NASA Technical Memorandum 104324

/ ,_ , F

" .... i" " +7 Z_

Design and Integration of

an Actuated Nose Strake

Control System

Bradley C. Flick, Michael P. Thomson, Victoria A. Regenie,

Keith D. Wichman, Joseph W. Pahle, and Michael R. Earls

October 1996

National Aeronautics and

Space Administration

NASA Technical Memorandum 104324

Design and Integration of

an Actuated Nose Strake

Control System

Bradley C. Flick, Michael P. Thomson, Victoria A. Regenie,

Keith D. Wichman, Joseph W. Pahle, and Michael R. Earls

NASA Dryden Flight Research Center

Edwards, California

1996

National Aeronautics and

Space Administration

Dryden Flight Research Center

Edwards, California 93523-0273

DESIGN AND INTEGRATION OF AN ACTUATED

NOSE STRAKE CONTROL SYSTEM

Bradley C. Flick, Michael E Thomson, Victoria A. Regenie

Keith D. Wichman, Joseph W. Pahle, Michael R. Earls

NASA Dryden Flight Research Center

Edwards, California

ABSTRACT

Aircraft flight characteristics at high angles of attack can be improved by controlling vortices

shed from the nose. These characteristics have been investigated with the integration of the actuated nose strakes for enhanced rolling (ANSER) control system into the NASA F-18 High Alpha

Research Vehicle. Several hardware and software systems were developed to enable performance

of the research goals. A strake interface box was developed to perform actuator control and failure

detection outside the flight control computer. A three-mode ANSER control law was developed

and installed in the Research Flight Control System. The thrust-vectoring mode does not command

the strakes. The strakes and thrust-vectoring mode uses a combination of thrust vectoring and

strakes for lateral--directional control, and strake mode uses strakes only for lateral-directional

control. The system was integrated and tested in the Dryden Flight Research Center (DFRC) simulation for testing before installation in the aircraft. Performance of the ANSER system was monitored in real time during the 89-flight ANSER flight test program in the DFRC Mission Control

Center. One discrepancy resulted in a set of research data not being obtained. The experiment was

otherwise considered a success with the majority of the research objectives being met.

NOMENCLATURE

ADC

AIL

ANSER

AOA

AUTO

BADSA

BIT

BRT

CAS

cmd

CONT

analog-to-digital converter

aileron

actuated nose strake for enhanced rolling

angle of attack

automatic

backup airdata sensor assembly

built-in test

brightness

control augmentation system

command

contrast
DAC

DDI

DEGD

DFRC

DPRAM

EEPROM

EHV

FCC

FCS

G

g

GN

GND

HARV

HATP

HUD

h

INS

I/O

L

LaRC

LEF

LVDT

M

MC

MCC

MNVR

NASA

N,

OAV

OBES

P

PAL

digital-to-analog converter

digital display indicator

degrade

Dryden Flight Research Center, Edwards, California

dual-port random access memory

electrically erasable programmable read-only memory

electrohydraulic valve

flight control computer

flight control system

gain set

acceleration due to gravity, 32.2 ft/sec 2

gain

ground

High Alpha Research Vehicle

High-Angle-of-Attack Technology Program

head-up display

altitude, ft

Inertial Navigation System

input--output

left

Langley Research Center, Hampton, Virginia

leading-edge flap

linear variable displacement transformer

Mach number

mission computer

Mission Control Center

maneuver

National Aeronautics and Space Administration

lateral acceleration, g

normal acceleration, g

onboard augmented vehicle

onboard excitation system

pitch

programmable array logic

2
PDS

PLA

Pos

PROC

R

RAV

RFCS

ROM

Rt

RUD

S

SIB

SOV

STAB

STV

SV1

SV2

TD

TEF

TH1

TV

TVCS

TVV

TVVS

V

Vdc

VME

Vrms

WOW

Y

6¢

parameterdisplaysystem

powerleverangle,deg

position

processor

roll

remotelyaugmentedvehicle

researchflight controlsystem

read-onlymemory

right

rudder

strakemode

strakeinterfacebox

shutoffvalve

stabilator

strakesandthrust-vectoringmode

servovalve1

servovalve2

trail damped

trailing-edgeflap

threshold

thrust-vectoringmode

thrust-vectoringcontrolsystem

thrust-vectoringvane

thrust-vectoringvanesystem

volts

volts,directcurrent

VersaModuleEurocard

volts root-mean-square

weightonwheels

yaw

angleof attackrate,deg/sec

angleof sideslip,deg

angleof sidesliprate,deg/sec

sum
INTRODUCTION

The F-18 High Alpha ResearchVehicle(HARV) flown at DrydenFlight ResearchCenter

(DFRC),Edwards,California,is a flight testplatformfor theHigh-Angle-of-AttackTechnology

Program(HATP). OneHATPobjectiveinvolvesimproving aircraftcontrollability andagility at

highanglesof attack(ref. 1).Thefirst phaseof theprograminvestigatedhigh-angle-of-attackaerodynamicsin a baselineF-18configuration.Duringthesecondphase,theaircraftwasmodifiedby

integratingaresearchflight controlsystem(RFCS)with thebasicF-18controlsystem.TheRFCS

allows variousflight control lawsto beflown. ThestandardF-18systemprovidesinput,output,

redundancymanagement,andsafebackupfor theRFCS.A thrust-vectoring(TV) systemwasalso

addedto theHARV. Thismodificationincludestheadditionof TV vanesto deflectengineexhaust

for high-angle-of-attackmaneuvering.

The third and final phaseinvestigatedan alternatecontrol methodologyfor high-angle-ofattackmaneuvering.LangleyResearchCenter(LaRC),Hampton,Virginia,developedasystemto

applyforcing momentsto thenoseof theaircraftbycontrollingthevorticesshedfrom thenoseat

highanglesof attack(ref.2).Theactuatednosestrakefor enhancedrolling (ANSER)systemconsistsof two actuatedcontrol surfacesinstalledin aspeciallyconstructedradome(fig. 1).

Figure1.TheF-18HARV.

EC96-43598-11

TheDFRCintegratedtheANSERsystemwith theRFCSandaircraftsystemsandconducted

the flight researchto accomplishthesetasks.Severalhardwareand softwaresystemswere

designedanddeveloped.Thispaperdescribesthesystemsdevelopedto integratetheANSERinto

theaircraftandintegrationtesting.Someof thetechnologiesdevelopedtoenableaccomplishment

of the researchobjectivesarediscussed.Useof tradenamesor namesof manufacturersin this
documentdoesnot constitutean official endorsement of such products or manufacturers, either

expressed or implied, by the National Aeronautics and Space Administration.

SYSTEM OVERVIEW

The ANSER installation was designed to use full actuator authority to provide 90 ° of strake

deflection (fig. 1). These strakes are commanded independently by control laws residing in the

RFCS computers with two independent channels controlling each strake. The RFCS computers are

F-18 flight control computers modified by the addition of a second processor running parallel and

sharing data with the main processor. Figure 2 shows a top-level diagram of the control system.

