NASA Technical Memorandum 4771

Thrust Vectoring on the

NASA F-18 High Alpha

Research Vehicle

Albion H. Bowers and Joseph W. Pahle

November 1996

NASA Technical Memorandum 4771

Thrust Vectoring on the

NASA F- 18 High Alpha

Research Vehicle

Albion H. Bowers and Joseph W. Pahle

Dryden Flight Research Center

Edwards, California

National Aeronautics and

Space Administration

Office of Management

Scientific and Technical

Information Program

1996

THRUST VECTORING ON THE NASA F-18 HIGH ALPHA RESEARCH VEHICLE

Albion H. Bowers

Joseph W. Pahle

NASA Dryden Flight Research Center

Edwards, CA

ABSTRACT

Investigations into a multiaxis thrust-vectoring system have been conducted on an F-18

configuration. These investigations include ground-based scale-model tests, ground-based fullscale testing, and flight testing. This thrust-vectoring system has been tested on the NASA F-18

High Alpha Research Vehicle (HARV). The system provides thrust vectoring in pitch and yaw

axes. Ground-based subscale test data have been gathered as background to the flight phase of the

program. Tests investigated aerodynamic interaction and vane control effectiveness. The groundbased full-scale data were gathered from static engine runs with image analysis to determine

relative thrust-yectoring effectiveness. Flight tests have been conducted at the NASA Dryden

Flight Research Center. Parameter identification input techniques have been developed.

Individual vanes were not directly controlled because of a mixer-predictor function built into the

flight control laws. Combined effects of the vanes have been measured in flight and compared to

combined effects of the vanes as predicted by the cold-jet test data. Very good agreement has

been found in the linearized effectiveness derivatives.

NOMENCLATURE

Axial

b

c

C dpv

C dyv

CAaero

CDaero

Ct

Cm

Cmaero

total axial force, lbf

span, fl

reference chord, fl

linearized pitch vane effectiveness

linearized yaw vane effectiveness

aerodynamic axial force coefficient

aerodynamic drag force coefficient

aerodynamic lift force coefficient

aerodynamic rolling moment coefficient

pitching moment

aerodynamic pitching moment coefficient
CNaero

Cnaero

Ct

CY.ero

dl

d2

d3

d4

d5

d6

apv

dyv

f(x)

HARV

l

LN

NASA

Normal

NPR

Pt

P_

pitch

RFCS

roll

S

Side

TVCS

V

vane 2angle

vane Jangle

aerodynamic normal force coefficient

aerodynamic yawing moment coefficient

thrust force coefficient

aerodynamic side force coefficient

upper left vane deflection, deg

outer left vane deflection, deg

lower left vane deflection, deg

upper right vane deflection, deg

outer right vane deflection, deg

lower right vane deflection, deg

generalized average pitch vanes deflection, deg

generalized average yaw vanes deflection, deg

net propulsive force, lbf

arbitrary function of variable x

High Alpha Research Vehicle

length of moment arm from center of gravity to vanes, ft

total lift force, lbf

National Aeronautics and Space Administration

total normal force, lbf

nozzle pressure ratio, Pt Ip**

jet total pressure, lbf/ft 2

ambient pressure, lbf/ft 2

total pitching moment, ft.lbf

research flight control system

total roll moment, ft/lbf

reference area, ft 2

total side force, lbf

thrust-vectoring control system

velocity, ft/sec

radial vane angle for vane 2, deg

radial vane angle for vane 3, deg

2
yaw

£t

8j.._

6tef

6r

8t_f

6v

P

total yaw moment, ft/lbf

angle of attack, deg

angle of side slip, deg

aileron deflection, deg

elevator deflection, deg

jet-turning angle, deg

leading-edge flap deflection, deg

rudder deflection, deg

trailing-edge flap deflection, deg

vane deflection, deg

density of air, slug/ft 3

INTRODUCTION

A strong interest in thrust vectoring 1,2 has led to many experiments designed to incorporate

thrust vectoring into current- and next-generation aircraft. Most of these studies have focused on

vectoring in the pitch ]?lane to improve the pitch control power 3,4 or in the yaw plane to improve

yaw control power. 5-_ Some studies have examined vectoring in the pitch and yaw planes. 8-10

To date, only a few research aircraft have flown with the capability of vectoring in both pitch

and yaw. Two of these aircraft have completed research programs using axisymmetric nozzles with

postexit vanes for thrust vectoring. These two aircraft are the U. S. Navy X-31A 11 and the NASA

F-18 High Alpha Research Vehicle (HARV) with a thrust-vectoring control system (TVCS)

installed. 12-15 Based on previous developmental tests, 8 both aircraft selected thrust-vectoring

systems that employed three postexit vanes radially displaced about their axisymmetric nozzles.

Early information was required to properly characterize an axisymmetric nozzle with postexit

exhaust vanes like that applied to the NASA F-18 HARV. These data were provided by testing

ground-based models of the F-18 HARV with the TVCS installed. Two aerodynamic interaction

tests were performed at the NASA Langley Research Center. The first test was conducted in the

Full-Scale Facility 30- by 60-ft wind tunnel. The test characterized the aerodynamic interaction

effects on a full-configuration, 16 percent-scale F/A-18 model that were caused by vectoring the

exhaust plume at low speeds. 16 The second test was conducted in the 16-Ft Transonic Tunnel

using a wingtip-supported 10 percent-scale F/A-18 model. This test was conducted between

Mach 0.3 and 0.7 and angles of attack between 0 ° and 70 °, with military-power and fullafterburner nozzles at nozzle pressure ratios (NPRs) between 1.0 and 5.0,17 to further examine

aerodynamic interaction effects caused by vectoring the exhaust plume.

A static cold-jet test was also conducted at the NASA Langley 16-Ft Transonic Tunnel to

characterize the thrust-turning effectiveness of the postexit vanes on a 14.25 percent-scale nozzle

3
model.18,19Thistestwasonly of thenozzleandmodeledonly theinternalflow of thenozzleand

thevanes.

A static ground test,called the "hot loads"test, wasconductedwith the aircraft after the

TVCS hadbeeninstalledat theNASA DrydenFlight ResearchCenter.The aircraftwassecured

to the EdwardsAir ForceBaseUniversal Horizontal ThrustStand.The thrust standwasonly

capableof resolvingforcesin theaxialdirection,andcomparisonsof theseresultswith thecoldjet test axial thrust lossdatawere made.19Infrared imagesof the exhaustflow were usedto

estimateplume-deflectionangles.Hot loadstestplume-deflectioninfrared estimateswerealso

comparedto cold-jettestresults.

Flight testswerealsoconductedat NASA Dryden.In-flight plumedeflectionis difficult to

measuredirectly, soindirectmethodsto estimatethrust-vectoringeffectivenesswereused.20-22

Parameteridentificationinput techniquesweredeveloped,23bothpilot-flown andpreprogrammed

throughthe useof anon-boardexcitationsystemthat controlledthe postexitvanes.Individual

vaneswere not directly controlled becauseof a mixer-predictorfunction built into the flight

control laws.As such,combinedeffectsof the vanesweremeasuredin flight andcomparedto

combinedeffectsof thevanesaspredictedby thecold-jettestdata.

