NASA/TM-97-206224

Performance of an Electro-Hydrostatic

Actuator on the F-18 Systems ResearchAircraft

Robert Navarro

Dryden Flight Research Center

Edwards, California

National Aeronautics and

Space Administration

Dryden Flight Research Center

Edwards, California 93523-0273

October 1997

https://ntrs.nasa.gov/search.jsp?R=19970034702 2018-07-12T00:42:41+00:00Z

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PERFORMANCE OF AN ELECTRO-HYDROSTATIC ACTUATOR ON THEF-18 SYSTEMS RESEARCH AIRCRAFT

Robert Navarro

NASA Dryden Flight Research CenterRO. Box 273

Edwards, California 93523-0273USA

ABSTRACT DAC

DCIAn electro-hydrostatic actuator was evaluated at

NASA Dryden Flight Research Center, Edwards, EHACalifornia. The primary goal of testing this actuatorsystem was the flight demonstration of power-by-wire °Ftechnology on a primary flight control surface. The FCCelectro-hydrostatic actuator uses an electric motor to FCSdrive a hydraulic pump and relies on local hydraulicsfor force transmission. This actuator replaced the F- 18 FMETstandard left aileron actuator on the F-18 Systems FWTResearch Aircraft and was evaluated throughout theSystems Research Aircraft flight envelope. As of g

July 24, 1997 the electro-hydrostatic actuator had H/Waccumulated 23.5 hours of flight time. This paper Hz

presents the electro-hydrostatic actuator system IBITconfiguration and component description, ground andflight test plans, ground and flight test results, and Ibox

lessons learned. This actuator performs as well as the lbfstandard actuator and has more load capability than LMCSrequired by aileron actuator specifications ofMcDonnell-Douglas Aircraft, St. Louis, Missouri. Theelectro-hydrostatic actuator system passed all of its LVDTground tests with the exception of one power-off test MCT

during unloaded dynamic cycling. MDA

NOMENCLATUREMDC

MIL-STD-A/C aircraft 1553

ADC analog to digital converter ms

ATP acceptance test procedure mV

BIT built-in-test NASA

°C degree Celsius

CPU computer processing unit PCME

digital to analog converter

Dynamic Controls, Incorporated,Dayton, Ohio

electro-hydrostatic actuator

degree Fahrenheit

flight control computer

flight control system

failure mode and effects test

flightworthiness test procedure

gravity

hardware

Hertz

initiated built-in-test

interface box

pound-force

Lockheed-Martin Control Systems,Johnson City, New York

linear variable differential transformer

Mos Control Thyristor

McDonnell-Douglas Aircraft, St. Louis,Missouri

McDonnell-Douglas Corporation, LongBeach, California

Response Multiplex Data Bus(Military Standard)

milliseconds

millivolts

National Aeronautics and SpaceAdministration

power control and monitoring

PCUPBWramSOVSRAVWPAFB

electronics

powerconversionunitpower-by-wireshaftof theactuator

shutoffvalve(solenoidoperatedvalve)SystemsResearchAircraftvoltageWright-PattersonAir ForceBase

INTRODUCTION

The F-18 SystemsResearchAircraft (SRA) [1],(fig. 1), at the National Aeronauticsand SpaceAdministration(NASA) Dryden Flight ResearchCenter(DFRC)isa dual-purposetestbedbenefitingboth commercialand military developments.Aprimary goal is to identify and flight test newtechnologieson the SRA that will be beneficialto subsonic, supersonic,hypersonic,or spaceapplications.Oneofthesetechnologiesistheelectro-hydrostaticactuator(EHA) system,providedbyWright-PattersonAir ForceBase(WPAFB),Dayton,Ohio.UsingtheEHAeliminatesthecentralhydraulicsystemand reducescomplexity,maintenanceandsupportpersonnel.Thetwointerfaceboxes(Ibox)andthepowerconversionunit(PCU)usedfor thesetestswereprovidedby DynamicControls,Incorporated(DCI), Dayton, Ohio. Lockheed-MartinControlSystems(LMCS),JohnsonCity,NewYork,integratedthe actuator,motor,pump,andelectronics.DowtyAerospace,Duarte,California; Vickers,Jackson,Mississippi; and Electromech,Wichita, Kansas,providedtheactuator,pump,andelectricmotor.

Testingthe EHA systemdemonstratedin flightpower-by-wire(PBW)technologyonaprimaryflightcontrolsurface.ForthistesttheEHA(fig.2) replacedthestandardF-18left aileronactuatoron theSRA.The flight test program consistedof severalmaneuversthroughouttheSRAflightenvelope.ThisreportpresentsEHAtestresultsduringthemaneuversandcomparesthemto thestandardaileronactuator.Lessonslearnedduring ground testingare alsopresented.

SYSTEM DESCRIPTION

The SRA was retrofitted with a larger left wing hingehalf to accommodate the larger size of the EHA(fig. 3). Figure 4 shows the EHA integrated into theleft wing of the SRA.

The EHA system consists of the following elements:

• The electro-hydrostatic actuator (EHA)

• Two independent interface boxes (Ibox)

• One power control and monitoring electronics(PCME) unit

• One power conversion unit (PCU)

• The pilot control panel located in the F- 18 SRAcockpit

Figure 5 shows the interface of the EHA componentsand the two flight control computers (FCC's). TheFCC's (FCC ch 1 and FCC ch 4) generate servo-current position commands which are fed to theIboxes, one Ibox for each FCC channel. The Iboxes

receive the servocurrent position commands from theFCC's, generate commands to the PCME andgenerate a simulated feedback response to the FCC's.The simulated feedback is used for loop closure and tofool the FCC's into thinking that there is a standardactuator on board. The PCME receives the command

signals generated by the Iboxes, then compares thesesignals to verify that they are within a set limit andaverages them into a single command signal for theEHA. The EHA feeds status and health informationback to the PCME to be used for fault detection. The

PCU provides the high power necessary to the EHAthrough the PCME. Figure 6 shows the location of theEHA system components in the SRA.

Two EHA systems sets, PCME and EHA, weredelivered to DFRC as a primary and a backup. Thesets were not interchangeable and they were classifiedas set 1 and set 2. To meet performance requirements,each EHA motor resolver had a specific calibrationconstant, which was coded into the software of the

PCME. Therefore, the EHA and PCME were pairedtogether by the resolver calibration constant Thus,

replacing one of the units would require a softwaremodification to the PCME for the calibration constantof the resolver.