The RFCS software is written in Ada and cross-compiled on a UNIX®-based workstation. The

compiled software is then downloaded to the RFCS.

FCC

m

/°t

RFCS P

control I-- R Processor

laws | A

I M

Left strake

(channels 1, 4)

Strske command

Aerosurfaces/

TVCS and

sensors L

Strake go

SIB

I

Channel A Channel B

strake I strake

servoloop servoloop

and failure I and failure

detection detection

I

Right strake

(channels 2, 3)

Figure 2. The F-18 HARV ANSER control system.

960402

The RFCS control laws can be engaged with weight on the wheels or when the aircraft is in a

predefined Mach number, M, and altitude, h, flight envelope (15,000 ft < h < 45,000 ft, M < 0.7)

by first arming the system with a cockpit-mounted toggle switch, then engaging with a push button

on the control stick. Research objectives required operation of the strakes both in conjunction with

and independent of the TV system, resulting in a multimode control law implementation. Upon

engagement, the RFCS defaults to TV mode in which the strakes are commanded closed at all

times. Strakes and thrust-vectoring (STV) and strake (S) modes are selectable via a push button on

the digital display indicator (DDI).

®UNIX is a registered trademark of UNIX System Laboratories, Inc.
Unlike the other control surfaces, which are only commanded by the RFCS control laws while

RFCS is engaged (ref. 3), the strakes are commanded by the RFCS control laws at all times through

previously unused analog outputs. Similarly, actuator control for the conventional surfaces and TV

vanes are performed by the flight control computer (FCC); meanwhile, the strakes are controlled

by an external interface box.

HARDWARE DESIGN AND IMPLEMENTATION

Control of the strakes is performed by two electrically controlled, hydraulically powered

servoactuators. These actuators are F-18 aileron actuator servovalves mounted to a longer, narrower main piston and cylinder. The strake actuator piston length is 5.68 in. _+0.05 in. in place of the

4.38 in. _+0.05 in. as installed on the standard actuator. The actuators are dual redundant. Either

channel can drive the surface. Hydraulic supply is single string to both actuators. The hydraulics

are plumbed to the gun hydraulic quick disconnect located in the nose barrel section of the aircraft.

Integration of the strake actuators with the FCS required the addition of actuator control loop

closure and failure detection for the strake actuators. These functions are normally performed in

the FCC software. However, insufficient FCC input--output (I/O) availability required implementation of these functions external to the FCC in a strake interface box (SIB). The RFCS strake commands are passed through a previously unused digital-to-analog converter (DAC) to the servoloop

located in the SIB. The servoloop performs loop closure and sends a rate command to the actuator.

Several control loop signals are monitored by the failure-detection circuitry, which sends a go/no

go signal back to the RFCS via a previously unused analog-to-digital converter (ADC). If a no go

signal is received, the RFCS automatically reverts to TV mode and commands the strake to close.

The SIB contains two independent channels of actuator control and failure detection. Each SIB

channel controls one channel of each actuator (fig. 2).

Because the actuators used to drive the strakes are modified F-18 actuators, spare servoloop

circuit cards were modified for use in the SIB to reduce manufacturing costs for a newly designed

card. These cards were originally designed for use in the thrust-vectoring vane (TVV) system. The

servoloop cards contain three identical servoloops of which two were used for the strake system.

Figure 3 shows a block diagram of the servoloop board. Servoloop integrity is monitored by comparing the command output to that of a servoloop model which runs in parallel with the actual

servoloop. A miscompare results in generation of a failure signal which is passed to the failuredetection circuitry for system monitoring. In addition to the servoloop function, this card performs

all signal conditioning necessary for signal usage by the failure-detection logic. The card also receives discrete information back from the failure-detection logic for use in the servovalve and shutoff valve driver circuitry. This card was originally intended to be addressed by the FCC, so minor

modifications were made to permanently enable discrete I/O of the board. A 1.8432-MHz clock

was also added to these boards to trigger a programmable array logic (PAL) integrated circuit.

The software failure detection routines implemented in the FCC for the TVV actuators were

also implemented in the SIB for the strake system. These routines include a comparison of servovalve command current to servovalve position (fig. 4), a main ram linear variable displacement

6
Level

Left strake i---------...------7 _ _ detector Servomodel T. Filter Scaling Limiter _ /_ I , _ current I .I v _ r I .... I - _ I ......... I _-e-_T._ Filter _ __ fail to

I c°mman° I _ :_ .. .._ Z A y I , i 1 failure

" [_- Model ram position _ T T detection

[ Reset J I I Servocurrent

[__t Scaling ] to future detection

r

1 _ ,

, ] I -,r I Servoetatus

Left strake _).__ Filter _ Scaling H Limiter _ Convertto _,h I To actuator

surface current I .... I " I

command J _ _ _ _ I " _ v _..__._:_ iTo failure detection

-- I IToremote Y

EHV relay on --_-_ _'_

Ram position to

failure detection

I Left strake _ Convert

ram position to Vdc

from actuator

LVDT Convert

to Vdc

LVDT sum to

failure detection

_0_3

Figure 3. Servoloop card.

Spool position

Servocurrent _

-n'l -TH1 0 TH1

Spool position

[ I ]_- versus

servocurrent fail P-,.

Hydrau,c _ "_

pressure off _- 96o4o4

Figure 4. Spool position versus servocurrent test.

transformer (LVDT) monitor (fig. 5), and a hydraulic pressure sensor monitor (fig. 6). Each of

these tests generates a discrete signal which is passed back to the servoloop board for actuator shutoff valve and servovalve relay control. Figure 7 shows the relay control logic. As shown in figure 7,

the failures would self-reset upon clearing with no manual reset capability. A built-in test (BIT)

circuit was integrated into the design which sequentially interrupts various actuator control signals

to verify the integrity of the failure-detection circuitry (fig. 8). Because a processor was not

designed into the system, monitoring of the BIT was performed manually in the Mission Control

Center (MCC) before each flight. Although not as efficient as an automatic system, this approach

proved adequate for this system.
(

(

Ram LVDT sum

Is ram LVDT sum

above threshold?

No

Is ram LVDT sum

below threshold?

Yes

Yes

r

_No 1

Clear ram LVDT Set ram LVDT

monitor or fail monitor fail

960405

Figure 5. Ram LVDT monitor.

Hydraulic

pressure status

Is hydraulic pressure / No below 1100 Ib/in2?

Yes

Has pressure been

below 1100 Ib/In 2 for

5 sec?

Yes

Set pressure hydraulic fall

'r

iic'r--u"c pressure fail I

960406

Figure 6. Hydraulic pressure monitor.
Spool position

versus current

test fail 7

LVDT /

sum /

test fall _ _-

Hydraulic /

pressure 1

monitor fail --J

Failure detection board

Power

supply fail --I

 o.r L> t

on reset--,_ I _ Latch I _ To shutoff Se.o,oop --1 I va,vare,sy

fail _ _ r _ I /

! _ _ I " Toserv.ovalve

Servoloop board

Clock 960407

Figure 7. Failure logic.

