NASA Technical Memorandum 104329

Inlet Distortion for an F/A-18A

Aircraft During Steady

Aerodynamic Conditions up

to 60 ° Angle of Attack

Kevin R. Walsh

NASA Dryden Flight Research Center

Edwards, California

Andrew J. Yuhas

Analytical Services and Materials

Hampton, Virginia

John G. Williams and William G. Steenken

General Electric Aircraft Engines

Evendale, Ohio

Under NASA Dryden Flight Research Center

Contract NAS 3-26617

1997

National Aeronautics and

Space Administration

Dryden Flight Research Center

Edwards, California 93523-0273
NOTICE

Use of trade names or names of manufacturers in this document does not constitute an official

endorsement of such products or manufacturers, either expressed or implied, by the National

Aeronautics and Space Administration.
INLET DISTORTION FOR AN F/A-18A AIRCRAFT DURING STEADY

AERODYNAMIC CONDITIONS UP TO 60 ° ANGLE OF ATTACK

Kevin R. Walsh

NASA Dryden Flight Research Center

Edwards, California

Andrew J. Yuhas

Analytical Services and Materials

Hampton, Virginia

John G. Williams and William G. Steenken

General Electric Aircraft Engines

Evendale, Ohio

ABSTRACT

The effects of high-angle-of-attack flight on aircraft inlet aerodynamic characteristics were

investigated at NASA Dryden Flight Research Center, Edwards, California, as part of NASA's

High Alpha Technology Program. The highly instrumented F/A-18A High Alpha Research

Vehicle was used for this research. A newly designed inlet total-pressure rake was installed in front

of the starboard F404-GE-400 engine to measure inlet recovery and distortion characteristics. One

objective was to determine inlet total-pressure characteristics at steady high-angle-of-attack conditions. Other objectives include assessing whether significant differences exist in inlet distortion

between rapid angle-of-attack maneuvers and corresponding steady aerodynamic conditions,

assessing inlet characteristics during aircraft departures, providing data for developing and verifying computational fluid dynamic codes, and calculating engine airflow using five methods. This

paper addresses the first objective by summarizing results of 79 flight maneuvers at steady aerodynamic conditions, ranging from-10 ° to 60 ° angle of attack and from-8 ° to 11 ° angle of sideslip

at Mach 0.3 and 0.4. These data and the associated database have been rigorously validated to

establish a foundation for understanding inlet characteristics at high angle of attack.

NOMENCLATURE

A8R

/kiP

ALF

AOA

right exhaust nozzle throat area, percent

aerodynamic interface plane

aft looking forward

angle of attack, deg
AOSS

ARP

CFD

Cp

Cp*

DFRC

DP/PC

DP/PR

F/A

FS

FVG

G •

GEAE

HAP

HARV

HATP

HPC

HPL

HPVG

LEX

LPC

LPT

m

MIL-STD

n

N1SR

N2SR

P6REF

PAV

PCM

PLAR

PMIN

PRBAYC

angle of sideslip, deg

Aerospace Recommended Practice

computational fluid dynamics

pressure coefficient

pressure coefficient corresponding to local speed of sound

NASA Dryden Flight Research Center, Edwards, California

circumferential distortion descriptor

radial distortion descriptor

fighter-attack aircraft

fuselage station

fan vane guide position, deg

power spectrum density function, psi2/psi2/Hz

General Electric Aircraft Engines, Evendale, Ohio

HARV analysis program

High Alpha Research Vehicle

High Alpha Technology Program

high-pressure compressor

pressure altitude, ft

compressor vane guide position, deg

leading-edge extension

low-pressure compressor

low-pressure turbine

number of data samples

military standard

number of total pressure probes

right fan rotor speed, rpm

right compressor rotor speed, rpm

augmentor static pressure used as reference source for PS6 and PS7 pressures, psia

face average total pressure at the AIP, psia

pulse code modulation

right engine throttle position, deg

minimum total pressure at the AIP, psia

right engine-bay pressure used as reference source for PSE and PTE pressures, psia
PS0

PS211

PS212

PS3R

PS6(3)

PS7(4)

PSB1

PSB2

PSD

PSE

PSE8

psfd

PSFUELI

PSFUELO

PSI

psia

psid

PSK

PTE

PT0

PT56(20)

PTK

PW(pk-pk)