The Electro-Hydrostatic Actuator

The EHA utilizes a single positive displacementbidirectional bent-axis pump driven by a single three-phase, permanent magnet motor. The motor directiondetermines ram extension or retraction. The 41.5 lb

actuator is designed to require no active cooling. TheEHA contains a hydraulic fluid accumulator which is

packaged together with fluid components, and abypass shutoff valve (SOV). The hydraulic integratedmanifold contains two back-to-back check valves,two cross-port pressure relief valves and the solenoidoperated shutoff valve with an internal dampingorifice. The two back-to-back check valves allow the

reservoir to replenish fluid in the balanced actuatorcylinder. Cross-port pressure relief valves are used toprotect the actuator, mounting structure, and aileronsurface from overpressure and overstress. The SOVdisengages the actuator ram from the motor and pumpassembly, reverts cycling the fluid through an orificefrom one side to the other of the balanced actuator

ram, to provide the trail-damped mode. The EHA alsouses a linear variable differential transformer (LVDT)and a resolver for position feedback. It has atemperature transducer and a pressure transducer forstatus and health information used by the PCME. TheLVDT measures ram position, the temperaturetransducer measures motor winding temperature, thepressure transducer measures reservoir pressure, andthe resolver measures motor rotor position andvelocity.

The Power Control and Monitoring Electronics

The electronics required for controlling the power andposition commands to the EHA and the monitoring ofthe actuator responses are located in the PCME unit(fig. 7). The PCME is equipped with an 115 V accooling fan and weighs approximately 20 lb. Theelectronics monitoring provides continuous faultmonitoring and transfers the actuator to the trail-damped mode when any failure is detected. The

PCME communicates with the SRA FCC's throughthe two Iboxes. The inputs to the PCME providecommands for actuator mode and position. Theseactuator command signals are translated by the PCMEhardware and software into motor current commands

through appropriate control laws which enable theactuator to be driven to its desired position. ThePCME is responsible for actuator position, motorvelocity, and current (acceleration) loop closures. In

addition, the PCME software provides BITcapabilities to test the PCME computer hardware,sensors, servoelectronics and actuator mechanics.

The PCME receives the high power +135 V dc fromthe PCU and translates it to current commands for the

EHA through the Mos Controlled Thyristors (MCT).Regenerative energy from the EHA is stored in

capacitors. Any excess regenerative energy isdissipated through an external resistor bank.

The Interface Boxes

The Iboxes (fig. 8) provide the interface between theF-18 FCCs and the EHA system. The Iboxes providethe two FCC channels (FCC chl and FCC ch4) withall the required loop closures. Two Iboxes are used,one for each FCC channel. The Iboxes receive the

servocurrent position commands from the FCC's,generate the commands to be sent to the PCME, andgenerate the simulated feedback responses to theFCC's. The Iboxes also provide monitoring of theresearch actuator through the system and allowresearch and performance data to be monitoredthrough a MIL-STD-1553 Response Multiplex DataBus (1553 data bus) [2].

The PCME reports health status of the EHA system tothe Iboxes. Once the Iboxes receive a failure, the

simulated position feedback to the FCC's recognizeopen position feedback as a failure. Once the FCC'srecognize the failure, it de-energizes the SOV outputsignal which causes the Iboxes to command the EHAsystem into a trail-damped mode. The PCME alsode-energizes the SOV.

The Iboxes are designed to self-monitor internalfunctions. Because there is no cross-channel

communications between the Iboxes, discrepanciesbetween the two channel position commands receivedby the FCC's cannot be detected by the Iboxes.Therefore, each Ibox splits the servocurrent positioncommands into two redundant paths to generateposition command signals for the PCME. The twocommand signals are compared and their differencescalculated. If the difference is not within a set limit a

failure is declared and the position feedbacks areopened to the FCC's, transferring the EHA system tothe trail-damped mode. The Ibox self-monitoringsystem includes failures of power supply, softwareand central processing unit (CPU).

The Power Conversion Unit

The PCU (fig. 9) provides the +135Vdc, whichpowers the EHA system. The PCU takes the 115 V ac,400 Hz, three-phase F-18 aircraft power and convertsit to +135 V dc, which is delivered to the PCME. The

PCME then sends the high power through the MCT'sto the motor windings. The PCU is fused with two35-amp fuses to protect the aircraft electrical systemsand EHA system.

The Pilot Control Panel

The pilot control panel contains control switches forvarious experiments integrated into the SRA. Fourcontrol push-button switches and a toggle switch areallocated for the EHA system as follows (fig. 10).

• The lbox mode switch commands the EHA

system to be in Normal or Test mode, Normalmode puts the system in normal operation modeand Test mode is selected when performing anaircraft flight control system (FCS) BIT.

• The IBIT button commands the EHA systeminitiated BIT. IBIT is used by the PCME toverify the failure logic used to monitor correctoperation of the EHA prior to aircraft takeoff(weight-on-wheels is needed to invoke IBIT).After a successful completion of IBIT the buttonis lit to indicate a passed condition for the test.

• The reset button is used to reset the EHA systemafter IBIT or a failure. The system only resetswith weight-on-wheels (aircraft on ground).

• The PCME BIT enable switch, in the offposition, prevents one Ibox from sending anIBIT command to the PCME which preventsinitiating an IBIT on the ground. It is used as aprecaution to prevent an IBIT in flight.

• The left aileron disable switch allows the pilot toactively de-energize the SOV valve and transferthe EHA into a trail-damped mode.

GROUND TEST PLAN AND DESCRIPTION

The ground test plan includes the acceptance testprocedure (ATP) and flightworthiness test procedure(FWT) conducted by LMCS. These procedures werederived from the EHA statement of work producedby WPAFB, which referenced the MDA aileron

actuator specification. The validation test proceduresconducted by DFRC were developed in response to

the requirements set up by the statement of work,LMCS ATP, and FWT.

Testing by LMCS included electrical, mechanical,

and thermal testing necessary to qualify the EHA andPCME in the F-18 environment. The EHA portion of

the testing also included pressure stresses. In additionto the ATP and FWT, the PCME contained a preflight

and continuous built-in-test (BIT) to verify the systemoperation.

Flightworthiness Test Plan

The flightworthiness test (FWT) plan covered theenvironmental portion for the EHA system. This

test is a comprehensive one and ensures designrequirements are met for the EHA and PCME. The

FWT is comprised of the following tests:

• Altitude and rate of change

• Temperature (continuous, intermittent andshock)

• Voltage power transients (28 V dc and 270 V dc)

• Vibration (resonance, sinusoidal and random)

• Impact shock

• Threshold and Linearity

• Endurance

• Limit load

• Output force

• Ultimate load

• Insulation resistance

• Nondestructive inspection

Validation test

The validation test objectives for the EHA were met

by fully integrating the system with the F- 18 Ironbird.