Initiate switch ] ] ,.16

---_ Counter _J_ -.

Clock --_

Demultiplexer

Figure 8. The ANSER built-in test.

-'_'_ Model command

"_-------- Strake command

_-"_ Spool LVDT high

_-"_ Unused

_-"_-- Servocurrent

_-"_ Ram LVDT high

_'_---- Ram LVDT low

_'_-_- Hydraulic monitor

960408

The SIB is powered by two independent power supply modules which regulate aircraft 28 Vdc

and convert it to +15 Vdc, 5 Vdc, and 8 Vrms LVDT excitation. The LVDT excitation frequencies

are programmable by selecting one of various jumper configurations at the power supply. To eliminate beat frequency development between channels, the two strake channels were given different

LVDT excitation frequencies. As with the servoloop boards, existing FCC power supplies were

used to eliminate development costs. The power supply modules, servoloop boards, and failuredetection boards were housed in a locally manufactured enclosure (fig. 9). The box was mounted

in the radome aft of and between the strake actuators.
Figure 9. Power supply modules, servoloop boards, and failure-detection boards.

The design goal for the strake control system was fail-operate/fail-safe. The fail-safe mode

placed the affected actuator into trail-damped mode, in which the actuator shutoff valves are

opened, causing hydraulic fluid to bypass the main ram. The surface then floats freely, thus not

generating a forcing moment on the aircraft. However, testing showed that the airflow over the

strake did not generate enough force to overcome the breakout force and friction of the strake

actuator and linkage while in trail-damped mode. Thus, the strake was essentially stuck in its last

position. Evaluations of a stuck strake conducted in the DFRC F- 18 simulation showed that an unrecoverable departure could occur at angles of attack of 50 ° or higher with a strake fully deployed.

10
This discoveryled to implementationof apilot-initiatedsystemto bypassthecontrol systemand

forcethestrakesto thefully retractedpositionin theeventof astuckstrake.A switchaddedto the

cockpitprovidedpowerto four independentrelayswhich overrodethe shutoff valvesandcommandedthe actuatorsto fully retract independentof the control systemcommands.Figure 10

showsthis logic.

BIB Strake override switch

Essential 28 V

SOV driver

Command high

Command low

Strake retract switch

Retract

Trail damped

I

GND

I

3900

I

_-_ GND

Figure 10. Strake retract system.

Actuator

I Bypass

valve to

operate

Servovalve

to full

retract

96o409

Cockpit modifications to accommodate the strake system were minimal. In addition to the

strake retract switch, two toggle switches were added to initiate the BIT logic. Two override

switches were also added to manually interrupt the shutoff valve driver signal and place the

actuators in the trail-damped mode. The override switches were only used in ground operations

although no measures were taken to prohibit their functionality in flight. A failure indication was

added to the warning light panel to indicate any failures in the system. Display modifications that

required updates to the mission computer (MC) software included the addition of a mode transition

button and a gain select button to the FCS display on the DDI and a mode and gain set display on

the head-up display (HUD). Figures 11 and 12 show the DDI and HUD symbologies. The onboard

augmented vehicle (OAV) engagement button and maneuver selections button for the onboard

excitation system (OBES) were already present from phase 2 of flight test. The FCS DDI page was

used to select ANSER modes, dial-a-gain settings, OAV maneuvers, and OBES engagement. The

HUD was used to display RFCS engagement, OBES engagement, ANSER modes, and dial-a-gain

settings. No indication of strake position was presented in the cockpit.

11
OAV engage button

MODE select button --_.

GAIN select button --_.

NIG HT-._.._- AUTO

I I I I I

2-[-']

I I 000000 B4k4.-l-l-I

I N 000000 j_ __

I--'I I oooooo _ _ _

I_1 i oooooo _ 7, I_J

k cs °°°°°° y

DF1Flr-IF1/ oo 0

o,vm__u__//

select buttons -_ 98o4_o

Figure 11. Flight control system display.

/

350 000 010

I t I I I

^

sr 15

AOA 0.0 r-- --I

M 0.00

G 0.0

-

F RFCS indication

_'_F OAV Indication

TV[_f MODE indicatiOn

GN 0 _-- Gain set indication

5L ....... J5 --

/ -- \

/ /TI \

<_

960411

Figure 12. The HARV head-up display.

SOFTWARE DESIGN AND IMPLEMENTATION

The ANSER software was developed using two methods: hand coding from a specification and

automatic code generation. The software for the ANSER can be divided into nine elements. These

12
elementsconsistof the existingshell,lateral-directionalcontrollaws,longitudinalcontrol laws,

pseudocontrols,TV mixer(ref.4), modelogic,G-disengage,dial-a-gain,andOBES.This section

discussesthesesoftwareelementsanddesignmethodologiesin moredetailwith theexceptionof

theTV mixer.

The RFCSexecutesat 160Hz in 16minor frames;thelongitudinalcontrol laws executeat

80Hz in the even minor frames;and the lateral-directionalcontrol laws, pseudocontrols,and

TV mixer executeat 80 Hz in the oddminor frames.Eachminor frameis 6.25mseclong with

therequirementthat the RFCScommandsto the processorbe computedby the 2.2-msecmark.

These2.2msecwerethecritical throughputrequirement.Read-onlymemory(ROM) wasalso a

concern.The RFCS had only 32 kilobytes of electrically erasableprogrammableread-only

memory(EEPROM).

Hand-CodedSoftware

TheANSERcontrollaw softwarewasdesignedanddevelopedfrom theHARV ANSERcontrol law specificationwrittenby theVehicleDynamicsandtheDynamicsandControlsBranches

of LaRC.ExistingRFCSsoftwarewasusedasabaselinefortheANSERcontrollaws(ref.5).The

control law softwarewas removed,and the existing shell containingthe multirate structure,

disarm-disengagementlogic,grossthrustestimator,andRFCSI/O softwarewasretained(fig. 13).

TheTV portionof theANSERcontrollawswasimplementedandflown first independentlyfrom

thestrakeinterfaces,followedby thecompleteANSERcontrollaws.

Control

commands,

RFCS no go

Control Control

commands commands

_,_ °ut u, 1 Ic°n'r°'l

logic interface I

I RFCS no go

Control

commands,

RFCS no go

p '

R

I Input

Control disarm/

inputs, analog/_ _-_ disengage

discrete I/O Control logic

inputs,

analog/

discrete I/O

Control[

_---'1 input I

ContrOlinputs,I interface I

analog/

discrete I/O,

computed

values

Figure 13. The RFCS shell.

Control

inputs,

analog/

discrete I/0,

computed

values

RFCS

control

laws

960412

Because of limited memory and throughput of the RFCS, the code was periodically reworked

to reduce the amount of memory it used. To reduce the risk of throughput problems, the software

13
waswrittenasefficiently aspossible.Themultiratestructureonly processedinputsat therateand

frameat whichtheywereupdated.Airdataandinertialnavigationsystem(INS) angleswereinput

at 20 Hz. Angle of attack (AOA) and angleof sidesliprate([3) inertial wereinput at 40 Hz.