This paperpresentsanoverviewof the ground-basedtests,resultsof the infrareddataof the

hot loadstestof theinstalledTVCS,andestimatesof theeffectivenessof theTVCSfrom in-flight

parameteridentification.Thesefull-scaleresultsarecomparedto the subscaletests.All testsof

this systemarenow complete,andthisreportsummarizessignificantresultsof thefull-scaletests

in comparisonto predictedthrust-vectoringcontroleffectiveness.

HIGH ALPHA RESEARCHVEHICLEDESCRIPTION

TheF-18 HARV is ahigh-performancetwin-turbofanjet enginefighter-attackairplanebuilt

by McDonnell DouglasAerospace(St.Louis, Missouri) (fig. 1).12-15The HARV hasmidwingconfiguration wings, twin vertical tails cantedoutwardsat 20° from the vertical, midmounted

horizontaltailplanes,andleading-edgeextensionsthatrunforward alongthefuselageto nearthe

canopy.The enginesareGeneralElectric (Lynn, Massachusetts)F404-GE-400enginesratedat

10,000lbf staticthrustat sealevelin thefull military-powersettingand16,000Ibf staticthrustat

sealevelin themaximum-afterburnersetting.

The aircraft wasmodified by the addition of a rudimentaryTVCS that usesthreepostexit

vanesaroundeachengine.The geometry of the TVCS hardware uses three vanes mounted around

each engine of the F-18 airplane. Vanes replace the standard divergent section of the nozzle and

external flaps. The convergent section of the nozzle remains on the aircraft. The vanes were

designed to be stowed at -10 °, out of the exhaust plume. During vectoring maneuvers, a

maximum of two vanes on any one engine were commanded in contact with the flow at one time,

helping to alleviate thermal and load constraints on the vane system.

The thrust-vectoring vanes were controlled through use of modified flight control computers

that allowed the basic flight control laws to be bypassed with research flight control laws in a
960689

Figure 1. Three-view drawing of the F-18 HARV.

portion of the flight control computers called the research flight control system (RFCS). The

RFCS envelope of the aircraft was Mach 0.2--0.7 and initial altitudes of 15,000-35,000 ft, which

was later expanded to 45,000 ft. The basic F-18 flight control laws resided in the 701E (General

Electric, Lynn, Massachusetts) portion of the flight control computers and passed data back and

forth to the RFCS.

The output of the control laws was a simple command of pitch or yaw vectoring moments. To

make use of these commanded moments from the thrust vectoring, a schedule was devised, called
the mixer-predictor or the mixer, wherethe pitch and yaw commandswere converted into

individual vanedeflections.Severalreleasesof mixer softwareweremade,andtwo werecarried

to flight (Mixer 1andMixer 4.2).Mixer 1wasa cooperativeeffortbetweenMcDonnellDouglas

Aerospace,NASA Dryden,andNASA Langley.24Mixer 4.2 wasa cooperativedesignbetween

NASA LangleyandNASA Dryden.25

The ability to usepreprogrammedmaneuversby superimposingmotions over the control

surfaceoutputs,called the on-boardexcitation system,residedwithin one part of the RFCS.

Theability to separatethe individual control surfacemotions throughthe on-boardexcitation

systemallowed reductionin surfacemotion correlation andreduceduncertainty in the final

effectivenessestimates.

THRUST-VECTORINGCONTROLSYSTEMDESCRIPTION

Figure2 showsthe vaneconfigurationfor theleft engineonly. Theupper-vanecenterlineis

5° outboardof theverticalplane.Theoutboard-vanecenterlineis 118° counterclockwisefrom the

upper-vanecenterline.The outboard-vanecenterlineto the lower-vanecenterlinemeasurement

is 103.5° counterclockwise.Thelower-vanecenterlineto theupper-vanecenterlinemeasurement

is 138.5° counterclockwise.The uppervanewas largerthanthe outeror lower vanes(fig. 3)

becauseof theunevenradial spacingcausedby structuralconsiderations.Theexhaust-plumeside

of eachvaneis concave,with eachvaneformingpartof a sphericalsurfaceof 36in. radiusaxially

andlaterally.

The total amountof turningof thejet exhaustplume,or jet-turning angle,is definedasthe

rootmeansquareof theequivalentthrust-vectordeflectionanglein pitchandyawasmeasuredby

theresultantforce(fig. 4). Theaxialthrustlossforthedeflectedflow is definedasthelossin the

Upper vane

Maximum afterburner /t_"__j _'_"

nozzle 348 In2_ //__'/'/_/'/_/_ _'"_x

/ _//_J/L \\\ ',

/ l _/__ _ \ 138.5"

- ... ,

Mll.ry power i_ _ _//

Outer VlnI _ " ,_,_. /

103.5-- .... -." '_ _" 960693

Figure 2. Thmst-vectofing control system end-view of the left engine looldng forward. Radial vane

included angles are displayed. All vanes shown fully extended into the plume. The fight engine is

a mirror image of the left.

6
20 In.

-i--

12.591 In

' T

/

5.185 In. __ Radlus=361n. -gP--- 9.120 in. -_

20!n, /

Lower and

outer vanes

15 In.

Area = 263.64 In 2

Jr

m

Upper vane

20 In.

Area ,=358.76 In 2

Figure 3. Dimensions for flight vanes.

7

12,420 In.

Englne

nozzle

Normalized axial thrust =

Axial force

(deflected flow_/_ Undeflectedthrust vector ---_ Enginecenterllne7

-. -. angle

Axlel force f

Undeflected thrust thrust magnltuds 960707

Figure 4. Schematic of jet-turning angle and axial thrust loss definitions.

thrust of the axial force when compared to the undeflected thrust. The normalized axial thrust is

the absolute value of the axial force divided by the absolute value of the undeflected thrust.

TESTING AND MODELS

Several tests were conducted to characterize thrust vectoring on the F-18 HARV. These tests

were two subscale aerodynamic interaction tests, one of which is discussed here; a subscale

cold-jet test; a full-scale hot loads test; and the flight test (table 1).

The 16 percent-scale aerodynamic interaction model was tested in the 30- by 60-ft wind

tunnel at NASA Langley. 16 This particular F/A-18 model was used because, being a free-flight

model, it already had ducting for the inlets, exhausts, and high-pressure air built into the design. 4
Table 1. Test matrix.