The F-18 Ironbird has hydraulically operatedhorizontal stabilators, rudders and ailerons. Other

F-18 surfaces are simulated using analog actuatormodels. The Ironbird also includes a test bench to

perform open-loop and stand-alone testing of the

FCC's. The validation test objectives consisted of a

functional test, continuous operation test, frequencyand rate limit test, and the failure mode and effects

test (FMET).

4

Functional test--The functional test confirmed the

following operations of the EHA system:

• Perform a successful EHA reset

• Perform EHA BIT successfully multiple timesand consecutively

• Perform full-up and full-down commandssuccessfully

• Perform aircraft RIG mode operationsuccessfully

• Perform EPAD BIT in all flap settingssuccessfully

• Perform aircraft FCS IBIT and Test Group 10BIT successfully

• Perform hardware simulation and data

recording functional checks

Continuous test--The continuous test ensured that

the EHA system operated under normal conditionsfor an extended period of time. The test consisted ofoperating the EHA for five (5) 1-hour segments at0.3 Hz sinusoidal and are listed as follows:

• Amplitude of 50 percent of full stroke

• Aileron biased at 6 ° up

• Aileron biased at 0 ° deflection

• Flap switch set at half-flap setting (30 ° down)

• Flap switch set at Up/Auto setting (0 °)

Amplitude for some of the tests listed were adjustedto clear the stops of the actuator.

Frequency response and rate limit test--Thefrequency response test was conducted to obtainopen-loop and closed-loop response of the left (EHA)and right (standard actuator) ailerons. Therequirement is that the difference between the left and

right ailerons should not exceed 0.5 dB in gain and 5 °in phase. The open- and closed-loop frequencyresponse test was performed using the smallamplitude log sweep at 5 percent of full strokecommand over a frequency range from 0.1 Hz to15 Hz. The rate limit test was performed with theEHA system using a square wave input sweep withan amplitude of 95 percent of full scale with aconstant frequency of 0.1 Hz.

Failure modes and effects test--The purpose ofFMET is to obtain failure mode responses and effects

of failures on the EHA system. The FMET testconsisted of actuator signal management failures,

cockpit signal management failures and powerfailures during a simulated test condition. Theactuator signal management failures consisted ofinterrupting signals upstream and downstream ofIboxes, and verifying that the EHA fails to a trail-damped mode during a simulated flight condition.

The cockpit signal management failures includedinhibiting individual discrete signals during asimulated flight condition at different flap settings.The power failures consisted of interrupting power tocomponents during a simulated flight condition.

Actuator signal management upstream of Iboxes.

Failures for the actuator signal management upstreamof Iboxes consisted of interrupting the followingsignals at a simulated condition of Mach 0.6, at an

altitude of 25,000 ft during an aileron reversal, and

verifying EHA fails to a trail-damped mode onchannels 1 and 4:

• Left aileron command

• Left Aileron LVDTexcitation

• Left Aileron SOV

• Left Aileron pressure switch

• Left Aileron ram position

• Biasing aileron command

Actuator signal management downstream of Iboxes.For the actuator signal management downstream of

Iboxes, failures consisted of interrupting thefollowing signals at a simulated flight condition of

Mach 0.6, at an altitude of 25,000 ft during an aileronreversal, while verifying EHA fails to a trail-dampedmode on channels 1 and 4:

• Position command

• Biasing position command

• Position command PCME fault tolerancedetection

• Shutoff of actuator

• Position command fail

• Fail-safe mode

Cockpit Signal Management. The cockpit signalmanagement failures include inhibiting individualdiscrete signals at a simulated level flight of 160 kn

andanaltitudeof 25,000ft. Therequirementsareto(I) verify EHA fails to a trail-dampedmodewhencommandedbythedisableswitch,(2)verifytheEHAsystemdoes not reset throughthe resetswitch,(3)verify theEHA systemdoesnotperforma BITthroughtheEHAIBIT switch,and(4)verifyIBIT isnotperformedwithoutweight-on-wheelsdiscretes.

Component Power Failures. The power failures testsconsisted of interrupting power to the following

components at a simulated flight condition ofMach 0.6 and an altitude of 25,000 ft during anaileron reversal while verifying EHA fails to a trail-damped mode: (1) Iboxes, (2) PCME, (3) High powerto PCME through PCU, and (4) High power toactuator (all motor windings).

FLIGHT TEST DESCRIPTION

A test matrix was developed after the ground testdemonstrated that the EHA meets standard F-18

aileron actuator requirements. The test points weredevised using the F-18 simulation to map the F-18envelope with q-bar and aileron hinge moment as theprimary variables. The flight test matrix was brokeninto the following three phases, functional checks,envelope mapping and mission profiles and wereflown in that order. All three flight test phasesperformed the following maneuvers; doublets,windup turns, aileron reversals, straight and levelturns, lateral frequency sweeps, and aerobaticmaneuvers. A gradual hinge-moment buildupapproach was used in testing the EHA. Each phase offlight test was completed by conducting themaneuvers with medium-rate half-stick inputs andevaluating the performance of the EHA. A visualinspection of the EHA was performed between thefirst three flights. Upon completing the first threephases of flights, evaluating the data, and verifying byvisual inspection that no anomalies were found on theEHA; the phases of flights were repeated conductingthe maneuvers with abrupt full-stick inputs. The EHAwas required to complete 25 hours of flight time. Thefunctional checks phase was flown at a safe altitude of

25,000 ft and mach 0.4. The envelope mapping phaseof flights included altitudes of 10,000 ft, 20,000 ft,

35,000 ft and 40,000 ft with roach numbers rangingfrom 0.54 to 1.6. The mission profiles phase includedthe following aerobatic maneuvers; Wingover, BarrelRoll, Aileron Roll, 4-Point Roll, 8-Point Roll, Split S,Loop, Immelman, Cloverleaf, Cuban 8 and theChinese Immelman.

TEST RESULTS

Ground and flight test results are described next.Tables and figures are used to more clearly delineateresults.

Ground Test Results

The EHA successfully passed the ATP, FWT, andvalidation tests (functional, continuous andfrequency). Table 1 through 4 illustrate requirementsand test results. Figure 11 shows that the EHA andstandard open-loop frequency are within MDAaileron small amplitude specification limits. Figure 12shows that the EHA tracks closely to the standardactuator in the open-loop frequency response.