QBAR

REC

rills

RTTM

SAE

sps

T1R

T1RRB

T56HR

free-stream static pressure, psfd

fan exit static pressure at 300-deg position, psia

fan exit static pressure at 120-deg position, psia

combustor static pressure, psia

augmentor inlet static pressures, psid

nozzle inlet static pressures, psid

combustor static pressure at 270-deg position, psia

combustor static pressure at 90-deg position, psia

power spectral density function, psi2/psi2/Hz

low-response inlet surface static pressure, psid

low-response static pressures at the AIP

pounds per square foot, differential pressure

inlet fuel pressure of main fuel control, psia

outlet fuel pressure of main fuel control, psia

Pressure Systems Incorporated, Hampton, Virginia

pounds per square inch, absolute pressure

pounds per square inch, differential pressure

high-response inlet surface static pressure, psia

low-response inlet total pressure, psia

free-stream total pressure, psia

right turbine discharge pressures, psia

high-response inlet total pressure, psia

peak-to-peak variance of the planar wave descriptor

freestream dynamic pressure, lbf/ft 2

inlet recovery value

root-mean-squared

real-time thrust methodology

Society of Automotive Engineers

samples per second

right engine inlet temperature, °R

fight engine inlet temperature with bias, °R

fight engine turbine exhaust gas temperature, °R
TT0

TU

W1R

WFABM

W'FABMT

WFABP

WFE

WFEP

WFET

free-stream total temperature, °R

inlet turbulence parameter

right engine corrected inlet airflow, Ib/sec

right engine main augmentor fuel flow, gal/min

fuel temperature of WFABM, °F

right engine pilot augmentor fuel flow, gal/min

right engine combustor fuel flow, gal/min

right engine combustor fuel flow, lb/h

fuel temperature of WFE, °F

INTRODUCTION

Inlet pressure distortion effects on the propulsion system at high angle of attack (AOA) during

steady aerodynamic conditions, rapid aircraft maneuvers, and aircraft departures are not thoroughly understood. A team of NASA and industry researchers was formed as part of the NASA

High Alpha Technology Program (HATP) to investigate inlet characteristics, inlet and engine

compatibility, and prediction methodologies at high-AOA conditions. This effort addressed questions that have arisen during past aircraft development programs. These questions included the

following subjects:

At high-AOA, how do the inlet total-pressure characteristics, such as inlet recovery, circumferential and radial distortion, planar wave, and turbulence, behave as a function of angle

of sideslip (AOSS) and Mach number?

• How do inlet distortion levels during rapid, high-AOA maneuvers compare with those at

corresponding steady aerodynamic flight conditions?

What are the characteristics of the inlet during aircraft departures? What factors lead to

engine surges that have been experienced during aircraft departures? Are there any other

significant factors beyond inlet-induced distortion which account for engine surges?

• Can computational fluid dynamics (CFD) technology be used to accurately predict inlet

characteristics at high-AOA conditions?

• Are the engine airflow pumping characteristics affected at extreme maneuver conditions?

What is the best method for measuring airflow in flight at these conditions?

To address each of these questions, mapping of inlet characteristics as a function of AOA,

AOSS, and Mach number was conducted on the F/A-18A High Alpha Research Vehicle (HARV).

The inlet data obtained at steady aerodynamic conditions formed the foundation for this inlet

research effort.

4
The HARV aircraft, flown at the NASA DrydenFlight ResearchCenter(DFRC), Edwards,

California,providedtheidealplatformfor controlledexplorationof inlet characteristicsrelatedto

highly agilevehiclesatfull scale,l Thethrust-vectoringvanesystemprovidedtheability to maintain steady,high-AOA conditions.The aircraft washighly instrumented,with emphasison the

regionaroundandin the starboardinlet andengine.A newlydeveloped,40-probetotal-pressure

inletrakewasinstalleddirectlyaheadof theengine.2Surfacestatic-pressuretransducerswereinstalledat theinletrakelocationandaroundtheinlet lip. High-frequency-responseinstrumentation

was installed to monitor engineoperationand behavior.To maintainthe quality of the highresponsedata,aseriesof specificinstrumentationcalibrationswasperformedonthegroundandin

flight duringthecourseof thetestprogram.