Although pilot commandsandrateand accelerationfeedbackswereinput at 40 and 80 Hz, the

40Hz signalswereprocessedat80Hz andaveragedto providecontinuousdatato thecontrollaws.

Only processingrequired to computecommandoutputswas completedin the critical first

2.2msec.All othercomputationswereprocessedaftercompletionof thecommandoutputcomputationbutbeforethe6.25msecendof frame.

Usingthehorizontalblock diagramsprovidedin thecontrol law specification,a control law

input listing wasgeneratedto verify thatthe requiredinputswereavailableto the control laws.

Next, eachaxiswasmodularized,anddataflow pathsweredetermined.Figure14showsanexample of a dataflow pathdiagram.Oncethedataflow wasdetermined,theexecutionorderof the

moduleswaseasilydeterminedbasedon thedataflow. Theexecutionorderensuredthat outputs

werecomputedbeforebeingusedasinputsto otherportionsof the system.Oncethedataflow,

executionorder,computationrates,input availability,andsoftwaremodularityweredetermined,

the softwarewascoded.Thereviewprocessat thecodelevelconsistedof informalcodereviews

and limited modulelevel andend-to-endtesting.The modulelevel andend-to-endtestingwere

performedusingtheRFCSsourcecodeonaworkstation.

Command /Pedal\ Lateral pedal

inputs _._{ command _.........._.._..command

k ,pazn / _ Directional pedal

_ Lateralstick // ANSER _C°ntr°llaw

Command in__ ,./ command _ commands

p_ r summation --p;ih .... /Directional _ ..... J

stick command_

Feedback Inputs >-_k

\ path // signals

960413

Figure 14. Lateral-directional control law data flow.

Automatic Code Generation

In developing the ANSER control laws specification, LaRC extensively used computer

generated software code in both FORTRAN and Ada using MATRIXx .® Using automatic code

®MATRIX X is a registered trademark of Integrated Systems, Inc., Palo Alto, California.

14
generationreducedthe time requiredto makechangesto the control law and to evaluatethe

changesin simulation,thusshorteningthe time requiredbetweencontrol law modification and

flight test.

TheFORTRAN from theautomaticcodegenerationwasusedin theLaRCsimulationsandin

theDFRCbatchsimulation.Whenchangesweremadeto thecontrollaws,softwarewasregeneratedusingtheautomaticcodegeneratorandinterfacedto thesimulations.Check-casescriptsand

time historieswerealsoprovidedfor usein verifyingtheimplementation.Thisprocedureoffered

considerablesavingsin time overthetraditionalprocessof hand-codingthechanges,evaluating

them,andthenprovidingthechangesto programmersin theform of block diagramsto behandcodedin that simulation.Of course,small changes,suchaschangingthe valuesof a few variables,werestill completedby handdirectlyin thecode.

To completethis processof automation,carryingtheautomaticcodegenerationall theway to

flight wasdesired.Someobstacleswerein this path,however.Theautomaticallygeneratedcode

wasnotasefficient ashand-codedsoftwarein termsof executionspeedandmemory.This obstacle was seriousbecausethe RFCSwasoperatingnearcapacityin memory andexecutiontime.

Second,the softwaregenerationtool did not include a multirate capability. In spite of these

obstacles,it wasdecidedto usetheautomaticallygeneratedcodefor thepseudocontrolsportion

of the lateral-directionalcontrol law to run in the RFCS.The automaticallygeneratedcodefor

thepseudocontrolswasmodifiedto accommodatethe differencesbetweenthe simulationsand

theRFCS.

ModeLogic

Becausethe ANSER control lawshad threeseparatecontrol law modes(TV, STV, andS),

allowingin-flight selectionsof thesemodeswasrequired.A modeselectionsystemwas,therefore,

incorporatedinto theRFCS.Thrustvectoringwasselectedasthedefaultmode;therefore,wheneverRFCSwasnot engaged,thecontrolslaws werein theTV mode.In addition,upon RFCS

engagement,the TV modewas selected.Transitionsto the STV andS modeswere performed

throughasinglemodeselectbuttonontheFCSpageontheDDI. A singlebuttonwasusedbecause

of memoryconstraintsin theMC. The pilot usesthis modeselectbuttonto togglethroughthe

modes,TV ---)STV---)S---)TV (fig. 15).ThesemodeswerethendisplayedontheHUD anddownlinkedto thecontrolroom.

Becausea 1-to 2-secfadeexistedbetweenthemodes,theHUD modeindicationwouldbe"X"

sothatthepilot wasawarethata modetransitionwasin progress.TheRFCSdisengagesrequests

throughthe paddleswitch functionedfrom anymode.A strakefailure indication from the SIB

would forceamodetransitionbackto theTV mode.Early in thedesignthedecisionto allow the

pilot to remainin TV following astrakefailurewasmade.TheTV canprovidegreatercontrolpower to recoverfrom abnormalconditionsthat might occurwhenrecoveringfrom a strakefailure.

This failure indicationwaslatchedin theRFCSsoftwaresothatthefailure reasoncouldbedeterminedin the control roombeforethepilot wasallowedto reselecta strakemode.To limit the

15
aircraft to benignflight conditionsduringmodeselections,the listed modetransitionlimits for

TV _ STV andSTV _ Sweredefinedin thesoftwareasdonefor RFCSarming.

Production .

FCS

RFCS engaged

2 eec

TV

t

2 sec 1 sec

STV

2 sec

_'_" Pilot selected mode transition.

(RFCS disengage to production FCS was available from all modes.)

Automatic down mode to TV mode due to strake failure

(strake go = false). (Automatic downmode to TV gives pilot

option to recover from strake failure with TV enabled.) 96o414

Figure 15. The ANSER mode transitions.

Table 1 lists the mode transition limits. The limit on AOA was a place holder in the software

in case a smaller limit was desired. For S ---) TV transitions, these limits were set at or near maximum sensor values to allow the pilot to select the TV mode throughout the RFCS envelope.

Table 1. Mode transition limits.

Signal Limit

Roll rate

Pitch rate

Yaw rate

Lateral acceleration

Normal acceleration

Angle of attack

Sideslip rate inertial

< +30°/sec

< _ 15 °/sec

< _ 15°/sec

< +0.5 g

>-I gor<2g

< +360 °

< +5°/sec

OBES and Dial-A-Gain.

Although not new for ANSER, the OBES was used extensively during ANSER research

flights. The OBES portion of the RFCS software was originally designed as a structural

mode excitation system, but it has evolved into a multidisciplinary, on-aircraft test facility. The

16
growing numberof useful applicationsincludesaerodynamicmeasurementsandflow visualization, parameteridentification, control law evaluation,in-flight control law variability, handling

qualitiesandinnerloop stability analysis,simulationmodelverification,andgroundtestof new

aircraftsystems.