The FMET tests and results are presented in tables 5through 8. The EHA system successfully passed allFMET tests with the exception of one of the powerfailures (table 8). The power failures were devised tobe tested during an in-flight (simulation) dynamicmaneuver (aileron reversal). During the 28Vdcpower failure test of the PCME, the EHA system wentto a trail-damped mode as designed, but the powertransient caused damage to the MCT's. Results fromtroubleshooting indicated that the inverter drive usingthe MCT's was sensitive to loss of 28 V dc powertransients during dynamic cycling of the aileron(loaded and unloaded). The same test was conducted

previously with a static actuator (uncommanded) andit did not damage any MCT's in the PCME. Thecycling rate threshold that caused the failure was not

determined. No correction was incorporated into thehardware and the MCT failure was accepted as a lowprobability and severity risk.

Flight Test Results

The EHA has successfully flown on the SRA sinceJanuary 1997, and as of July 24, 1997 it hasaccumulated 23.5 hours of flight time. Table 9summarizes the first five flights, which include the in-flight failures and their causes.

The first EHA (set 1) flight took place onJanuary 16, 1996, on flight 564. Approximately 9 mininto the flight, the EHA system went into a trail-damped mode as designed. The SRA landed safelywithout extra effort from the pilot. Set 1 was qualifiedby similarity because it had completed and passed theATP and validation testing. Set 1 environmentaltesting was waived since the second EHA system

(set2) had completedand passedenvironmentaltesting.Duringtroubleshootingof thein-flightfailure,it wasfoundthatthePCMEwouldnotdetectanopenmotorphaseduringnormaloperationmode.Anopenmotorphaseis detectedby thepreflightBIT. Openmotorphasefailuredetectionwasincorporatedin thePCME software. In addition, load tests wereperformedtoquantifytheeffectofopenphasestothedynamicstiffnessof the actuationsystem.It wasdeterminedthat the EHA would have sufficientdynamicstiffnesswithoneopenphase.

ThesecondEHA system(set2) wasfirst flownonNovember4, 1996,on flight 583 andsuccessfullycompletedaircraft functionalflight checks.ThesecondflightwasflownonNovember12,1996,onflight584andincludedanaircraftenginecyclingtest.Theenginecyclingtestrequiresanengineshutdownandrestartin flight. Duringthefirstengineshutdown(right engine)the EHA systemwent into a trail-dampedmode. Thecauseof the EHA in-flightshutdownwasapowersurgecausedbyswitchingtotheleftgenerator,becausetherightengineiscoupledwith the right generator.During this generatortransition(less than 50 ms) the powersurgeispropagatedthroughthe28V dccontrolpowerto thePCME.The MCT's in the PCMEweredamagedbecauseof the powersurgeduringenginepowerdown.A powersurgefilteringcircuitwasinstalledonthe 28V dc powerto thePCMEto reducepowertransientstothePCMEduringengineshutdowns.

While set2 wasbeingrepaired,set 1(qualifiedbysimilarity)wasinstalledin theaircraft.Thecauseoftheflight564failurewassuspectedto bethelossof28V dc powerin the powersupplyinternalto thePCME.All componentsof thepowersupplyweretestedandno anomalieswerefound. A vibrationproblemwasspeculatedto bethecause,soset1wasvibratedtotheworst-caseaxis.Thisvibrationtestdidnot reproducethe failure. Set 1 was reflownonJanuary10, 1997,on flight 585 andshortlyaftertakeoffthe EHA systemwent into a trail-dampedmode. Thefailuresof set1for thetwoflights,564and585,werecorrelatedto altitude.Both in-flightfailuresoccurredat an altitudeof slightly above11,000ft. Thefailurewasreproducedin analtitudechamberand the causeof the shutdownwasdeterminedto bearcing. The leadsin thepowersupplywouldarcto acircuitboardlocatedabovethepowersupplyat higheraltitudes.The arcingwas

stoppedby addingalayerof captontapebetweentheleadsandthecircuitboard.

Figure 13 and 14 showthe in-flight failuresofflight564and585,respectively.Trace1showshowtheEHApositionfeedbackdepartsfromtheaveragepositioncommandatthetimeof failure.Theaveragepositioncommandisdesignedto moveto apositivet.9 volts,hardoverdown,oncetheEHA systemhasfailed.Thishardoverdowncommandduringafailurewasincorporatedtominimizethedifferentialbetweentheaileronsin casethe EHA systeminadvertentlycamebackonline,particularlywhenthe aircraftiscloseto theground.TheEHA is in a trail-dampedmodeonceafailurehasbeendetected.Trace2,withthe standardactuatoraileronpositioninvertedandtrim removed,is plottedagainstthe EHA aileronposition. As illustrated,theEHA positiondepartsfromstandardactuatorpositionatthetimeof failureasit failsto a trail-dampedmode.Trace3 illustratesSRAaltitudeatthetimeof failure.

Set2 wasintegratedinto theSRAagainandbeganflyingonJanuary16,1997withmediumrateinputs.As of July 24, 1997,the EHA system,set 2, hadcompleted24flights[3] totaling23.5hoursof flighttimeandhasflownflawlessly.Table10illustratesthemaximumandminimumtemperaturesthePCMEandEHAexperiencedandthemaximumhingemomentmeasured.

Figure15through18showtypicaltestmaneuvers.Trace 1 illustratestrackingof the EHA positionfeedbackwith theaveragepositioncommand.Thetimelag between command and feedback isapproximately30 ms and is associatedwith datasampling.In trace2 the standardactuatoraileronpositionis invertedandtrim is removed,andthenplottedagainsttheEHA aileronposition.HeretheEHA performsas well as the standardactuator.Trace3 showsthe EHA hingemoment,a negativehingemomentputs the actuatorin tensionand apositive hinge moment puts the actuator incompression.Figures19 and 20 illustratethe testpointswherethe EHA actuatorstalled.Figure21showsatestpointwherethestandardactuatorstalledandtheEHAdid not.

CONCLUDINGREMARKS

The ground test procedures were conducted to verifythat the electro-hydrostatic actuator and the power

control and monitoring electronicscomponentsmet design requirementsand were qualifiedasflightworthyfor theF-18aircraft.