Test Model Scale Comments

Aero 16 F-18 HARV 16 percent

Aero 17 F-18 HARV 10 percent

Cold jet 16,18 Left nozzle 14.25 percent

Hot loads 19 F-18 HARV Full

Flight parameter

identification23, 30

F- 18 HARV Full

External flow,

nonmetric nozzles

Internal/external flow,

nonmetric wing

Internal flow

Static ground engine

run test

Parameter identification

flight test

The model, designed for use in free-flight testing, was of very light construction, which resulted

in tests being conducted at low dynamic pressures to prevent structural damage. The radial

location of the vanes on the model was oriented in the same way as for the aircraft. The hinges

were located to allow correct vane position and deflection angles. A spin chute canister was also

included in the modifications to the model. The primary difference between the aircraft and the

aerodynamic interaction model configuration was the nozzle areas. The nozzles used in the

aerodynamic interaction test were 13.31 in 2 in area, which corresponds to a 520-in 2 nozzle area

on the full-size F- 18 airplane. The tailpipe area nominally used during flight tests is 348 in 2 in the

thrust-vectoring envelope. Differences between the aircraft and the low-speed aerodynamic

interaction model also included AIM-9 missiles (Raytheon, Bedford, Massachusetts) in place of

research airdata wingtip probes, the omission on the model of video camera fairings used on the

aircraft, and the lack of incorporation on the wind-tunnel model of the horizontal stabilizer area

reduction (1.6 percent in area) for thrust-vectoring fairing clearance used on the aircraft.

The 10 percent-scale test was conducted at NPRs between 1.0 and 5.0, free-stream Mach

numbers between 0.3 and 0.7, and angles of attack between 0 ° and 700.17 This model was

wingtip-supported and differed from the flight vehicle configuration by having faired-over inlets,

nose strakes, and larger leading-edge extension slots. Significant modification was undertaken to

make the 10 percent-scale F/A-18 model metric aft fuselage area geometrically resemble

the HARV configuration with postexit vanes and their associated actuator fairings and the spin

chute canister.

A 14.25 percent-scale model of the left exhaust nozzle with postexit vanes of the F-18 HARV

aircraft was tested. 16,18 The objective of this static test was to gather a comprehensive set of data

from direct jet-turning effectiveness of the postexit vanes. Data were gathered with a force and

moment balance. The cold-jet test was performed in the NASA Langley 16-ft Transonic Tunnel.

The comprehensive data set from the cold-jet test formed the baseline of comparison to all

subsequent tests. Two nozzle sizes were used in this test. These sizes corresponded to militarypower and maximum-afterburner nozzles at a nominal flight condition in the thrust-vectoring
envelopeof the aircraft. The vane angleswere measuredwith simple protractors, and their

positionsvariedfrom-10 ° to 30° deflection(approximately0.5° accuracy).TheNPR variations

werefrom 1.0(noflow) to 6.0for mosttestcases.

Thestatictestof the TVCS installed on the aircraft with running engines was called the "hot

loads" test. The hot loads test was the first test of the F-18 HARV with the TVCS installed and

was conducted at NASA Dryden using the thrust stand at Edwards AFB. One research objective

of this portion of the study was to determine the feasibility of using digitized infrared imaging as

a technique to determine thrust-vectoring effectiveness. Infrared imaging provided an ability to

directly compare exhaust-plume deflections to the cold-jet test ground facility predictions. With

the inability to measure actual vectoring forces developed during the hot loads test, infrared

imaging could offer an alternative to validate the flight hardware implementation of thrust

vectoring. Test conditions of the hot loads tests included military-power, midafterburner-power

(with a nozzle exit area of 56 percent), and maximum-afterburner nozzles with all combinations

of single-vane and two-vane equal deflections. The aircraft was secured to the thrust stand, which

could directly measure thrust loss during the test. 19

Flight tests of the F-18 HARV were conducted at NASA Dryden. Parameter identification

maneuvers were used to determine stability and control coefficients using the error output

technique and pEst code. 20-22 One parameter extracted from the flight data was thrust-vectoring

control effectiveness, of which some results have been reported. 23,26-28 This parameter

estimation technique does not differentiate between thrust-vectoring forces and moments and the

aerodynamic interactions caused by thrust vectoring. As a consequence, these two parameters are

inseparable and cannot be independently identified. During the F- 18 HARV flight testing of these

data, however, the aerodynamic interaction was considered to be very low because of the

low-dynamic pressures level-flight 1-g condition used during the parameter identification

maneuvers. As a result, the predominant characteristic estimated is that of the thrust-vectoring

effectiveness. The low-dynamic pressure data are a result of all test points being flown in t-g

level flight and high angle of attack confining the aircraft to relatively slow flight. Further

complicating the issue is the mixer function in the flight control laws preventing individual vanes

from being controlled. As a result, a collective vane function is identified, and estimates of

vectoring control power from flight are made and then compared to equivalent collective vane

effectiveness from the cold-jet test data.

THRUST-VECTORING RESEARCH RESULTS

The research conducted was divided into four areas. These four areas are: aerodynamic

interaction tests, the static cold-jet test, the static engine run ("hot loads") test of the installed

system on the aircraft, and the flight test of the aircraft.

Aerodynamic Interaction Tests

Aerodynamic interaction with thrust vectoring was an unknown; therefore, two tests were

conducted to investigate aerodynamic interaction. The tests were a 16 percent-scale low-speed

test with a subsonic ejector exhaust conducted in the NASA Langley 30- by 60-ft low-speed wind

9
tunnelanda 10percent-scaletestwith asonicnozzleconductedin the 16-ftTransonic Tunnel at

NASA Langley.

Aerodynamic forces and moments on the F-18 HARV with the installed TVCS are taken at

the 0.24-chord position. Vectoring direction definitions for plume deflections are as follows: up is

negative, down is positive, right is negative, and left is positive. When the thrust-vectoring vanes

are not installed, the nozzle configuration is defined as unvectored. For any one engine, when two

vanes are deflected into the exhaust, the third vane is retracted at- 10 ° deflection and is referred to

as the retracted vane.

Significant findings of the low-speed test were that forces were mildly adverse and moments

were mildly proverse. For these aerodynamic interaction tests, the plume deflections were

-17 ° up, 14 ° down, and -9 ° right. A change in lift coefficient of approximately 0.1 caused by

vectoring in the pitch plane was found over the entire angle-of-attack range (fig. 5) for the tested

condition. A change in pitching moment of approximately 0.1 was found near 0 ° angle of attack,

decreasing to 0 increment in moment near 65 ° angle of attack with vectoring in the pitch

plane (fig. 6). Vectoring was intended primarily as a moment-producing effector, so vectoring the

plume up would cause a noseup pitching moment. But the exhaust plume vectored up would

decrease the lift coefficient in the adverse direction while increasing the pitching moment

coefficient in the proverse direction. This behavior is analogous to a blown flap.

The 10 percent-scale test 17 was conducted to investigate aerodynamic interactions over a

representative configuration. The results of the 10 percent-scale test expanded the database and

supported previous conclusions of the 16 percent-scale test, in particular that substantial gains in

controllability are provided by a multiaxis thrust-vectoring capability.

2.5 _ Vector

direction

Unvectored

F o, Right

A

1.5

CL 1.0

.5

0

-.," I

0 10 20 30 40 SO 60 70

Angle of attack, (leg _OT0e

Figure 5. Pitching moment coefficient as a function of angle of attack, with thrust vectoring varying

direction; 8 e = -12 °, _ia = 0 °, 8 r = 0 °, and C t = 0.8.