Theinitial setof researchTV controllawsrequiredactivein-flight structuralmodeexcitation.

This requirementresultedfrom marginalaeroservoelasticstabilityanalysispredictions.Different

excitationconceptswereconsidered.An extemalmechanicalexcitationsystemwasnot selected

becauseof its intrusivenatureandintegrationdifficulties.Another concept considered was to use

the remotely augmented vehicle (RAV) facility. The RAV facility has been a critical test facility

at DFRC for remote maneuver guidance, control augmentation, and remote piloting. This facility

could be used to up-link an excitation signal to the aircraft which would then need to be received,

processed and passed to the RFCS. The signal would then be summed to the desired control surface

command. By driving the selected control surfaces at varying frequencies, the structural modes of

the aircraft could be excited. The frequencies of the structural modes of interest were well within

the bandwidth of the F- 18 actuators.

The OBES was created by adapting this latter concept. Taking advantage of the imbedded and

programmable RFCS, the functions of the RAV facility were effectively brought onboard the aircraft. With this OBES configuration, the RAV functions were performed internal to the aircraft,

thus eliminating the problems associated with the remote facility. Some of the problems addressed

include the up-link and down-link range and dropout problems, on-aircraft receiver requirements,

signal processing, redundancy management, and time delay. Other RAV drawbacks eliminated by

the OBES include the need for manpower support during the flights as well as dependence on a

complex system of hardware and software.

The OBES generates the desired function and sums the signal to the desired RFCS control law

path (fig. 16). The function generation logic and signal superposition are performed within the

RFCS Ada software, and the selected OBES function identifier is down-linked to the control room

for test point verification. For these flight test points, the test pilot would simply initiate the appropriate OBES maneuver and maintain flight conditions. The OBES performs the selected maneuver.

Lateral stick

command

OBES lateral

stick command

Lateral stick

deadband

OBES rudder

command

Rudder

command

___ Rudder limit

960415

Figure 16. The OBES command summation examples.

17
ForANSERflight testing,theOBESwasusedfor nosestrakeaerodynamicmeasurementsand

flow visualization,nosestrakeloadsflight clearance,closed-andopen-loopparameteridentification andnosestrakepreflight built-in testexcitation.

A dial-a-gaincapability wasaddedto the RFCSto add flexibility to the flight test of the

ANSERcontrol laws. Theoption of eight separategainsetselectionswasmadeavailable.For

ANSER, threegain setswere definedfor the longitudinalaxis with onegain set definedasa

default,baselineset.ThesegainsetswereselectedthroughthecockpitDDI, andthecurrentgain

setwasdisplayedontheHUD anddown-linkedfor controlroommonitoring.Gainselectionwas

allowed throughoutthe flight envelopebefore or after RFCSengagement;however,with any

RFCSdisengagement,heselectedgainsetdefaultedbackto thebaselinegainset.

g-Disengage

The RFCS and Ada software were originally intended to be a class B system to reduce the

software testing and verification requirements (ref. 3). Although many definitions of class B exist,

in general this designation meant that an error in the RFCS or software could not result in loss

of aircraft.

Vehicle modifications required by the installation the TVCS and ANSER systems added a

significant amount of weight at the extreme fore and aft ends of the aircraft. Because of this "flying

dumbbell" configuration, expected load levels in the forward fuselage area during maneuvering

flight were significantly increased. Structural analysis indicated that a reduction in symmetric and

asymmetric maneuvering load limits was required (table 2). Software modifications to enforce

the new limits could not be made to the basic flight control system, so a flight operations limit

was imposed.

Table 2. Production F/A-18 and HARV load limits.

Load limit Production F/A- 18 HARV

Symmetric -3.0 g to 7.6 g -2.0 g to 5.4 g

Asymmetric 0 g to 6.0 g 0 g to 4.2 g

Although the flight operations limit specifically restricts the pilot from maneuvering above the

symmetric and asymmetric load limits, the aircraft still has the command authority and control

power available to significantly exceed the aircraft structural limits in the lower right-hand corner

of the RFCS envelope (h = 15,000 ft, M = 0.7). During software testing, it was found that hard-over

stabilator commands from the RFCS could not be countered by the pilot in enough time to prevent

an over-g. As a result to balance the risk of an undetected RFCS software fault with the general

desire to design and test in a class B environment, a function to predict and reduce the possibility

of an over-g was implemented in the RFCS.

18
This function, calledthe g-disengage, was only active at high dynamic pressures where an

over-g was possible. Structural loads in the critical forward fuselage area were estimated from

flight control accelerations and rates and projected ahead in time with a lead filter. When the projected load exceeded the computed load limits, the RFCS stayed engaged. However, the RFCS stabilator command was ignored, and the stabilator was driven to the basic F/A-18 stabilator

command. Once the stabilator got within 5 ° of the basic system stabilator command, the RFCS was

disengaged. In this way, the g-disengage reduced the potential of an over-g caused by a failure in

the RFCS. It dynamically limited the stabilator commands.

Once the g-disengage software was implemented, it was complied and stored in a designated

portion of the EEPROM. This designation was done so that when software updates were made,

a bit-for-bit comparison could be completed using the previous version of g-disengage. This

comparison ensured that any software updates to the RFCS did not affect the g-disengage portion

of the software.

SYSTEM INTEGRATION AND TESTING

The ANSER system was extensively tested before it was integrated into the aircraft. Testing

was broken into four levels: card level testing, box level, integrated system, and system validation.

The designer performed card level testing to check the design and build of the card. Each card was

tested against the design as much as possible. The chassis wiring was also tested at this level.

Box level testing was a more intensive test and was integrated with an actuator and an actuator

model. Integrated system testing included the FCS and the actuators or actuator models. System

validation testing included the actuators or actuator models, FCC, and closed-loop simulation of

the vehicle dynamics.

An actuator model was developed for the integration and testing of the strake system. The

strake actuator model was designed to accurately model the response of the forebody strake actuators in software. The model receives strake commands as inputs from the SIB and hinge moment

from the simulation, and this mode gives out the spool and ram positions, voltages, and ram velocity as outputs to the SIB and then as outputs to the simulation. The actuator model calculates spool

and ram positions in normal operational mode or in trail-damped mode. The models are implemented on a Versa Model Eurocard (VME) system which operates at 800 Hz.

Box level testing checked the box and actuator interfaces. A function generator replaced the

FCC and provided the command voltages to the interface box. The initial tests verified the interface

box under steady-state conditions. A voltage was applied, and the actuator or actuator model position was verified. Once the steady-state inputs were evaluated, a full-scale ramp input was then

used to evaluate the time history of the interface box and of the actuator and actuator model. After

the ramps, a frequency response test was run to verify the dynamics of the servoloop. Failures were

then introduced while a ramp was input to verify that the failure-detection circuits operated as

expected. Table 3 indicates the failures that were introduced and in what level of testing these failures were performed. After the failure-detection tests were run, BIT was tested, and failures were

introduced to verify that the implementation detected the failures properly.