Flight testdatawereobtainedduringsteadyaerodynamicconditions,rapidchangesin AOA

andAOSS,3andaircraftdepartures.4Theseflight testdatawerealsousedto estimateengineairflow with cleananddistortedflows5andto improveaCFDapproachof predictingdynamicdistortion.6Thesteadyaerodynamictestmatrixconsistedof 79conditionswith theengineat maximumcorrectedairflow. Themajority of theflight datawasacquiredat Mach0.3 and0.4, with limited

testingperformedat Mach0.6,0.8,and0.9.Specialin-flight testingwasperformedto verify data

repeatability.An exhaustiveverificationof thedataprocessingandreductionsystemwascompletedbeforebeginningtheanalyses.

A sophisticateddatareductionprogramwasdevelopedby GeneralElectric Aircraft Engines

(GEAE),Evandale,Ohio,to performall dataqualityandanalysiscalculations.Thesecalculations

included inlet characteristics,suchas inlet recovery,circumferentialand radial total-pressure

distortion levels, and also planar wave and turbulencevalues.The program provided timeaveragedandpeak-patternscreeningof the inlet characteristics.The distortiondescriptorswere

calculatedconsistentwith Societyof AutomotiveEngineers(SAE), AerospaceRecommended

Practice(ARP) 1420,andAerospaceInformationReport(AIR) 1419standards,7.8andestablished

GEAE methodology.9A°

This papersummarizestheresultsof inlet datavalidationandinlet performanceanalysisfor

steadyaerodynamicflight conditions.Inlet datarepeatabilityis discussedfirst by showinginletrecoverypressurepatternsanddistortionlevelsatMach0.3,30° AOA and0° AOSS.A summaryof

theeffectof AOA andAOSSoninletrecovery,turbulencelevels,andpeakdynamiccircumferential andradialdistortionisthenpresentedfor Mach0.3and0.4.A summaryof theeffectsof AOA

andAOSSon thewall staticpressuresat theinlet exit for Mach0.3 is shown.Lastly, asummary

of the effectsof AOA at 0° AOSSontheinlet cowl lip andductsurfacepressuresis presented.

AIRPLANE

TheF/A-18A airplaneis atwin-engine,single-place,multimissionfighter/attackairplanewith

all-weatherinterceptandgroundattackcapability.TheHARV is a preproductionmodel(aircraft

no.6) of theF/A-18A airplanewhichwasmodifiedwith extensiveinstrumentationandmultiaxis

thrust-vectoringpaddles(fig. 1).Thrust vectoringprovidedthe HARV with the ability to fly at

sustainedaerodynamicconditionswhich exceedthe capabilitiesof conventionalaircraft. Wing

5
leading-edgeextensions(LEX) weremountedoneachsideof thefuselagefrom thewing rootsto

just forwardof thewindshield.

PROPULSIONSYSTEM

TheF/A-18Apropulsionsystemconsistsof two inletsandtwo GeneralElectricF404-GE-400

turbofanengineswith afterburnerandthrust-vectoringpaddles.Theright-handinlet andengine

(ALF) wereusedfor conductinginletresearch.

Air Induction

Two inletsandinlet ductsmakeuptheairinductionsystemof theF/A-18airplane.Theseinlets

areside-fuselage-mountedwith fixed geometryanda single5° externalcompressionramp.The

inlets arelocatedapproximately25 ft aftof theaircraftnoseundertheLEX of the wing (fig. 1).

Figure2 showsaschematicof theairinductionsystem,includingkey inlet dimensions.Theinlets

arelocatedapproximately5in. fromthefuselageto avoidingestionof thefuselageboundarylayer.

Thefuselageboundary-layerflow is divertedupwardanddownwardby awedge-shapeddiverter

betweentherampandthefuselage.

Theexternalandinternalgeometryof theinlet cowl lip wasoptimizedfor maneuveringin the

subsonic,high-AOA regionof theflight envelope.The lowerand lower-inboardportionsof the

inlet lip were cut back and thickenedto decreasethe compressor-facedistortion at extreme

aircraftattitudes.

The lengthof thesubsonicdiffuseris approximately13ft. The diffuserlengthto engine-face

diameterratiois 5.3,whichprovidesfor agradualtransitionfrom theinlet entranceto theengine

face.The inlet entranceis offsetfrom theenginecentedineby 10in. in thehorizontalplaneand

14in. in thevertical plane.A pair of vortexgeneratorswereinstalledon thelower surfaceof the

diffuser to preventpossiblelocal flow separationin this area.References11 and 12 provide

detaileddescriptionsof theF/A-18Aair inductionsystem.

