical system witha resulting increase in weight, volume, and cost, and a decrease inflexibility and reliability. Invariably he is forced to compromisebetween the desired performance and design requirements and a practicalmechanizaticn. Fly-by-wire offers not only to meet the demands ofthese advanced controi system design requirements, but also promisesto do so with a decrease in complexity, weight, volume, and cost andan increase in flexibility and reliability. Why then is fly-by-wirenot in comon use today? Th e answer to this question was given inthe first part of this paper; i.e., a lack of confidence in the conceptof fly-by-wire and a feeling of false security in mechanical flightcontrol systems. These are the principal factors which are now retard-ing the growth and general acceptance of fly-by-wire. The Air ForceFlight Dynamic-s Laboratory fly-by-hire programs are aimed at establish-ing the assurance level or level of confidence in fly-by-wire controlsystems among ilitary operators and aircraft manufacturers and designerswhich is necessary to overc. e this stigma. We recognize the inevitabl- 3existance of many engineering prebleas which m.ust be solved in goingfrom a drawing board design to flight worthy hardware. Although ourprograms will not necessarily establish the best solutions to theseproblems, they should demonstrate conclusively that the solutions arefeasible and practicable. We firmly believe that fly-by-wire is notonly inevitable for use in advanced military and connercial aircraftand aerospace vehicles but is, in fact, on the inmediate horizon. Ifthis paper has helped to convince you of this fact, or even encouragedyou to re-evaluate your previously held opinions, then it has servedits purpose well.

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Bibliography

1. Jenney, G. D.; A Study of Electro-Hydraulic Servo Control RedundancyMechanization, Report No. 78600011; Hydraulic Research and ManufacturingCompany, Burbank, California; October 1965.2. Malcom, L. G. and Tobie, N. N.; New Sho:xt Period Handling QualityCriterion for Fighter Aircraft, Docuwent No. D6-17841 T/N; Th e BoeingCompany, Airplane Division, P. 0. Box 707, Renton, Washington 98055;September 1965.3. Miller, F. L. and Erfinger, J. E.; Fly-by-Wire Techniques, AFFDL-TR-67-53;Flight Control Division (FDCL), Air Force Flight Dynamics Laboratory, Wright-Patterson Air Force Base, Ohio; June 1967.4. Rinde, J. E., et al ; Investigation and Development of Redundancy Tech-

to Achieve Dual Fault Corrective C!.pability in Flight ControlActuators, AFFDL-TR-67-17; Flight Control Division (FDCL), Air Force FlightDynamics Laboratory, Wright-Patterson AFB, Ohio; January 1967.5. Sethre, V. C., et al ; Design Technigu s and Laboratory Develovpent ofan Electrical Primary Flight Control Systsa, ASP-TDR-62-46; Flighz ControlDivision (FDCL), Air Force Flight Dynamics Laboratory, Wright-PattersonAir Force Base, Ohio; April 1962.6. Darrieus, Bernard; "Concorde Flight Control Systen", Paper given at SAEComittee A-6, San Francisco, California; 20 October 1966; SUD Aviation,633 Third Avenue, New York, New York 10017.7. Howell, G. C.; "Flight Experience of ?.ate Demand Control Using ElectricSignalling in the Avro 707C Aircraft", Paper given at A.G.A.R.D. FlightMechanics Panel Session I, RAE; Farnborough, Hampshire, U. K.; 10 May lc66.8. Ostgaard, M. A.: "Flight Control Syst(m Design for Supersonic Transport",Paper given at SAE-NAEW&M Meeting; Los Angeles, California; 8 October 1962;Air Force Flight Dynamics Laboratory, Wright-Patterson Air Force Base, Ohio.9. Gaul, J. W. , et al; Application of Optimal Control to VTOL ControlSystem Design, AFFDL-TR-67-102; Flight Control Division (FDCL), Air ForceFlight Dynamics Laboratory, Wright-Patterson Air Force Base, Ohio;May 1967.10. Johannes, R. P.; "Adaptive Control of Flexible Aircraft StructuralModes", Paper given at Joint Meeting of FligLt Mechanics and Guidance andControl Panels of A.G.A.R.D.; Oslo, Norway; 3 - 5 September 1968; AirForce Flight Dynamics Laboratory (FDCL), WrigIt-Patterson AFB, Ohio45433.

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11. Gut, S. J., '"odular Hydraulic Servos", Paper given at SAE AerospaceFluid Power Systems and Equipment Conference, Los Angeles, California,18-20 May 1965; Kearfott Division, General Precision, Inc.12. Boulton Paul Aircraft Limited brochure; "Douty Packaged System ofAircraft Control", March 1968; olverl'npton, England.

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Date post:	06-Apr-2018
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