19
Table3.Integrationtestingmatrix.

Failures

Card Box Integrated

level testing system

System

validation

Ram position,

high open

Yes One channel One channel

Two channels with actuator Two channels

and actuator model

One channel

Two channels

Ram position,

low open

Yes No No One channel

Two channels

Spool position,

high open

Yes One channel One channel

Two channels with actuator Two channels

and actuator model

One channel

Two channels

Spool position,

low open

Yes No No One channel

Two channels

FCC command

to strake

interface box,

open

Yes One channel One channel

Two channels with actuator Two channels

and actuator model

One channel

Two channels

Strake interface

command to

actuator or

actuator

model, open

No One channel One channel

Two channels with actuator Two channels

and actuator model

One channel

Two channels

LVDT excitation,

open

No One channel One channel

Two channels with actuator Two channels

and actuator model

One channel

Two channels

Pressure discrete,

open

Yes One channel One channel

Two channels with actuator Two channels

and actuator model

One channel

Two channels

Power, open Yes One channel One channel

Two channels with actuator Two channels

and actuator model

One channel

Two channels

Shut off valve, Yes One channel One channel One channel

open Two channels with actuator Two channels Two channels

and actuator model

Servocurrent, Yes One channel One channel One channel

open Two channels with actuator Two channels Two channels

and actuator model

20
Threemajoranomalieswerediscoveredduringbox leveltesting.First, a groundingproblem

requiredthe additionof a second1.832-MHzclock insteadof relying ona singleclock to drive

bothservoloopboards.Second,asignerrorwasdiscoveredin thespoolpositionversuscommand

currentalgorithm, requiring minor redesign.Third, the counterusedto stepthrough the BIT

sequencewouldintermittentlyresetorarbitrarilyjump to adifferenttest.Troubleshootingrevealed

asharedgroundbetweenthedigitalBIT circuitry andtheanalogcircuitry onthefailure-detection

board.This problemwasdocumentedbut wasdeemeda nuisance.No fix was implemented,so

theproblemwascarriedthroughto theaircraftandoftenrequiredmultipleBIT operationsbefore

eachflight.

Integratedsystemtestingwasthe nextlevel. The RFCSprovidedthe ability to programthe

OBESto generateidenticalrampsandsquarewavesasthosefromthefunctiongenerator.Thetests

which hadpreviouslybeenperformedduringbox leveltestingwere repeatedwith the interface

box, actuatoror actuatormodel,andFCC integrated.The only discrepancynotedduringsystem

integrationtestingwasalackof fidelity in thestrakeactuatormodel.Thismodelwasimprovedand

testingcontinued.

Uponcompletionof integratedsystemtesting,systemlevelvalidationwasrun.A high-fidelity

simulationof thevehicledynamicswas integratedwith theinterfacebox, FCC, andactuatorsor

actuatormodels.The flight softwarewas installedin the RFCS,andtestingsimilar to that performed in the integratedsystemstesting wasperformed(table 3). Time histories, frequency

responses,and inducedfailureswere run at variousflight conditions.The MC modifications,

RFCS,andtotal systemwerevalidated.Thisvalidationincludeddisplaysandswitchologyaswell

asfailures,modetransitions,timehistories,andfrequencyresponses.Duringvalidationtesting,a

singlefailureintheFCCcommandto theSIBwouldresultin theactuatortryingto respondtoboth

commandssimultaneously.Both channels'servovalvecommandsversusspoolpositionmonitors

wouldfail, placingtheactuatorin trail-dampedmode.Thissinglepointfailurecombinedwith the

discoverythattrail-dampedisanunacceptablefailuremode,resultedindevelopingandimplementatingthestrakeretractsystem.

Additionally certainconditions,suchasa singlefailure in a main ram LVDT signal,could

causea failure in the secondchannel.Suchfailuresalso resultedin the trail-dampedmode.

Troubleshootingrevealedthat the failure in the first channelram positiondrove the command

currentin thepartnerchannelhighermomentarilyto compensate.The servopositionversuscommandcurrentmonitorfailurediscretewouldmomentarilytrip andthenself-reset.This transition

typically lasted20to 25msec.Theopticalrelaywhichcontrolstheactuatorshutoffvalveis specified to recoverfrom transitionsin lessthan50msec.Thatrelaywouldlatchto a failed stateand

couldnotberesetbecauseof thedesigndecisionnotto includea manualreset.

Aircraft integrationtestingwas fairly limited becauseof high confidencein the simulation

implementation.Testingconsistedof time historiesin thethreeRFCSmodes.Openwire failures

were inducedto verify failure-detection,andcockpit functionswereverified. The only known

anomalywastheBIT problemdiscoveredduringintegrationtesting.

21
RFCSVERIFICATION

With theimplementationof theg-disengage software, the class B designation of the rest of the

RFCS software was used to define the level of verification testing required before flight. Independent testing has shown that the RFCS envelope limits and the OBES are the only elements of RFCS

which affect flight safety.* Since those tests, the vane load limiting was also determined to affect

flight safety. With this affect in mind, the ANSER verification testing requirements were defined.

The RFCS envelope limits (15,000 ft < h < 45,000 ft, M < 0.7) were verified. A bit-for-bit comparison of the g-disengage software was performed to verify that g-disengage executable did not

change from a previously verified version. The RFCS software configuration identifier was verified to ensure that the software release identification was correct. Thrust-vectoring vane system

(TVVS) command limiters were verified in the lower right- and left-hand corners of the envelope.

Because the OBES could also affect safety of flight, each OBES maneuver was flown at flight conditions starting in the lower right-hand corner of the envelope to determine where in the flight envelope the individual maneuver affected flight safety. Once this determination was completed, a

flight operations limit was defined for that maneuver. Although some informal testing was performed to gain confidence in the operation of the software, no other RFCS software testing was

required. Any errors existing in the software were considered to be only mission critical. The

project decided to accept this risk knowing that a complete verification would take an unreasonable

amount of time.

FLIGHT TEST MONITORING

Flight test monitoring of the RFCS and the processor FCS for phase 2 was done on two control

room display pages. An FCS display page was designed to look similar to the FCS DDI display in

the aircraft (fig. 17). This page displays leading-edge flap, trailing-edge flap, aileron, rudder, stabilator and TVV position and failure status; lateral stick, longitudinal stick and rudder pedal positions, and failure status; and FCS analog and discrete I/O failure status. The RFCS display page

was used to monitor the RFCS I/O and system functionality during RFCS engagement (fig. 18).

Aircraft rates, accelerations, AOA, and rates of change in angles of attack and sideslip, 6t and _,

were displayed with RFCS arm and RFCS engage-disengage limit windows. The TVVS commands and positions were displayed with maximum limits. The RFCS Mach number and altitude

were displayed with RFCS envelope limits incorporated. Engine parameters included power lever

angle (PLA), nozzle area, turbine discharge pressure, and gross thrust. The RFCS abort and failto-arm flags were also displayed. The TVVS temperatures and aircraft loads as a function of percent of maximum allowable were also displayed to allow for the option of not having a structural

loads engineer in the control room.