Engines

The F404-GE-400is a low-bypass,twin-spool,axial-flow turbofanenginewith afterburner.

The 3-stagefan (low-pressurecompressor)and7-stagehigh-pressurecompressorareeachindependentlydrivenby a single-stageturbine.In thefan,theinlet guidevanesandthe statorsof the

first stagearevariable.In thehigh-pressurecompressor,theinletguidevanesandthefirsttwo stator stagesarevariable.Thesevariableinlet guidevanesdirecttheinlet air atanoptimumanglefor

efficient engineoperation.Thethrough-flowannularcombustorusesatomizingfuel nozzles.The

mixed-flow augmentorburnsairfrom boththebypassandthehigh-pressurecorein theafterburner

6
section.The enginecontrol is an integratedsystem,usingboth hydromechanicalandelectronic

controlcomponents.Thesealevelstaticmilitary thrustof eachengineis approximately10,700lbf,

andthemaximumafterburnerthrustis approximately16,000lbf. Themaximumcorrectedairflow

throughtheengineis approximately144lb/sec.

Thrust-VectoringVaneSystem

Thethrust-vectoringvanesweremountedontheaircraftandpositionedabouttheperipheryof

theengine,behindthenozzleexhaust.Thecomersof eachvanewereclippedto avoidinterference

with adjacentvanesat full deflection.Electrohydraulicactuatorsprovidedvaneactuation.

Theenginesweremodifiedto accommodatethrust-vectoringvaneinstallationbyremovingthe

divergent sectionof the nozzle.The remainingconvergentnozzle hardwarewas modified to

maintainstructuralintegrity.Reference1 providesfurtherdetailsof the HARV thrust-vectoring

vanesystem.

INSTRUMENTATION

TheHARV inlet researchobjectivesrequiredaccuratemeasurementof specificinlet, engine,

and airdataparameters.Thesemeasurementsneededto be recordedduringsteadyaerodynamic

flight conditionsup to 60° AOA, rapid AOA maneuvers, and aircraft departures with possible

engine surges. This section summarizes the instrumentation used for the HARV inlet research program. Table 1 summarizes the important pressure transducers that were used.

Table 1. Summary of inlet research transducers.

Description

Measurement

Type Range sample rate, sps

High-response pressure transducers

(PTKxy and PSKxy)

Low-response pressure transducers

(PTExy and PSExy)

Inlet reference pressure (PRBAYC)

Fan exit diagnostic pressures (PS211,

PS212)

Absolute 0-20 psia 2143

Differential +5 psid 67

Absolute 0.4-19 psia 67

Absolute 0-50 psia 800

Aircraft and Engine

Aircraft instrumentation included accelerometers, rate gyros, surface position measurements,

and airdata. Data monitored from the MIL-STD-1553 data bus included 226 flight control system,

7
21airdata,37engineandaircraftdiagnostic,and32inertial navigationsystemparameters.Two

high-AOA flight researchairdatasystemsconsistingof swivelingpitot probeswith conventional

AOA andAOSSvanesweremountedonbothwingtips.13

Both engineshadbasicinstrumentationfor monitoringengineoperationandwere equipped

with areal-timethrustmeasurement(RTTM) system.14Figure3showstheright engineinstrumentation.Additionalflight testinstrumentationwasinstalledtomeasureguidevanepositions,fanand

compressorspeeds,fuel flows,diagnosticpressures,andRTTM pressuremeasurements.

Inlet Rake

An innovative inlet total-pressure distortion measurement rake was developed for the A/B/C/D

models of the F/A-18 airplane inlet by a team of NASA Dryden and GEAE personnel. Figure 4

shows the inlet rake installed at the aerodynamic interface plane (AIP) with the engine installed.