10
'4 F Vector

r_ direction

3 _ 0 Unvectored

• [] Up

,_ _ _ Right

Cm

-.3 I l I I I [ I

0 10 20 30 40 50 60 70

Angle of attack, deg 96o7o9

Figure 6. Pitching moment coefficient as a function of angle of attack, with thrust vectoring varying

direction; 81el= 34 °, 8tel= 0% _5e = -12 °, 8 a = 0% 8r= 0% and Ct = 0.8.

Thrust-Vectoring Effectiveness Tests

Although many tests were conducted to examine the effectiveness of thrust vectoring, three

tests in particular are examined here. The tests in this section are the cold-jet test, the hot loads

test, and the in-flight parameter identification tests.

Cold-Jet Test

The objective of the cold-jet test was to gather a comprehensive set of data of the direct

jet-turning effectiveness of the postexit vanes. Data were gathered with a force and moment

balance, and plume-turning angles were inferred from the force and moment data. Because of the

nonaxisymmetric installation of the vanes, pitch and yaw deflection angles are not discussed; the

total root mean square of the pitch and yaw vane deflections is used instead.

Control effectiveness is defined as the angle through which the exhaust plume is turned by the

vanes, regardless of the ultimate forces or moments produced. This effectiveness definition is an

important distinction, as the thrust available to be turned is much larger in maximum-afterburner

nozzle than in military-power nozzle.

The large upper vane produced greater plume deflection than the two small (lower and outer)

vanes did. In the military-power nozzle case, the large vane was 50 percent more effective than

the small vanes were; and in the maximum-afterburner nozzle case, the large vane was 25 percent

more effective than the small vanes. The vanes did not become effective until they impinged on

11
theedgeof theexhaustplumeatapproximately15° vanedeflectionfor military-powernozzleand

approximately10° deflectionfor maximum-afterbumernozzlefor the selectedNPR. This vane

ineffectivenessresults in larger deadbandfor the smallermilitary-power nozzle thanfor the

larger maximum-afterburnernozzle. The slopesof the vanedeflection-to-plume deflection

curvesarealsoclosebetweenthe two testednozzles,with themilitary-power nozzlehavinga

slightly largerslopethanthemaximum-afterburnernozzle.

Also as expected,the NPR largely affectseffectiveness.As theexhaustplume expansion

increaseswith increasedNPR,thevanesbecomeeffectiveat decreasedvanedeflections(fig. 9).

The slopeof the variouscurvesarerelatively insensitiveto NPRvariationin the 15°-20° vane

deflectionregion.Thesedataarefor all single-vanedeflectionswith themilitary-powernozzleat

anNPRof 2.

Two vanesusedsimultaneouslyshowgreatereffectiveness(fig. 10)for the samedeflection

whencomparedwith singlevanes(fig. 7). Effectiveness increased to approximately 100 percent

in this case. Also, two vanes radially spaced closer together (fig. 3) show greater effectiveness

than vanes spaced further apart. This difference could be explained by flow "leaking" through the

gap between vanes, and a smaller gap will leak less. The lower and upper vanes (138.5 ° radial

spacing (fig. 3)) were less effective than the lower and outer vanes (118 ° radial spacing (fig. 3)).

Cold-jet test results determined that vane spacing is a very strong and important parameter, more

important than vane size. These data are for the military-power nozzle at an NPR of 2.

Axial force loss was another measurement of the cold-jet test. Axial force loss data

are normalized axial thrust. In all cases, increases in vane deflection caused greater axial force

loss (fig. 11). In this sample case of single-vane deflections, the large upper vane caused the

lowest normalized axial force (0.82) at 30 ° vane deflection, with the two equally sized vanes

producing equal amounts of thrust loss (0.87) within experimental error. These results are for the

military-power nozzle case with an NPR of 2.

The extreme edge of the plume-deflection envelope in pitch and yaw is shown with variation

in NPR (fig. 12) for the maximum-afterburner nozzle for vane deflections of 30 °. The extreme

envelope is found by holding at least one vane at 30 ° deflection and at least one other vane

at -10 °. Retracted vane interference effects can be observed at three of the comers, near pitch and

yaw coordinates (11, 9), (-25, -3), and (16, -20). The loss of turning effectiveness caused by

increasing NPR is most evident near these comers, especially near coordinate (11, 9) in pitch and

yaw. The other three comers, (18, -3), (-7, -10) and (-9, 9), are single-vane deflection comers

and show no retracted vane interference effects. The data shown are for the left engine only and

display the actual pitch and yaw plume-deflection angles, not the plume-deflection angle to the

centerline as previously shown (figs. 7 to 11).

Full-Scale Static Thrust Test Infrared Data

Infrared imaging can visualize the hot exhaust plume of a jet engine. Infrared intensity is

related to temperature, although in a nonlinear function. The infrared representation of the

exhaust plumes, however, could be digitized and treated mathematically to calculate plume

deflections. Reducing the data obtained during hot loads testing to simple digital values with no

mention of infrared intensity or wavelength and circumventing classification of the results was

possible. The fundamental assumption is that the infrared intensity and the plume profiles

are equivalent.

12
35

30

25

8 idr_gs, 20

15

10

5

Ol

-10

-- Upper vane Vane

._l _ Upper

-- .J" _,. -'B-- Outer

Ov:tne: _ L:aW:r _ Lower

-- Left engine

- 5 0 5 10 15 20 25 30

8 v, deg

960710

Figure 7. Jet-turning angle as a function of single-vane deflection for all three vanes with the

military-power nozzle and NPR = 2.

35 -- Upper vane Vane

i__ Upper

30 -- _ Outer

Outer Lower -_-- Lower

vane vane

25 -- Left engine

8irma, 20 --

deg

15

10

- 10 - 5 0 5 10 15 20 25 30

8 v, deg

960711

Figure 8. Jet-turning angle as a function of single-vane deflection for all three vanes with the

maximum-afterburner nozzle and NPR = 2.

13
_Jrma'

deg

35

30

25

2O

15

Upper vane NPR

_ 3

Outer Lower _ 4

vane vane _ 5

Left engine _ 6

10

o_

- 10 - 5 0 5 10 15 20 25 30

8v, deg

960712

Figure 9. Jet-turning angle as a function of single-vane deflection (upper) and varying NPR with

the military-power nozzle.

35

30

25

20 --

15

10 --

5 --

-10

Outer

vane

Upper vane

Left engine

Vane

Upper and outer

Outer end lower

Lower end upper

-5 0

0

10 15 20 25 30

_v, deg

960713

Figure 10. Jet-turning angle as a function of two-angle equal deflections for all two-vane

combinations with the military-power nozzle and NPR = 2.

14
Normalized

axial

thrust

1.05 --

1.00

,95 -

.90 --

.85 --

.80 --

.75 --

.70

-10

Outer

vane

Upper vane

_Lo_r

vane

Left engine

r i i

-5 0 5

_ "'13-- Outer

I I r I I

10 15 20 25 30

_v, deg

960714

Figure 11. Normalized axial force thrust loss as a function of vane deflection angle for all singlevane deflections with the military-power nozzle and NPR = 2.