*Earls, M.R., "Nasty RFCS Test Report," HA92-84-601, 1992. (This report is an unpublished working paper and

is not available to the public.)

22
Left Right

1234 1234

Lateral stick position

Longitudinal stick position

Rudder pedal force

0

0

-4

-10

-10

-10

Leading-edge flap 0

Trailing-edge flap 0

Aileron 0

Rudder 0

Stabilator - 4

Top vane - 10

Outboard vane - 10

Inboard vane - 10

Pitch rate/acceleration

Roll rate

Yaw rate/acceleration

Stick

Pedal

AOA

Airdata

Processor

Degrade

Figure 17. The HARV RFCS monitoring page.

System failures

960416

RFCS ARM RFCS ENGAGE MODE RFCS FAIL TO ARM RFCS ABORT FCS WOW

Temperatures

Left top Cmd

vane Pos

Left out Cmd

vane Poe

LoadsI I Left

inboard Cmd

Pea

vane

Pitch rate II I I 11 Right top Cmd

vane Poe

Roll rate II I I [I

Yaw rate II I I II Right out Cmd

vane Pos

AOA EL .U

Right Cmd

inboard Poe AOA rate 11 II vane

13rate II l I II Mach

Nz II I I I I number

Altitude

Ny II II II

I I I

I I I I

Lt PLA

Rt

Lt A8

Rt

Turbine Lt

Discharge

Pressure Rt

thrust Gross Lt Rt I

Left Cmd

etrake Pos

Right Cmd

strake Pos

Strake go/no go

Trail damped

ORES

ORES index

Gain set

RFCS abort flags

Figure 18. The RFCS monitoring page.

II

960417

23
In addition to the existing RFCS and FCS display pages used on phase 2, other parameters were

incorporated into display pages for the phase 3 (ANSER) portion of flight test. A new ANSER

parameter display system (PDS) display page was designed for monitoring strake actuator commands versus actuator positions, servocurrents, spool positions, and failure-detection circuitry during flight (fig. 19). A BIT index was displayed for preflight BIT monitoring. The strake go flag,

ANSER mode (TV, STV, S), trail-damped indication, FCS caution, and strake retract switch position were also displayed. The FCS caution was displayed to alert the systems engineer that an FCS

problem existed. The systems engineer could then display the FCS display page to troubleshoot the

FCS problem. The RFCS display page was modified for phase 3 to include ANSER mode, strake

command, strake position, strake go flag, and current gain set selected. To indicate an ANSER

mode transition (for example, TV _ STV) to the systems engineer, the ANSER mode would flash

at approximately 5 Hz during the transition (figs. 18-19). Some of the more critical parameters,

such as ANSER mode, trail damped, and strake go, were displayed on the ANSER PDS and the

RFCS display pages. This combination display reduced the systems engineer's scan during flight

test when critical calls to the test pilot, flight controller, or both, were required. To isolate possible

RFCS anomalies from SIB anomalies, the RFCS page displayed command and position data from

the RFCS, and the ANSER PDS page displayed command and position data from the SIB.

MODE TD FCS RETRACT

Left command channel A

Left ram position channel A

Left command channel B

Left ram position channel B

[

Left servocurrent local channel A

[

Left spool position channel A

Left servocurrent local channel B

Left spool position channel B

Ram LVDT sum channel A

Ram LVDT sum channel B

Shutoff Local Logic Servo- Sen/ovalve sen/oemp fall poaitlonl amp

on off current fal/

fail

Channel A bit index

LVDT HydrllullcHydraulic

sum pressure pressure

monitor fall off

fail

Power Power

supply A supply

Power A

on reset failure

Right command channel A

Right ram position channel A

Right command channel B

Right ram position channel B

I

Right servocurrent local channel A

I

Right spool position channel A

I

Right servocurrent local channel B

Right spool position channel B

Ram LVDT sum channel A

Ram LVOT sum channel B Strike go/no go

Shutoff Local Logic Sen/o- Servo- LVDT Hydraulic Hydraulic

valve servoemp fall position/ amp sum pressure pressure

on off current fail monitor fail off

fail fail

Power Power

supply B supply

Power B

on reset failure

Figure 19. The ANSER PDS page.

Channel B bit index

960418

24
FLIGHT TEST RESULTS

The ANSER system was installed and flown for the final 89 flights of the HARV flight

program. With one exception, flight performance of the system was generally successful. The

servoposition versus command current test would fail and self-reset during large amplitude lateraldirectional maneuvers in the 30 ° to 50 ° angle-of-attack range. Typically, one channel of one actuator would indicate a failure. Troubleshooting revealed a divergence in the RFCS command to the

servoloop between the two channels of the affected actuator. The two independent servoloops

would command a different servovalve position based on the FCC commands. The servovalve

would respond to a position midway between the two commands. When the commands diverged

sufficiently, the failure-detection algorithm would declare a failure (fig. 20). The test tolerance was

increased to allow the commands to differ by 8.5 ° or nearly 10 percent of full scale. This increase

reduced the frequency of the failures, but it did not eliminate them completely.

Left atrake

command,

inches of

stroke

2.0

1.5

1.0

.5

-.5

Channel A command

Channel B command i'_

Strske fail I

-- /I I

|

_ _,'

,\

I I I I I

.2 .4 .6 .8 1.0

Time, sec 960419

Figure 20. Example of in-flight command miscomparison.

Investigations into the cause of the divergent FCC commands are ongoing. The RFCS channels

appear to be receiving different values for lateral stick commands resulting from the quad signal

selection algorithm in the FCC. The high gain nature of the RFCS control laws amplifies this difference and results in the different strake commands. Documentation of these findings is hindered

somewhat because internal FCC data are only available on two of the four channels through the

MIL-STD-1553 data bus. Project scheduling concerns prohibited a resolution of this problem;

therefore, a set of research data in the flight regime affected was not obtained. The probable fix

would have been a persistence counter in the RFCS on the strake go discrete. These failures were

25
generallyone or two frames in length. A five-flame persistence on the strake go discrete would

have allowed the S or STV mode to remain engaged while the failure self-reset.

Other problems encountered in the flight program were minimal. An integrated circuit failure

occurred causing loss of one flight for repair. A failed spike suppression diode in an unrelated

system caused corruption of strake power. This diode failure resulted in an self-start of the ANSER

BIT logic in flight. The retract system was used, and the aircraft was recovered normally.

LESSONS LEARNED

Several lessons were learned in the development of the ANSER system. These lessons can be

applied to future system integration efforts. An assumption was made early in the program that

placing the actuators to trail-damped was an acceptable failure mode. This assumption drove a

great deal of the system design and eventually proved false. A substantial effort was then required

to implement the strake retract system. Had the assumption been disproved earlier in the design

phase, the normal failure mode could have been to a strake-closed configuration.