The AIP is defined as the measurement plane between the inlet and the engine where inlet distortion and performance is determined. For the F/A-18, the AIP is located 4 in. in front of the bullet

nose of the engine. The inlet rake consisted of a streamlined centerbody and eight aerodynamic

rake legs. The rake legs were equally spaced, oriented 45 ° from each other starting at 9 ° counterclockwise off top center aft looking forward (ALF) for the right-hand inlet. Each rake leg consisted

of five dual probes located at the centroids of five equal areas. The 40 dual probes measured highand low-frequency-response total pressures. Figure 5 shows the rake orientation and nomenclature

used to assign each pressure port. This inlet rake is fully described in references 2 and 15. The orientation of the rake was consistent with that used in previous F/A-18 inlet tests, tl

The inlet rake probe configuration was designed to be insensitive to flow angularity, an important criterion when measuring total pressure in distorted flows. This configuration allowed the sensors to read true pressure levels of local airflow at yaw angles from +_25° and pitch angles from 15 °

to -25 °, with positive angles being in the direction of the engine centerline. The blockage of total

airflow at the AIP caused by inlet rake probes was 0.4 percent. The maximum airflow blockage

caused by the rake structure was less than 8 percent, located 1.5 in. downstream of the AIP.

Inlet Low-Response Pressures

The low-frequency-response pressures (PSEs and PTEs) were measured using differential

pressure transducers (ESP-320S BSL, Pressure Systems, Inc., Hampton, Virginia). To obtain

absolute pressures, an accurately measured reference pressure was added to the differential pressures. The right-hand ALF engine-bay was selected as the reference pressure location.

Each differential transducer unit provided the measurement of 32 individual pressures. The

units were thermally stabilized to minimize zero drift associated with temperature variations. This

stabilization was accomplished by wrapping the transducer unit in a temperature-controlled

thermal blanket. The differential pressure transducers were capable of in-flight calibrations. The

8
calibrationsystemallowedfor the reference-pressurepneumaticsourceto be suppliedsimultaneouslyto eachsideof thedifferential-pressuresensors.Thesedataallowedfor anychangein the

zerointerceptof thecalibrationto beremovedandtherebyincreasetheoverallaccuracyof differentialpressuremeasurements.In-flight calibrationswereperformedatstraightandlevelconditions

beforeinitiating eachsetof flight researchmaneuvers.

Inlet High-ResponsePressures

The high-frequencyresponsepressures(PSKs andPTKs) were measuredusing individual

temperature-compensatedpressuretransducers(XCS-27L-093-20A,Kulite SemiconductorProductsInc.,Leona,NewJersey).Eachtransducerwasmountedattherakemeasurementportsin close

proximity to thelow-frequencyresponsepressureprobe.Thesetransducerswereselected,in part,

becauseof theirpassivetemperaturecompensationcapability.Thisability minimizedany calibration drift which occurredbecauseof changesin thethermalenvironmentat thetransducersensor.

To furtherincreasethe accuracyof thetransducermeasurement,a seriesof pressurecalibrations

wereperformedovertheentirerequiredpressureandtemperaturerange,upto 20psiaandat 395,

425,460,535,and610°R. Along with themeasuredengineinlet temperature,thesecalibrations

allowed any remainingzerothermaldrift to be removedduringpostflight dataprocessing.The

flight calibrated,low-responsepressuremeasurementswerethenusedto removeanybiasor drifting of thehigh responsepressuremeasurementsduringsteadyaerodynamicflight conditions.This

high-responsetransducersetupallowedfor accuratemeasurementof totalandstaticpressuresduring rapid aircraftmaneuversanddepartures.Thesamplerateof thesetransducerswasconfigured

to 2143sps.

InletWall StaticPressures

The inlet wall staticpressuremeasurementsurveyedtheinlet, primarily at its entranceand

exit. Twenty-eightlow-responsemeasurementsand 14high-responsemeasurementswere obtainedusingthesametransducersetupastheinletraketotal pressuresdescribedin thesectionson

inlet low- andhigh-responsepressures.Table2 showshow theinlet wall staticpressuremeasurementlocationsweredistributed.

Table2. Inletwall staticpressureprobedistribution.