Thegroundtestingphasepointedout thattheMosControlThyristorswhichareusedfor powercontrolin thepowercontrolandmonitoringelectronicsunit,aresensitiveto a lossof 28V dc powertransientsduringdynamiccyclingof theactuator(loadedandunloaded).Mos Control Thyristor'sare a newtechnologyand the manufacturerdid not havesufficienttestdatato characterizetheperformanceofthem.

Theflight testplanwasdevisedto testandevaluatetheelectro-hydrostaticactuatorperformancethrough-out the F-18 SystemsResearchAircraft flightenvelope.The electro-hydrostaticactuatorpositionfeedbacktrackedwellwiththepositioncommandandthesystemhasflownflawlesslysincetheadditionofthe28V dcpowersurgefilter to thepowercontroland monitoringelectronicsunit. The electro-hydrostaticactuatordid stall twice,asexpected,athigh hingemomentmaneuverswheretheexternalloadwasgreaterthanthemaximumoutputload.Theelectro-hydrostaticactuatorappearstohavemoreloadcapabilitythanrequiredby actuatorspecifications,andhasperformedaswell asthestandardactuatorthroughoutthe envelopeof the F-18 SystemsResearchAircraft. Generalperformanceof theelectro-hydrostaticactuatoris good.The fail-safedesignandtrail-dampedmodeworkedwellafterthreein-flight failureswereencountered.Pilots indicatethatflyingwithanelectro-hydrostaticactuatorontheF-18SystemsResearchAircraftfeelsthesameasif astandardactuatorwereonboard.

LESSONS LEARNED

The following lessons were learned throughout theEHA system program.

Controller to actuator interchangeabilit3,: In order to

meet position accuracy requirements, the system iscalibrated as a controller and actuator set. Two sets

were provided to DFRC for testing, a primary set anda backup set. Thus, the units are not interchangeable.Replacing any one of the units requires a softwaremodification to the calibration constant of the motor

resolver. Software management is an issue, especiallydeveloping acceptable tests to prove system

functionality, confirm successful software loading,and minimum acceptance criteria. It would have been

beneficial to the program to have interchangeableunits.

Open phase detection: The PCME detected an open

phase during the pre-flight BIT on the ground but did

not detect an open phase to the actuator during normal

operation. A design modification was required toprovide continuous detection on an open motor phase.

In addition, load tests were performed to quantify the

effect of open phases to the dynamic stiffness of the

actuation system. It was determined that the EHA

would have sufficient dynamic stiffness with one open

phase. It is recommended that input and output powerpaths of all components be tested.

Avionics power transients: Transients in avionics

28 V dc power may result in damage of the MCT's.

There were two contributing factors in the analysis of

this failure. First, the avionics 28 Vdc power and the

270V dc actuator control power were completely

isolated, which resulted in a hot 270 V dc bus duringthe avionics 28 V dc power failure tests. Second, the

MCT's, which uses avionics power, did not ensure a

fail-safe position of switching cells with loss of the

28 V dc power gate drive. As a result, the cells could

short the 270Vdc buses and damage controller

hardware (MCT's). A power surge filtering circuit was

incorporated to handle the control power of the MCT

as backup devices to handle temporary power spikes

or surges. Our experience with the MCT indicated

that they are sensitive to power surges and that theyrequire constant non-interruptible control power. The

PCME was designed with the understanding that the

28 V dc was on a non-interruptible (battery backed-

up) bus. The 28 Vdc bus in the aircraft is backed up

with a battery, but the switching time to the battery or

good generator (of less than 50 ms) was not pointedout in the design requirements.

Flight qualified by similari_: Two in-flight failureswere experienced because the PCME set 1 was

qualified by similarity. All tests were completed onset 1 with the exception of the environmental test. In

addition to necessary tests, when qualifying

components by similarity it is recommended to

include environmental tests that reveal workmanship

variationsor provideathoroughdetailedcomponentspecificationset of documentsand inspectionprocesses,sonon-conformingpartsorprocesseswillbeevident.

REFERENCES

[1] Sitz, Joel R., F-18 Systems Research Aircraft

Facili_, NASA TM-4433, Dec. 1992.

[2] Aircraft Internal Time Division Command/Response Multiplex Data Bus, MIL-STD-1553(DOD), Sept. 8, 1986.

[3] Systems Research Aircraft, NASA F/A- 18 # 845,Flight Report, Flights 583-590, FR-97-01 andFlights 591-600, FR-97-02, 1997.

[4] Aircraft Electric Power Characteristics,MIL-STD-704(DOD), May 1, 1991.

Table1.Electro-hydrostaticactuatoracceptancetestprocedureresults.

PCMEandParameter Units Requirement actuatorresults

Actuator sensor BIT actuator shutdown at sensors passed

shorted/grounded

Position command ch 1 = ch 2 = 0 + 50

match mv ch 1 cmd = 0 passed

mismatch V ch 4 cmd = +2.1 passed& ch 4 cmd = +1.9

Bus voltage too low (270 V) V Vbus=190 +10 passed

Demand BIT pass BIT passed

Hot/cold temp test of PCME

at -40 °C Perform BIT and normal passed

at 71 °C operation for 2 min passed

Output stall force lbf 13107 +500 13300sec 10 minimum > 30 minimum

Steady load lbf 5000 +500 5100min 20 minimum > 20 minimum

Ram output stroke in. (mechanical) +2.25 +2.25

in. (electrical) +2.19 +2.19

No load ram velocity in./sec 6.7 minimum 7.7 minimum

Load vs rate

load lbf 6329 6329

rate in./sec 4.88 minimum 6.0 minimum

Frequency response

+0.5 percent command Hz/dB/deg. 7/-7.25/92 maximum 7/-3.272/66

+5.0 percent command Hz/dB/deg. 7/-7.25/92 maximum 7/-3.27/65

Hysteresis percent of full stroke 0.1 0.080

Threshold percent of full stroke 0.05 0.0375

Linearity percent of full stroke +0.5 +0.1875

Dynamic stiffness lbf/in. 270,000 minimum 282,000

10

Table2.Flightworthinesstestresults.*

Parameter Units Requirement

Altitude ft 70,000Rateof change ft/min 40,000

TemperatureContinuous °F -40 to 160Intermittent °F 18010rainShock °F -40 to 160

Inputpowervariations28V dctransients V MIL-STD-71M[4]28V dcsteadyand V

rippletest Hz MIL-STD-704[4]270V dctransient V MIL-STD-704[4]

Vibration(eachaxis) perMDCA3376Resonantsurvey Hz 5to2000Dwellatresonance Hz 44.8& 49.8SineCycling Hz 5to 50to 5Random Hz 50to 2000

Shock(eachaxis)half-sine

g g 20-z,15x andy,EHAg 35-z,15x andy,PCME

time msec 11Threshold in. firstmotionnotto exceed

0.002in.command

Linearity percentof _+0.5full stroke

Endurance cyclesLimit load

25°up lbf0° neutral lbf45° down lbf

OutputforceCompression lbfTension lbf

Ultimateload25°up lbf0° neutral lbf45° down lbf

500,000

12,093(compression)15,800(tension)

13,106.67(tension)

12,093for 10sec13,106.67for 10sec

18,140(compression)23,700(tension)19,660(tension)

*PCMEandEHA,passed.