- 3O

- 2O

-10

10 --

20 --

30

30

-- Upper

vane

Outer _ L.o_r

-- vane vane

Left engine

NPR

.......... 3

4

5

6

I I I I

20 10 0 - 10

By,deg

(16, - 20)

I I

- 20 -3O

960715

Figure 12. Maximum jet-turning angle envelope, in pitch and yaw, where at least one vane is

deflected 30 °. Maximum-afterburner nozzle and varying NPR showing retracted vane interference

near comers with two-vane equal deflections. Symbols omitted for clarity.

15
Limitationsof this techniquearethatthedataareonly two-dimensional.Geometricsetupof

thecamerawascarefullyalignedwith theplaneof theexhaustnozzlesandtheexhaustplumesso

that metric datacould be madefrom theinfrareddata.An additionallimitation is that in cases

wheremultipleplumesexist,reconcilingtheresultantfrom theindividualplumesis difficult.

Five test conditions were selected. Thesetest conditions were: lower-vane deflection

variationsin military power,lower-vanedeflectionvariationsin midafterburnerpower(a nozzle

exit areaof 56 percent),upper-vanedeflectionvariationsat maximumafterburner,equallowerandouter-vanedeflectionvariationsatmilitary power,andequallower-andouter-vanedeflection

variationsat midafterburnerpower.

Theprocessto gatherdataandanalyzeeachtestconditionwasasfollows: only oneenginewas

activein eachtest,andtheotherenginewasat idle. Theidle conditionresultedin nosignificant

signaturein thefinal images,sonoattemptwasmadeto subtracttheidle plumefrom the image.

Somebackgrounddatadid exist,however,whichtendedto biasresultsto themiddleof theimage.

This backgroundwasremovedby simplysettinganartificially highthresholdjust abovethelevel

of the background.This backgroundis actualsignalandnot noisein the data.Noisewithin the

infrared datawasnot aproblem.Typical variationsthat couldbe ascribedto noisewere much

lower than the level of the signal, with the signal-to-noiseratio being approximately 50:1.

Approximately5 secof stabilizeddatawererecordedateachtestpoint,resultingin approximately

63framesof data.These63datasetswereaveragedfor thefinal analysis.

After capturingthedigital data,thedatawereconvertedto analphanumericformatthatcould

beeasily manipulated.The format correspondedto rows andcolumnsof the pixels from the

captureddata(fig. 13).("Row" and"column"referto thealphanumericrepresentationof thedata

manipulation and not to the physical geometric presentationof the pixels if the data were

reassembledto makeaphotographicimageasin thefigure.)By examiningeachcolumnandrow

Screen has 1:1 aspect ratio

Plxel has 2:1 aspect ratio

280

columr

960716

Figure 13. The layout of the rows and columns in one frame of infrared data.

16
individually, curvesof infrared intensity were generated. Routines were constructed to integrate

the areas under the curves and provide weighted averages of the curves. Simple calculations

provided the centroid of each area corresponding to the representative column. Plotting these

columns gives a graphical representation of the exhaust plume infrared intensity curves and the

plume distribution (fig. 14). Comparing centroids of individual plume curves with vane deflection

to centroids of undeflected plumes provides a direct measurement of the plume-deflection angle.

6OOO , Col

8o

....... 90

5000 .... 100

.......... "-......... "- ........ _......... " ......... " ..... _--- 120

4000 .......... ,-........ - ...... ,-......... ,-......... ,-.....

/

Measured

infrared 3000

Intenslty !

2oo0 ......... :i-....... /Lr, :_ii ....I- -_'a-_ ,, ,.i-'_ " ......... :_......... ::-.....

___ /_, ....... ......... ;_____'

: " /, , ,

-- . ,

i

o iso 21)o 21o

Row numbers

960717

Figure 14. Typical distributions of infrared intensity in vertical sections of one frame of data.

A trapezoidal integration was applied to each column of pixels to find the area under each

curve. The area of each trapezoid is given as:

J_xi+lf(x)dx " f(xi) +f(xi+ 1)(Ax) = _(fi +fi+ 1) (1)

i 2

Because each vertical column of pixels is exactly the same width, the calculation is given as:

x n

L_f(x)dx " _ _(fi +fi+l)

i=1

(2)

When the centroid of each area is found, it is stored as the centroid of that column. Each

column is treated likewise, and the complete figure of geometric plume centroids can be plotted

for each vane deflection case (fig. 15). The low or negative vane deflections with small variations

in plume deflection show little or no potential effectiveness.

17
210

Plume 200

centerline

location,

row number 190

180

170

160

6q

, i i i ,

b t t J _ , o. *S ,

i 4 t d sis.° _ ._.s. -'_*,

', , , oo" , ..--"7 , --"7

, ..__" ___.....,......

,,___.-.-

:,,, ,_,,.,;,;

• "-:'' " -:" - -: :

_ t t i r i

iO 80 90 100 li0 120 li0 40

Column number

960718

250

....... 20 °

15 °

10 °

0 o

iiiiiii _5 0

..... 10 °

Figure 15. Typical plot of the loci of the centroids of vertical sections of data, or calculated

plume centerlines.

Assuming the locus of centroids is equal to the theoretical plume centerline gives the

deflection angles. Using the raw loci in such a manner and plotting them was referred to as the

one-region technique.

The one-region technique was modified by considering the area above and below the

centerline of the nondeflected plume as separate areas. Weighted averages of the two areas were

made to determine the mean plume centerline when the plume was split, as in two-vane deflection

cases. This second method was called the two-region technique. The two-region technique was

developed to handle the complex flow structures resulting from two-vane equal deflections into

the flow. In these cases, large secondary jets were found, and the results of the one-region

technique were poor. As the geometric setup of the infrared system could only measure angles in

pitch only, the plume was assumed to move in the line of action directly opposite to the vane

radial angle. By using this assumption, the angle of the plume could be estimated in yaw.

Comparison of hot loads test infrared data with the NASA Langley cold-jet test data shows

varying correlations for many conditions. Examples are for the lower-vane deflection with

military-power nozzle (fig. 16), lower-vane deflection with midafterburner-power nozzle

(fig. 17), upper-vane deflection with maximum-afterburner nozzle (fig. 18), lower- and outervane deflections with military-power nozzle (fig. 19), and lower- and outer-vane deflections with

midafterburner-power nozzle (fig. 20).

Differences between the two analysis techniques, one-region and two-region, were small for

lower-vane deflection in military-power setting (fig. 16) and were larger for midafterburner-power

setting (fig. 17). The two techniques did produce varying results. Military-power setting and

18
single-vanedeflectionworkedwell withtheone-regionmethodwhencomparedto thecold-jettest

results (fig. 16). In the midafterburner-powerwith two-vanedeflection case,the one-region

methodalsoseemedto matchthecold-jettestdatabetter,within 2° for mostof therange(fig. 20).

The two-regionmethodseemedto matchcold-jet testdatabestwith themidafterburner-power

settingandsingle-vanedeflection(fig. 17),althoughnotaswell asasinglevanein military power.