Low-frequency High-frequency

Location response,PSE response,PSK

Inlet entrance 17 4

Inlet throat 2 2

Vortexgenerator 1 0

AIP 8 8
The low- andhigh-responsetranducerswerelocatedaboutthe aircraft fuselage station (FS)

positions (0 FS is the aircraft nose tip location) (table 3). The curvature of the inlet entrance made

exact fuselage station locations difficult to measure. At the inlet entrance an alternate method to

define the fuselage station location was used to specify distance from the inlet tip. Arc measurements were obtained by placing a string alongside the differential pressure transducer locations at

each clock position (0 °, 90 °, 180 °, 225 ° ALF) and marking the position on the string. The 0 ° position is located at the top of the inlet. The tip probe locations (2A, 2C, 2E, 2F) were used as the 0 in.

reference for each clock position. A positive length indicates that the port is located in the inlet,

downstream of the inlet tip. A negative distance indicates that the port is outside of the inlet, downstream of the inlet tip.

Table 3. Locations of the low- and high-frequency-response static pressure measurements.

Low-pressure transducer static pressures

Distance from

Parameter Fuselage Angular inlet cowl lip

identification station, in. position, deg* highlight, in.

PSER1A FS411.5 0 ° (outer skin) -0.750

PSER2A FS411 0 ° (lip tip) 0.0

PSER3A FS411.5 0 ° (inner skin) 0.938

PSER4A FS412.5 0 ° (inner skin) 1.875

PSER 1C FS411.25 90 ° (outer skin) -0.813

PSER2C FS410.75 90 ° (lip tip) 0.0

PSER3C FS411.25 90 ° (inner skin) 0.688

PSER 1E FS415.5 180 ° (outer skin) -0.563

PSER2E FS415 180 ° (lip tip) 0.0

PSER3E FS415.5 180 ° (inner skin) 1.063

PSER4E FS416 180 ° (inner skin) 1.750

PSER5E FS422 180 ° (inner skin) 7.750

PSER6E FS426 180 ° (by throat) 11.750

PSER7E FS495 180 ° (by vortex generator) 89.0

PSER 1F FS415.5 225 ° (outer skin) -0.813

PSER2F FS415 225 ° (lip tip) 0.0

PSER3F FS415.5 225 ° (inner skin) 1.000

PSER4F FS416.25 225 ° (inner skin) 2.188

PSER5F FS418 225 ° (inner skin) 4.250

PSER6F FS426 225 ° (by throat) 12.313

PSER8A FS551 12.25 ° (Alp) 140.0

PSER8B FS551 58.75 ° (ALP) 140.0

PSER8C FS551 97.50 ° (ALP) 140.0

PSER8D FS551 144 ° (ALP) 140.0

PSER8E FS551 189 ° (Alp) 140.0

10
Table3. Concluded.

Distancefrom

Parameter Fuselage Angular inlet cowl lip

identification station,in. position,degt highlight, in.

PSER8F FS551 234° (AIP) 140.0

PSER8G FS551 283° (AIP) 140.0

PSER8H FS551 323 ° (AIP) 140.0

High-pressure transducer static pressures

PSKR2A FS411 3 ° (lip tip) 0.0

PS KR2C FS410.75 93 ° (lip tip) 0.0

PSKR2E FS415 183 ° (lip tip) 0.0

PSKR2F FS426 183 ° (by throat) 11.75

PSKR6E FS415 228 ° (lip tip) 0.0

PSKR6F FS426 228 ° (by throat) 12.31

PSKR8A FS551 9.25 ° (AIP) 140.0

PSKR8B FS551 55.75 ° (AIP) 140.0

PSKR8C FS551 94.5 ° (AIP) 140.0

PSKR8D FS551 141 ° (AIP) 140.0

PSKR8E FS551 186 ° (AIP) 140.0

PSKR8F FS551 231 ° (AIP) 140.0

PSKR8G FS551 280 ° (AIP) 140.0

PSKR8H FS551 320 ° (AIP) 140.0

*0° = Top, clockwise, ALE

Figure 6 shows the inlet wail static locations about the inlet entrance. The low-response wall

static positions at the AIP were equally spaced between each rake and were in the same plane as

the rake probe tips. High-response positions were approximately clocked 0.5 in. (3 ° ) clockwise and

ALF from the corresponding low-pressure transducer locations.

Data Acquisition and Reduction

Data acquisition for the HARV inlet program used three pulse code modulation (PCM)

systems. Two PCM systems telemetered aircraft and engine data to ground-based computers. The

third recorded inlet specific data onboard the aircraft. The high-response absolute-pressure transducers had an onboard antiaiiasing filter applied to the measurement signals. The onboard,

six-pole, low-pass analog filter provided a nominal cutoff frequency (-3 dB point) of 400 Hz with

a 36 dB per octave rolloff. Additional filtering was applied to the signals during data reduction.