11

Table3:Validationfunctionaltestresults.*

Parameter Requirement

EPAD-reset Multipletimes

EPADBIT(Up/Auto) Multipletimes

CommandsFullup,deg 25Fulldown,deg 42

A/CRigMode Nofailures

EHABITHalfflaps NofailuresFullflaps Nofailures

A/CFCSIBIT Nofailures

A/Ctestgroup10BIT Nofailures

*PCME and EHA, passed.

Table 4. Validation continuous test results.*

Parameter Requirement

0.3 Hz sinusoidal at

AmplitudeAileron biased

Aileron Biased

Half-Flap

Up/Auto

50 percent of full stroke (- stack)

6 ° up0 °

-30 ° down

~0 °

*PCME and EHA, passed.

12

Table5:ValidationFMETresults(failureupstreamof Iboxes).*

Parameter Requirement

Interruptcommandchannel1 Openchannel4 Open

InterruptLVDTexc.channel1 Openchannel4 Open

InterruptSOVchannel1 Openchannel4 Open

Interruptpressureswitchchannel1 Openchannel4 Open

Interruptrampositionchannel1 Openchannel4 Open

Biascommandchannel1 +8 percent of commandchannel 4 _+8percent of command

*PCME and EHA, passed.

13

Table6:ValidationFMETresults(failuredownstreamof Iboxes).*

Parameter Requirement

Interruptcommandchannel1channel4

InterruptPCMEfaulttolerancedetectionchannel1channel4

InterruptSOVchannel1channel4

Interruptcommandfailchannel1channel4

Interruptfailsafemodechannellchannel4

Biascommandchannellchannel4

OpenOpen

OpenOpen

OpenOpen

OpenOpen

OpenOpen

+8 percent of command

+8 percent of command

*PCMEand EHA, passed.

14

Table7:FMETvalidationcockpitsignalsfailures.*

Parameter Requirement

InhibitdiscretesignalsofdisableswitchInhibit1 OpenInhibit2 OpenInhibit1& 2 Open

InhibitdiscretesignalsofresetswitchInhibit1 OpenInhibit2 Open

InhibitdiscretesignalsofIBIT switch Inhibit1Open

Inhibit2Open

*PCMEandEHA,passed.

Table8:ValidationFMETPowerfailures.

PCMEandactuatorParameter Requirement results

InterruptPowerIbox1 Remove28V dc PassedIbox4 Remove28V dc Passed

InterruptPowerPCME Remove28V dc Passed(staticEHA)

Failed(cyclingEHA)

InterruptPowerPCU Remove_.+135V dc Passed

InterruptPowerEHAwindings Remove+135 V dc Passed

15

Table9:Flighttestsummary.

Flight EHA/PCMEDate number hardware,set Failure

1-16-96 564 ! Powersupplyarcsabove11,000ft causingMCTdamageinPCME.

11-4-96 583 2

11-12-96 584 2

1-10-97 585 1

1-16-97 586 2

None

Powertransientduringin-flightengineshutdowncausedMCTdamagein thePCME

Powersupplyarcsabove11,000ft causingMCTdamageinPCME.Sameasthein-flightfailureinflight564.

(NoFailureshavebeenencounteredtodatesincethisflight.)None.

Table10:Flightresults.

Parameter Result FlightCondition

WarmestTemperatureEHA 35°CPCME 44°C

ColdestTemperatureEHA 1.25°CPCME - 1.77°C

HighestHingeMoment 57,300in-ibf

Mach0.42,Altitude25,000ftMach0.7,Altitude25,000ft

PriortotakeoffMach0.6,Altitude25,000ft

Mach1.6,Altitude35,000ft, q-bar890lbf/ft2

16

Figure 1. Systems Research Aircraft (SRA).

EC93 42065-06

Source: Lockheed Martin Control Systems, Johnson City, New York

Figure 2. Electro-hydrostatic actuator (EHA).

17

Figure 3. Hinge half housing.

@

®

EC97 43875-02

Figure 4. Electro-hydrostatic actuator installed in SRA.

18

n

Interface box ,-

Yl channel1 .o¢.,

,--orbank H C

I "t °28 vdc o

essential _¢

bus _=t.

-' }channel 4 -- _in_L _

Figure 5. Systems Research Aircraft system architecture.

EHA actuator

Aircraftinstrumentation

system

970874

970e_ j

Figure 6. Electro-hydrostatic actuator system component locations on SRA.

19

EC93-41023-03

Figure 7. Power control and monitoring electronics (PCME).

EC93 41023-09

Figure 8. Interface box (Ibox).

2O

EC9343334-01

Figure9.Powerconversionunit(PCU).

_ INSTR TM1 _ TM2 UPLINK EVENT(TAPE

MASTER REC

Or_'_3_'_ _ FADS FADS ACT _ IBOX FIBOX/ACT-1RESET POWER POWER MOOE IBIT

OF

ARTS ARTS _ PCME BIT ZEROPOWER RESET ENABLE CAL

Figure 10. Electro-hydrostatic actuator disable switch and control panel in cockpit.

970875

21

-2

-4

-6

Amplitude - 8ratio,

dB

i i i i i ! i i ! _Upp;rlimit i ! ! i i i

, _ , _ _ , "_.-P._..._.._. , , , , , , ,

-20 ......... _ ..... ;----i ! __.__ _ _'__ ....... -_"_.:,,-;_,_ _'_ _. _--_Standard_ '__:_'_.........

-_o.........................._-:-:-_:....._ ......,--_,-.................

- 60 ........ - ............ ,- .............

Phase .... ,, ,, : ;, ,, : : :-_,, :-_,: :

,+ _+o ,, .........i....deg

i :-100 ...................... '- .............................. ..+...... _, .... +,_ ......