Andin thecaseof military powerwith twovanes,mostof theoperatingregionof thecold-jettest

datais matchedbetterby thetwo-regiontechnique(fig. 19).Trendsineachof thesecasesarewell

definedandcorrelatewell, butabsolutevaluesof plumeturning--effectivenessresultsaremixed.

Plume

deflection,

(leg

One mglon

Two region

i

- - |

p

-2 i I I I

0 5 I0 15 20 25

Vane deflection, (leg 96o719

Figure 16. Lower vane deflection with nozzle in military power for one-region and two-region

infrared compared with cold-jet test data.

Plume

deflection,

dog

12

10

One reglon ' '

TWO region , , .

........... ....... ............................................ iiiiiiiiiiiiiiiiiiiiiii 

5 10 15 20 25

Vane deflection, dog 96o720

Figure 17. Lower vane deflection with nozzle in midafterbumer for one-region and two-region

infrared compared with cold-jet test data.

19
20

One reglon

Cold jet

)

15 ]

Measured

plume 10

deflection,

deg

5

(

[

0 5 10 15 20 25

Vane deflection, deg 960721

Figure 18. Upper vane deflection with nozzle in maximum afterburner for one-region infrared

compared to cold-jet test data.

25

One r'egion

Two _eglon

20____-____c,_,:_:-" - .......... :............ :.... --------_,---------z_

]

Plume 15

deflection, 10

dog

, ', ,

, b n

0

, i b

_5 I I I I

0 5 10 15 20 25

Vane deflection, dog 960722

Figure 19. Lower and outer vanes deflected together with military-power nozzle for one- and tworegion infrared compared with cold-jet test data.

20
3O

One re*ion

Two region

Colalet

20 )

,,,.m.

deflection, 10

deg

o<.

D []

-10 ....

0 5 10 15 20 25

Vane deflection, deg s_0723

Figure 20. Lower and outer vanes deflected together with midafterburner nozzle for one- and tworegion infrared compared to cold-jet test data.

Axial force, or thrust, was the only parameter for which direct measurement could be made

during the hot loads tests. The thrust stand at Edwards AFB is a single-component stand and

measures axial force only. These axial force results showed excellent prediction (fig. 21) and very

close agreement of absolute values of thrust loss, with less than 4-percent error in thrust between

the cold-jet test and the thrust-stand test data. 19 These data are for the maximum-afterburner nozzle.

100 , , i ' '

Remaining 90

axial gross O Langley cold Jet(eat ': i i _ _ :

thrust, [] F-18;,Ircrafttut: ', : ', \ \ ',

percent 85 1 : ! i

........ 2 ........ ._ ........ '.. ........ ,......... ,.. ...........

VIlne 1

flx_',at -10 °

........ _ ........ * ........ r ........ (" ................. p........ 75

- 10 - S 0 5 10 15 20 25

Dual vane 2 and 3 deflection, deg 960724

Figure 21. Thrust-vectoring axial force comparison between hot loads thrust-stand and cold-jet test

data for vanes 2 and 3 equal deflections.

21
Thecold-jettestdatawerefoundto havegoodcorrelationof trendswith thehotloadstestdata

from thrust vectoring, although the comparison varied in absolutelevels with complex jet

structures.The infraredresultsshowsomepromisein beingableto characterizethrust-vectoring

effectiveness,especiallywith simplejet structures.

Thrust-VectoringFlightTests

Parameterestimationtechniquesfrom flight testdataproducelinearizedestimatesof stability

and control derivatives. Other complications arise becauseof the mixer of the control law

preventingindividual vanemotions.In thiscase,anaveragedeflectionof thevanesis madebased

on themeasuredvanedeflections.

In pitch, the uppervanesactin oppositionto thelower andoutervanes.A generalizedpitch

deflectionof thevanesis made:

(dl d2 +d3)+(d4-d52d612

dpv = 2 (3)

where the vane deflections are defined as follows: dl is the upper left vane, d2 is the outer left vane,

d3 is the lower left vane, d4 is the upper right vane, d5 is the outer right vane, and d6 is the lower

right vane.

Likewise in yaw, the left vanes of each engine act in opposition to the right vanes of each

engine. In this case, the upper vanes do not act strongly in this axis, and deflections of the upper

vanes are ignored in yaw. The generalized yaw deflection is then defined as:

(d2 - d3) (d6- d5)

+

2 2

dyv - 2 (4)

Estimates of the generalized pitch and yaw deflection vane control power are made in a

similar manner. These generalized pitch and yaw control power estimates from the cold-jet test

data are used as the baseline to which the flight data are compared. The linearized terms are made

from the cold-jet test data in the 15°-20 ° vane deflection range with military-power nozzle. This

range of vane deflections corresponds to the operational area, away from the effects of the plume

boundary and nonlinearities near the extreme vectoring envelopes (fig. 12). Referring to the coldjet test results for single-vane deflections (fig. 7) results in an upper-vane effectiveness of

approximately 0.92 ° of plume deflection for each degree of vane deflection. Doing the same for a

two-vane deflection (fig. 10) for the lower and outer vanes gives a slope of approximately

0.97 deg/deg; the generalized average of these is approximately 0.94 deg/deg.

dy dl = 20 dy d2, d3 = 20

dx dl = 15 dx d2, d3 = 15

Cdpv = 2 (5)

22
0.92 + 0.97 (6)

Cdpv - 2

Cdp v -- 0.94 (7)

Doing the same in yaw for the lower and outer vanes (fig. 7) requires rotating the axes to the

yaw plane by multiplying the cosine of each vane angle to the yaw axis (0.73 and 0.84,

respectively) to their slopes of 0.72 deg/deg (the lower and outer vanes are identical in size and

control power). Following this procedure, the generalized yaw vane effectiveness of

approximately 0.56 deg/deg is found.

dy dv = 20 dy dv = 20cos(vane 3angle )

cos(vane 2angle ) + d--x dv 15 dx dv = 15

Cdy v = 2 (8)

0.72cos(5 + 118 - 90) + 0.84cos ( 138.5 - 5 - 90)

C dyv = 2

(9)

Cdy v = 0.56 (10)

In the generalized pitch and yaw effectiveness terms, only the slopes of the control power

curves remain. The slopes were relatively insensitive to nozzle size or NPR in the operational

region of the vanes (figs. 7 to 9), so these values of vane effectiveness from the cold-jet test data

did not vary much for the operating envelope of the aircraft with the engines in military-power to

maximum-afterburner settings.

An estimate coefficient was added to the summation of forces and moments equations in the

pEst program to model the flight data. The nondimensionalizing parameters for the vane

effectiveness coefficients are different than the aerodynamic parameters, so these equations were

split into two parts.