The decision to have failures automatically reset was adequate in the majority of cases, but

when a main ram LVDT failure occurred, the partner channel servocurrent versus spool position

monitor failed and reset faster than the electronics could recover. This rapid automatic reset resulted in an unresettable trail-damped condition. The addition of a manual reset capability along with

the automatic reset would have alleviated this problem.

Monitoring the preflight built-in test manually rather than automatically proved to be a valid

approach for this application. This monitoring eliminated the development of a processor-based

system to perform this function.

A persistence counter on the strake go discrete may have prevented the reversions to TV mode

in the 30 ° to 50 ° angle-of-attack range during the flight phase. Additionally, a cross-channel comparison of RFCS parameters may have detected the command miscomparisons, and a routine could

have been implemented to average these signals.

In the RFCS area, increased throughput and ROM would have eased the software design.

Streamlining the code a great deal was necessary to accommodate these limitations of the RFCS.

CONCLUDING REMARKS

A system to enhance high-angle-of-attack flight handling qualities was developed by NASA

Langley Research Center and integrated into the F-18 High Alpha Research Vehicle by NASA

Dryden Flight Research Center. The actuated nose strake for enhanced rolling system proved reliable throughout 89 test flights. This system consists of two actuated control surfaces installed on a

specifically constructed radome.

26
Actuatorcontrolwasperformedbyastrakeinterfaceboxwithcommandsgeneratedbythecontrol lawsresidingin theresearchflight controlsystem(RFCS).Thesecontrol laws consistedof

threemodes:thrust vectoring(TV), strakesandthrustvectoring(STV), andstrake(S). TheTV

modeusesthe TV systemfor longitudinal andlateral-directionalcontrol.The strakesarenot

commandedin thismode.Thesystemrevertsto theTV modein theeventof astrakefailure.The

STV modeusesTV for longitudinalcontrol anda combinationof strakeswith TV for lateraldirectionalcontrol.TheSmodeusesTV forlongitudinalcontrolandstrakesfor lateral-directional

control.TheRFCSalsoresidesin theonboardexcitationsystemandallowsopen-andclosed-loop

functionsto besummedinto theRFCScontrollawsfor researchflexibility.

Extensivetestingwasperformedto qualifythesystemfor flight. Thistestingrevealedonemajor andseveralminorsystemproblems.Themajorproblemwasthatthefailuremodefor the systemwhichhadbeendesignedto placethefailedstrakesin trail-dampedmodewasinadequate.This

inadequacycouldresultin a out-of-controlsituation.Thisproblemresultedin theimplementation

of a manuallyactivatedretractsystemto closethestrakesif atrail-dampedsituationoccurred.

In the30° to 50° angle-of-attackrange,strakecommandswoulddiffer betweenthetwo flight

controlchannelsdriving aparticularstrake.Thisdifferenceresultedin areversionto TV mode.A

correctionwasnotimplemented,andaportionof theflight researchprogramwasnotcompleted

becauseof thisfailure.With thisexception,theprogramwasgenerallysuccessful,andthemajority

of theresearchobjectivesweremet.

REFERENCES

1Gera,Joseph,"Flight TestStatusof theNASA High-Angle-of-AttackTechnologyProgram,"

High-Angle-of-Attack Projects and Technology Conference Proceedings, NASA CP-3137, 1992.

2Murri, D.G., Shah, G.H., and DiCarlo, D.J., "Preparations for Flight Research to Evaluate

Actuated Forebody Strakes on the F-18 High-Alpha-Research Vehicle," Fourth High-Angle-ofAttack Technology Conference Proceedings, NASA CP- 10143, 1994.

3Chacon, Vince, Pahle, Joseph W., and Regenie, Victoria A., Validation of the F-18 High Alpha

Research Vehicle Flight Control and Avionics System Modifications, NASA TM- 101723, 1990.

4Bundick, W. Thomas, Pahle, Joseph W., Yeager, Jessie C., and Beissner, Fred L., Jr., Design

of a Mixer for the Thrust-Vectoring System on the High-Alpha Research Vehicle, NASA

TM-110228, 1996.

5Regenie, Victoria A., Earls, Michael, Le, Jeanette, and Thomson, Michael, Experience with

Ada on the F-18 High Alpha Research Vehicle Flight Test Program, NASA TM-104259, 1992.

27
REPORT DOCUMENTATION PAGE FormApproved

OMB No. 0704-0188

Public reporting burden _or this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources,

gathering and maintaining the data needed, and cornpleting and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 Jefferson Davis

Highway, Suite 1204, Arlington, VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington, DE: 20503

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

October 1996 Technical Memorandum

4.TITLE AND SUBTITLE

Design and Integration of an Actuated Nose Strake Control System

6. AUTHOR(S)

Bradley C. Flick, Michael P. Thomson, Victoria A. Regenie, Keith D.

Wichman, Joseph W. Pahle, Michael R. Earls

7.PERFORMINGORGANIZATIONNAME(S) ANDADDRESS(ES)

NASA Dryden Flight Research Center

P.O. Box 273

Edwards, California 93523-0273

9.SPONSORING/MONITORING AGENCYNAME(S)ANDADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-3191

5. FUNDING NUMBERS

WU 505-68-30

8. PERFORMING ORGANIZATION

REPORT NUMBER

H-2134

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM- 104324

11. SUPPLEMENTARY NOTES

Presented at the 5th High Angle of Attack Technology Conference, NASA Langley Research Center,

Hampton, Virginia, Sept. 17-19, 1996.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified--Unlimited

Subject Category 05

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

Aircraft flight characteristics at high angles of attack can be improved by controlling vortices shed

from the nose. These characteristics have been investigated with the integration of the actuated nose

strakes for enhanced rolling (ANSER) control system into the NASA F-18 High Alpha Research

Vehicle. Several hardware and software systems were developed to enable performance of the

research goals. A strake interface box was developed to perform actuator control and failure detection

outside the flight control computer. A three-mode ANSER control law was developed and installed in

the Research Flight Control System. The thrust-vectoring mode does not command the strakes. The

strakes and thrust-vectoring mode uses a combination of thrust vectoring and strakes for lateraldirectional control, and strake mode uses strakes only for lateral-directional control. The system was

integrated and tested in the Dryden Flight Research Center (DFRC) simulation for testing before

installation in the aircraft. Performance of the ANSER system was monitored in real time during the

89-flight ANSER flight test program in the DFRC Mission Control Center. One discrepancy resulted

in a set of research data not being obtained. The experiment was otherwise considered a success with

the majority of the research objectives being met.

14. SUBJECTTERMS

Ada, Digital systems, F-18 High Alpha Research Vehicle, Flight control, Nose

strakes, Servoloop, Servoactuators

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

OF REPORT OFTHIS PAGE OF ABSTRACT

Unclassified Unclassified Unclassified

NSN 7540-01-280-5500 Avai/ab/e from the NASA Center for AeroSpace Information, 800 Etkridge Landing Road,

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

15. NUMBER OF PAGES

31

16. PRICE CODE

AO3

20. LIMITATION OF ABSTRACT

Unlimited

Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std Z39-18

298-102