: : : : : : : ', ,' : : : : :_ , •,\

_12o........................._ .....::__::_i___,,

- 140 ......... -_..... :.... :--- "- ........................... \-\_ ;,+;_: i ', ', : i ', ', ' ', i : : ', : : \ ,.

-160 .................... ',- -',- - _- -',- _+-:--' ................ ',--- "- - -:- -i - _ -',- -',-:....

-180 _ r _ ' , + I , , r r e ' , ' I _ P10 -1 10 0 101

Hz970876

Figure l I. Electro-hydrostatic actuator and standard actuator open-loop small amplitude frequency response.

22

-10

S_n0.,0_ : ii ilii : : i i i ilil

-20 ........ -' ..... '.... '--- ,'-- -;- -:--:- ,'--',- ...... --_- - - - -;- -¥ -:-- -;-- ;- }/-;- -:- ;- -_-',,-p

-25 ...................... i____.__J_i_;.............. ___i\.....f_i_i_;_i....Amplitude .......... _, _ , I .....

ratio, ' _\ /

- 3o ........ _..... _---_--,_--:--:-_-:-:......... :..... _:,........i i i i i iiii i : i'=i'_iiii

_ 40 ......................................

i , , _ q

, i , i i

i , i b , , , , i , , ,

-45 , i J J = , _ _ , , , L , = h , , ,

Phaselag,deg

0l .... _-- _L_ i i ! /--Standard i i i i i i i i, , ,, ,, __.;../ , , , : : ; : : :

-5oI.........';........._-C-I-_-- --'-"-,.-:......:....:---:-_--';-',--:-';.............. _ , , ,/-_ : ' ,,I : : ,', i _-;,i "_',,, : : :,,_,:::

_,ooI........_::.....:....:_:_::::: : ,,: : , ,,::::i i_i_ii........._i, _i_!,__!_!............... \" ' If ' \\' ' '

'%;t '1,i

_,oo......................!-!-i-il.............i---i.....i-!-i--i\,--250 ........ : ..... ' .... '- --;- --;--:--;-:- 4 ......... '...... ' .... '- - -:- - ',-- .;-- ,'--:- ;-- - -

- 30010-1

, , , , , _ , , , , , ,

i i i I I I I I I I I I I i i I I I

100 101Hz

970877

Figure 12. Electro-hydrostatic actuator and standard actuator closed-loop small amplitude frequency response.

23

2.5

Averageposition

commandand

feedback,V

Alleronposition,

deg

Trace 1

bsck --1.0 .................... :--...........e

o .................... i .................. -----=---n- ....- .5 ........ :....... i.....

I

- 1.0 ....................................... _ .................... : ...... I .....: ' ' I

- 1.5 , , ,

50 1 Trace 2

4O

3O

2O

10

-10

................ J

f Left................... r ........... i ....... _ .... ,............... n ...........

',

....... ± .... £ .............. J ...........

' _Right_;;:___; _- ._. _ ..........

16 103

Altitude,ft

14

12

10

8

6

4

2

0

-2

Trace 3

iiiiii:iiiiiiiiiiiiiiiiiiiiiiiiiiilIiiiii:iiiiiiiiiiiiiii:::................... _ ................... _ ...... £ ............ _ ...........

200 400Time, sec

Figure 13. In-flight failure, January 16, 1996, flight 564.

600

970878

24

Averageposition

commandand

feedback,V

2.0

1.5

1.0

.5

0

--.5

- 1.0

35

Aileronposition,

deg

Trace 2 ' '

o°15

p Ii p

5 ................. ................ -R-igl_t-.......... ..................

200 x 102

Altitude,ft

150

100

5O

Trace 3 __o_

J

1 0 200 300 400

Time, sec 970879

Figure 14. In-flight failure, January 10, 1997, flight 585.

25

Averageposition

commandand

feedback.V

1.0

.5

-.5

- 1.0

- 1.5

2.'

2O

Trace 1

............. _''_..................Feedback--/ t _,t!-- -_,- - - i_, !: .............. :_.............

..............i.......---)_,----#-........._..............i.............Command--J .__ jl_ _

Trace 2i

15 .............. _.............. ._L-eft........._ _ .............. ._.............

10

5Aileron

position, 0 ' - - - - '_...... -_.............. _= - ........ --_deg

-5

-1o ............. !............. !________..,o__t__i .............. i .............

-15 ................ L .............

q

- 20 , , , ,

15 x 103

Hingemoment,

in.-Ibf

Trace 3

5 ............. 7 ............. ,............. 7 .............. r .............

0

_1oiiiiiiiiiii-15

- 20 I

0 4I J

2 3Time, sec

Figure 15. Abrupt roll doublet--full stick (left).

M=0.72, altitude=25,000 ft, q-bar=284 lbf/ft 2.

5

970880

26

"2/Trace1 i i ! :ll i

andlblv ii!i- .6 - - -' ............. ! '- .....

Average

positioncommand

feedback,

10

0Aileron

position,

deg-5

r

Trace 2

i

-10 ....................................................................

_Right

-15 , , , , ,

15 x 103

Trace 3 ' ' :

101 ............ ! ............. i ............. i ............. :............. i ....

5 | ............ i, ................................................. _ ....

Hinge

iomen 0 ............ _ ............. :.......... _ ............. _ .....in.-Ibf

- 5 ..... ' ........... _ .....

-10 ' :

-15

moment,

h 1'5 L5 10 20

Time, sec

Figure 16. Slow--fast lateral frequency sweep.

M=0.83, altitude=25,000 ft, q-bar=376 lbf/ft 2.

i

25

970881

27

- .15

- .20

- .25

Average - .30position

command - .35

end - .40feedback,

V - .45

- .50

- .55

- .60

Tr.ac_e1 ...... : ........... : ........ _--i ........... i ........... i ...........

i : ' ' '

........... _ ................... _ ........... _ ........... J ...........

Feedbacki _

Aileronposition,

deg

-1

Trace 2

................................

b

0x 102

L-7 ........... '; ........... _ ...........

i i i

Hingemoment,

in.-Ibf

- 50

- 100

- 150

- 2OO

- 2500

Trace 3 : : : ;

iiiiiiiiiii,iiiiiiiii_ , , "5 1() 1'5 20 2'5 30

Time, sec 970882

Figure 17. Windup turn to 20 ° c_.

M=0.83, altitude=25,000 ft, q-bar=376 lbf/ft 2.

28

1.0

Averageposition

commandand

feedback,V

.5

0

--.5

- 1.0

- 1.5

-2.0

20

Aileronposition,

deg

15

10

5 ........... _r...........

o ! ....-5 ........... ! ...... --_-_ ....