1 2

Normal= _pv S(CNaero)+ Fnp(Cdp v dpv) (11)

1 2

Axial = _pv S(CAaero) (12)

1 2

pitch= _2pv Sc(Cmaero)+ Fnp l(Cdp v dpv) (13)

1 2

Side = _pv S(CYaero) + Fnp(Cdy v dyv) (14)

roll = _9v2Sb(Cl )

(15)

1 2 l( dyv)

yaw = _pv Sb(Cnaero) + Fnp Cdy v

(16)

23
Thesesummedforces and momentsareimplementedin the equationsof motion just as

conventionalaerodynamicforcesandmomentswouldbeaddedto theequationsof motion.An

estimateof thenetpropulsiveforce,Fnp, is required for this technique to work, and a simplified

gross thrust method was used. 29

Several maneuvers were executed for parameter identification during the flight test program

of the F-18 HARV. Three different sets are examined here: traditional pilot-flown doublets, 21 an

optimal programmed set of inputs, 30 and a series of individual doublets commanded to separate

control surfaces. 23,26-28 Conventional doublets resulted in excessive levels of correlation

between control surfaces, and transient motions were entirely nonexistent. The use of

conventional pilot-input doublets, with their large surface correlations, created large Cramer-Rao

uncertainty bounds on the estimates and little confidence in the estimates produced. The optimal

maneuvers and the series of individual doublets produced low surface correlations with low

Cramer-Rao bounds, improving confidence in the estimates produced.

Across the range of angles of attack, little variation was found in the vectoring effectiveness

terms (figs. 22 and 23). Uncertainties in the parameter identification estimated thrust-vectoring

effectiveness may be from several areas, such as aerodynamic interaction between the thrust

vectoring and the external aerodynamics of the F-18 HARV aircraft, possible measurement

errors, and residual correlation of control surfaces. The external aerodynamics errors can be

thought of as modeling error caused by the model having insufficient degrees of freedom. The

residual control surface correlations may be thought of as a deficiency in maneuver design.

However, comparison of this vectoring effectiveness, between flight and linearized cold-jet test

effectiveness terms, shows an excellent agreement between the two different data sets.

Cdpv,

rad/sec

.040

.035

.030

.025

.020

.015

.010

.005

0

! / i i i co,,,jet t. -T; : ; _ Parameter estimate

..... :......... ; ........ _ .......... from flight with

.... , Cramer Rso

i J t t

................................... i.................

t

.......

"_" ...... ---1-'"

.

.......................... '..........................

J

,

: :

I I I I I

10 20 30 40 50 60

Angle of attack, deg 96o725

Figure 22. Pitch-vectoring effectiveness comparison between flight data parameter identification

estimates and cold-jet test data using vane average positions.

24
Cdyv,

rad/sec

.010

.005

0

- .005

- .010

- .015

- .020

- .025

- .030

i i b

_iiiiiii_,iiiiiii__iilti__ iiiiiiii,iiiiiii-

_._---_ _ CoJ_ jet_data_ ' '

Parameter estid_ete ' ' '

..L

....... from fll.ght w.i_ ....... ' ....... ' ...... ' .....

Cromer Rao ' '

li0 l i i i 0 20 30 40 50

Angle of attack, deg

960726

6O

Figure 23. Yaw-vectoring effectiveness comparison between flight data parameter identification

estimates and cold-jet test data using vane average positions.

In pitch effectiveness, the parameter identification results show a slight reduction in

effectiveness in the 250-40 ° angle-of-attack range (fig. 22). Because the pitch effectiveness of the

thrust vectoring is independent of angle of attack, little or no sensitivity to this parameter is

expected. The slight sensitivity that is found could be caused by the effects of the entrained air

around the exhaust plume influencing the aerodynamics around the aircraft, which is a direct

result of a modeling error. In addition, it is possible that the large error at low angles of attack in

pitch are also caused by aerodynamic interaction. At low angles of attack, the dynamic pressure is

elevated in the 1-g flight condition, which makes the moment increment caused by aerodynamics

large. The results of the cold-jet test data are also shown along with the in-flight estimates, and the

correlation is extremely good. The cold-jet test data value of 0.94 deg/deg is also shown. This

agreement is extremely good.

In yaw effectiveness, similar results are found with respect to angle of attack in the parameter

identification results (fig. 23). A similar reduction in effectiveness, as seen in the pitch

effectiveness, is also evident in the yaw effectiveness for the same angle-of-attack region. The

yaw vane effectiveness from the cold-jet test data is 0.56 deg/deg, which is also extremely close

to the in-flight parameter identification data.

CONCLUDING REMARKS

Data from ground-test investigations and a flight-test investigation of axisymmetric nozzles

with postexit vanes that were used to vector the thrust on the NASA F-18 High Alpha

Research Vehicle (HARV) aircraft have been presented and compared. Results of these tests

25
were usedto assistin the evaluationof an operationalresearchsystemthat was installed on

theNASA F-18HARV.

Aerodynamicinteractiontestswereconducted.Aerodynamicinteractiontestdatashowthe

plumedeflectionto haveamild effectin forcesandmomentswith vectoring.Thevectoredplume

tendedto behavesimilarto ablownflap.

Thecold-jet testwasusedto predictthethrust-vectoringsystemplume-turningeffectiveness.

Theresultsof thecold-jettestwerethatlargervanesaremoreeffectiveby asmuchas50percent

with themaximum-afterburnernozzle,andasmuchas25percentwith themilitary-powernozzle,

at a nozzlepressureratio (NPR) of 2 whencomparedto the smallervanes.Increasesin NPR

(NPR= 6) canresultin asmuchas5° moreturning of theexhaustplume thanat lower NPRs

(NPR = 2) with the military-power nozzle, althoughslopesof the vane-to-plume deflection

curveswere relatively insensitiveto NPR.Two-vaneequaldeflectionscanproduceasmuchas

100percentmoreplumedeflectionatanNPRof 2 for themilitary-powernozzlewhencompared

to single-vanedeflectionresults.Military-powernozzleresultsshowslightly moreplume-turning

control powerthanmaximum-afterburnernozzleresults,althoughslopesareclosebetweenthe

two cases.Axial force loss results from the cold-jet test showed increased losses at increased vane

deflections. The vanes are stowed out of the plumes in the -10 ° deflection position and showed no

effectiveness until deflected approximately 10 ° into the plume for maximum-afterburner and

approximately 15 ° for military-power nozzle at low NPRs.

A ground test of the installed thrust-vectoring system on the F-18 HARV aircraft was

conducted. An infrared data technique to measure physical exhaust-plume deflection was used

and showed good correlation with the cold-jet test data trends, but also showed mixed results in

absolute values. Static-load thrust loss was also measured during this test, and the cold-jet and

thrust-stand test data were within 4 percent agreement.

The flight test of thrust-vectoring effectiveness of the vanes was made with parameter

identification techniques. Maneuvers were designed and implemented using an on-board

excitation system, and pilot-flown maneuvers were analyzed. Because of control system

scheduling of the vanes, individual vane effectiveness could not be directly measured, but overall

combined effects could be compared between the cold-jet and flight test data. Flight data

estimates of pitch-vectoring effectiveness show approximately 0.9 ° plume deflection for each

degree generalized pitch vane deflection, and 0.6 ° plume deflection for each degree of

generalized yaw vane deflection in yaw. This pitch and yaw vane effectiveness showed extremely

good correlation between cold-jet test data and flight parameter identification techniques.

REFERENCES

1Chambers, Joseph R. "High-Angle-of-Attack Technology: Progress and Challenges:' HighAngle-of-Attack Technology, NASA CP-3149, vol. 1, May 1992, pp. 1-22.

2Chambers, J. R., "High-Angle-of-Attack Aerodynamics: Lessons Learned" AIAA-86-1774,

June 1986.

26
3Berrier,B. L. and Leavitt, L. D., Static Internal Performance of Single-Expansion-Ramp

Nozzles With Thrust-Vectoring Capability up to 60 °, NASA TP-2364, Oct. 1984.

4Murri, Daniel G., Grafton, Sue B., and Hoffler, Keith D., Wind Tunnel Investigation and FreeFlight Evaluation of a Model of the F-15 STOL and Maneuver Technology Demonstrator, NASA

TP-3003, Aug. 1990.

5Sappington, Jeffery W. and Thompson, Robert L., "F-14A Yaw Vane Technology

Demonstration Program," Society of Experimental Test Pilots 31st Symposium Proceedings,

Sept., 1987, pp. 187-200.

6Mason, M. L. and Berrier, B. L., Static Investigation of Several Yaw Vectoring Concepts on

Nonaxisymmetric Nozzles, NASA TP-2432, June 1985.

7Mason, Mary L. and Berrier, Bobby L., Static Performance of Nonaxisymmetric Nozzles With

Yaw Thrust- Vectoring Vanes, NASA TP-2813, May 1988.

8Berrier, Bobby L. and Mason, Mary L., Static Performance of an Axisymmetric Nozzle With

Post-Exit Vanes for Multiaxis Thrust Vectoring, NASA TP-2800, May 1988.

9Capone, E J. and Mason, M. L., Multiaxis Aircraft Control Power From Thrust Vectoring at

High Angles of Attack, NASA TM-87741, June 1986.

l°Capone, E J. and Berrier, B. L., Investigation of Axisymmetric and Nonaxisymmetric Nozzles

Installed on a O.lO-Scale F-18 Prototype Airplane Model, NASA TP-1638, June 1980.

l lTam_rat, B. E and Antani, D. L., "Static Test Results of an Externally Mounted Thrust

Vectoring Vane Concept," AIAA-88-3221, July 1988.

12Regenie, Victoria, Gatlin, Donald, Kempel, Robert, and Matheny, Neil, "The F-18 High

Alpha Research Vehicle: A High-Angle-of-Attack Testbed Aircraft," AIAA-92-4121, Aug. 1992.

13Gera, Joseph, "Hight Test Status of the NASA High-Angle-of-Attack Technology Program;'

High-Angle-of-Attack Projects and Technology Conference, NASA CP-3137, vol. 2, Apr. 1992.

14Kempel, Robert, F-18 High Alpha Research Vehicle Description, Uniform Resource Locater

http://www.dfrc.nasa.gov/Projects/HARV/VDD/kempel2.html

15Bowers, Albion H., Regenie, Victoria A., and Hick, Bradley C., "F-18 High Alpha Research

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16Bowers, Albion H., Noffz, Gregory K., Grafton, Sue B., Mason, Mary L., and Peron, Lee R.,

Multiaxis Thrust Vectoring Using Axisymmetric Nozzles and Postexit Vanes on an F/A-18

Configuration Vehicle, NASA TM- 101741, Apr. 1991.

17Asbury, Scott C. and Capone, Francis J., Multiaxis Thrust-Vectoring Characteristics of a

Model Representative of the F-18 High-Alpha Research Vehicle at Angles of Attack From O° to 70 °,

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18Mason, Mary L., Capone, Francis J., and Asbury, Scott C., A Static Investigation of the Thrust

Vectoring System of the F/A-18 High-Alpha Research Vehicle, NASA TM-4359, June 1992.

19johnson, Steven A., Aircraft Ground Test and Subscale Model Results of Axial Thrust Loss

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27
2°Maine,RichardE. and Iliff, KennethW., Identification of Dynamic Systems: Theory and

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21Maine, Richard E. and Iliff, Kenneth W., Application of Parameter Estimation to Aircraft

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23Gates, Russell J., Bowers, Albion H., and Howard, Richard M., "A Comparison of Flight

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"Estimation of the Lateral-Directional Aerodynamic Parameters from Flight Data for the NASA

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29Ray, Ronald J., Evaluation of Various Thrust Calculation Techniques on an F404 Engine,

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28

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

November 1996 Technical Memorandum

4. TITLE AND SUBTrrLE

Thrust Vectoring on the NASA F-18 High Alpha Research Vehicle

6. AUTHOR(S)

Albion H. Bowers and Joseph W. Pahle

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

NASA Dryden Flight Research Center

P.O. Box 273

Edwards, California 93523-0273

9. SPONSORING/MONrrORING AGENCY NAME(S) AND ADDRESB(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

5. FUNDING NUMBERS

WU 505-68-30

8. PERFORMING ORGANIZATION

REPORT NUMBER

H-2139

10. SPONSORINGJMONITORING

AGENCY REPORT NUMBER

NASA TM-4771

11. SUPPLEMENTARY NOTES

Presented at the NASA Langley High-Angle-of-Attack Technology Conference, Langley Research Center, Hampton, Virginia,

Sept. 17-19, 1996. Albion H. Bowers and Joseph W. Pahle, NASA Dryden Flight Research Center, Edwards, California.

1211.DISTRIBUTION/AVAILABILITY STATEMENT

Unc lassified--Unlimited

Subject Category'_LO" _ '7

12b. DISTRIBUTION CODE

13. ABSTRACT (Maxlmum 200 words)

Investigations into a multiaxis thrust-vectoring system have been conducted on an F-18 configuration. These

investigations include ground-based scale-model tests, ground-based full-scale testing, and flight testing. This

thrust-vectoring system has been tested on the NASA F-18 High Alpha Research Vehicle (HARV). The system

provides thrust vectoring in pitch and yaw axes. Ground-based subscale test data have been gathered as

background to the flight phase of the program. Tests investigated aerodynamic interaction and vane control

effectiveness. The ground-based full-scale data were gathered from static engine runs with image analysis to

determine relative thrust-vectoring effectiveness. Right tests have been conducted at the NASA Dryden Right

Research Center. Parameter identification input techniques have been developed. Individual vanes were not

directly controlled because of a mixer-predictor function built into the flight control laws. Combined effects of

the vanes have been measured in flight and compared to combined effects of the vanes as predicted by the

cold-jet test data. Very good agreement has been found in the linearized effectiveness derivatives.

14. SUBJECTTERMS

Aircraft parameter identification, Cold jet testing, Flight test, Thrust vectoring,

Wind tunnel testing

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

OF REPORT OFTHIS PAGE OF ABSTRACT

Unclassified Unclassified Unclassified

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

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

15. NUMBER OF PAGES

32

18. PRICE CODE

AO3

20. UMITATION OF ABSTRACT

Unlimited

Standard Form 298 (Rev. 2-89)

Pmecnl)ed IW ANSI Std. Z39-10

298-102