-10 ............ '_.... -- _,----

-2o

30 x 103

Trace 2t

........... P............ L__

i

i i i

t q t

i I t n

Hingemoment,

in.-Ibf

20

10

0

-10

- 20

- 30

- 400

Trace 3

I I I I

1 2 3 4Time, sec

Figure 18. Abrupt aileron reversal--full stick (left).

M=0.84, altitude=25,000 ft, q-bar=-382 Ibf/ft 2.

6

970883

29

1.5

Averageposition

commandand

feedback,V

1.0

.5

0

--.5

- 1.0

- 1.5

- 2.0

2O

Trace 1 _ '

..................... i_m:" _.... i .................... i .........

..................... :__r_F_e_-b'_ck_i,_...... j .................... ':.........

li ............

t

J I I

Trc215 ..................... ,................ _ .................... r .........

............. _', :

10

5

-20 , , i

60 x 103

Trace 3 A i ', :/

40 ............... : ................ " .................... _ .........

Hinge i20 ............. i ................................iiiiiiiiiii_i .........moment,

in.-Ibf

-2

-40 .................... ' l--------J .......................... 'r .........

- 60 i0 4 6

Time, sec 970884

Figure 19. Aileron reversal--half stick (left).

M= 1.58, altitude=34,000 ft, q-bar=890 lbf/ft 2.

3O

1.5

Averageposition

commandand

feedback,V

1.0

.5

0

b,5

- 1.0

- 1.5

- 2.0

2O

Trace 1

-comma_ ii -i i I

Feedback .--_ _ li I_

........_l', t........._ ......._....

Aileronposition,

deg

Trace 2

15 .............

Left_ IS

10 .....................

5

0 --- lL' Right

-15

60 x 10 3/

40 I Trace 3

L20

Hingetomen 0

in.-Ibf _-4020l

- 60

moment,

200 400 600Time, sec

Figure 20. Abrupt aileron reversal--full stick (right).

M= 1.6, altitude=34,000 ft, q-bar=-925 lbf/ft 2.

800

970885

31

1.0

EHA

averageposition

commandand

feedback,V

Standardactuatorposition

commandand

feedback,deg

Aileronposition,

deg

.5

0

-- ,5

- 1.0

- 1.5

- 2.0

25

20

15

10

5

0

-5

-10

-15

- 20

2O

15_i

loi

5

0

-5

-10

-15

- 200

, , , , i

Trace 1

...........i...........i...........i...........i...........Feedback" // i 4, _ : : :

5/,,"7,,"......i...."--'i,.-./i i......---t-"--P_ ........_ ........'...........:...........

;;mm;n:_-.S'-.......i...........i..........._..........._...........

.:::=_e:.....:......:_:: .....i...........i............i...........

"e--c'i---r,'--7....i....._, : : :

Trace 3 : _i_-_-ii!i i i

f_.--_-_-_-

......::::::::-<_--<-;_iiii ! _......,__............::::::::::::::::::::::::::::::::::::i i

I I I I I

1 2 3 4 5 6

Time, sec 970886

Figure 2 l. Abrupt aileron reversal--full stick (left).

M=0.95, altitude=20,O00 ft, q-bar=600 lbf/ft 2.

32

REPORT DOCUMENTATION PAGE FormApprovedOMB No. 0704-0188

Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering andmaintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection o_ information,including suggestions for reducing this burden, to Washington Heaclquarters Services. Directorate for Information Operations and Reports. 1215 Jefferson Davis Highway, Suite 1234, Arlington,VA 22202-4302. and to the Office of Management and Budget. Paperwork Reduclion Project (0704-0188), Washington. DC 20503.

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORTTYPE AND DATES COVERED

October 1997 Technical Memorandum

4. TITLE AND SU BT1TLE 5. FUNDING NUMBERS

Performance of an Electro-Hydrostatic Actuator on the F- 18 SystemsResearch Aircraft

6. AUTHOR(S)

Robert Navarro

7.PERFORMINGORGANIZATIONNAME(S)ANDADDRESS(ES)

NASA Dryden Flight Research CenterP.O. Box 273

Edwards, California 93523-0273

9.SPONSORING/MONITORINGAGENCYNAME(S)ANDADDRESS(ES)

National Aeronautics and Space AdministrationWashington, DC 20546-0001

529-31 -14-00-36-00-SRA

8. PERFORMING ORGANIZATION

REPORT NUMBER

H-2210

10. SPONSORING/MON_ORING

AGENCY REPORT NUMBER

NASA/TM-97-206224

11.SUPPLEMENTARY NOTES

Presented at the 16th Digital Avionics Systems Conference, Irvine, California October 26-30, 1997

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified--Unlimited

Subject Category 05

12b. DISTRIBUTION CODE

13, ABSTRACT (Maximum 200 words)

An electro-hydrostatic actuator was evaluated at NASA Dryden Flight Research Center, Edwards, California.

The primary goal of testing this actuator system was the flight demonstration of power-by-wire technology on

a primary flight control surface. The electro-hydrostatic actuator uses an electric motor to drive a hydraulicpump and relies on local hydraulics for force transmission. This actuator replaced the F- 18 standard left aileron

actuator on the F- 18 Systems Research Aircraft and was evaluated throughout the Systems Research Aircraft

flight envelope. As of July 24, 1997 the electro-hydrostatic actuator had accumulated 23.5 hours of flight time.

This paper presents the electro-hydrostatic actuator system configuration and component description, ground

and flight test plans, ground and flight test results, and lessons learned. This actuator performs as well as the

standard actuator and has more load capability than required by aileron actuator specifications of McDonnell-

Douglas Aircraft, St. Louis, Missouri. The electro-hydrostatic actuator system passed all of its ground tests withthe exception of one power-off test during unloaded dynamic cycling.

14. SUBJECT TERMS

Electro-hydrostatic actuator, F-18 Systems Research Aircraft, performance,ground test, flight test

17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION

OF REPORT OF THIS PAGE OF ABSTRACT

Unclassified Unclassified Unclassified

NSN 7540-01-280-5500 Available from the NASA Center for AeroSpace Information, 800 Elkriclge Landing Road,

Linthicum Heights, MD 21090; (301)621-0390

15. NUMBER OF PAGES

37

16. PRICE CODE

AO320. LIMITATION OF ABSTRACT

Unlimited

Standard Form 298 (Rev. 2-89)Prescribed by ANSi Std Z39-18

298-102