This digital filter used a 9-scan rolling window average providing a cutoff frequency equivalent to

100 Hz which is consistent with the F404-GE-400 engine response to effects of dynamic distortion.

11
To synchronizethe threePCM systems,anembeddedtime-codewas insertedinto eachdata

stream.The telemetereddatawere synchronizedwith the onboardrecordeddataat the fastest

samplerate,2143sps.Dataobtainedfor eachtestconditionwerereducedandprocessedthrough

specialdataqualityandanalysiscomputerprogramsto assurehighqualitydata.Thedatasynchronization,pressuredatacorrections,dataqualitychecks,andanalysiswereconductedpostflight.

DataQualityAssurance

A quality assurancesoftwareprogramwasdevelopedto review and verify the validity of

HARV inlet data.For eachresearchmaneuver,theaverage,maximum,minimum,andrmsvalues

werecalculatedfor all parameters.An outputsummaryfile providedaveragesandratiosof the

high-andlow-responsepressuredata,andalsoof simpleinlet flow qualitydescriptors.

Data Tolerance Levels

The quality assurance software validated data sets by comparing the current value of each

input parameter with its previous value to determine whether the latter exceeded user-defined

tolerances. These tolerances were defined by either a percentage or an absolute change in the

value. The quality assurance software replaced any out-of-tolerance data by holding the last validated value. Any out-of-tolerance data were investigated to determine whether tolerance limits

required modification. The software also calculated the average, maximum, minimum, and root

mean square (rms) values over the entire input file for all parameters.

Probe Substitution

For the 40 rake total and 8 wall static high-response pressures measured at the AlP, any transducer that was known to be inoperative was substituted. The substitution procedure was based on

information from adjacent circumferential pressure measurement locations.

DATA ANALYSIS

The HARV Analysis Program (HAP) was specifically developed to analyze HARV inlet data.

Two output files were generated for database storage. One contained the computed information for

each scan of the data record. The other contained a summary of the time-averaged dynamic and

peak parameter information obtained from each flight test maneuver.

Time-averaged values were calculated for all input parameters. Inlet and engine pressure

distortion indices were computed to describe the flow field characteristics. Output from the HAP

consisted of circumferential and radial elements of distortion descriptors displayed in tabular and

graphical form. Isocontour plots of the AIP pressures were graphically displayed to correspond to

specified peak distortion descriptors. This HAP data also provided information to assess the influence of the descriptors on the engine stall line.

12
AppendixA containstheproceduresfor computingtheparametersusedto describe inlet airflow quality. These procedures reflect a history of successful correlation of inlet distortion effects

for the F404 engine and other military and commercial engines. These procedures are based on industry standard guidelines 3'4 and are consistent with GE methodology. 5 The inlet total pressure and

flow stability descriptors described are based on the F404-GE-400 Engine Model Specification. 6

FLIGHT TEST DESCRIPTION

A matrix of aerodynamic flight conditions was required to evaluate the inlet research objectives. These conditions consisted of steady attitudes of AOA and AOSS, positive and negative

sweeps of AOA, and aircraft departures. All testing was performed between military (maximum

dry) and maximum afterburning power settings to maintain a constant engine airflow of approximately 144 lb/sec. When on flight condition, the power setting was held steady while research data

were recorded. The maneuvers at steady aerodynamic conditions were conducted at Mach 0.3 and

0.4 and from an altitude of 20,000 to 35,000 ft. Figure 7 summarizes the matrix of inlet data

obtained during steady aerodynamic conditions for a given AOA, AOSS, and Mach number.

In-Flight Calibrations

During flight, calibrations of the inlet rake and duct, low- and high-response pressure transducers (PTE, PTK, PSE, and PTE) were performed prior to a series of steady aerodynamic flight

maneuvers. The calibration required constant altitude and airspeed for 15 sec. The calibrations

were conducted in two steps. The PTE and PSE transducers were calibrated at their zero differential pressure condition by applying the reference pressure to both sides of the transducer sensing

element for 10 sec. Immediately following this calibration, 5 sec of PTK and PSK transducer calibration was obtained. The aircraft was required to maintain steady aerodynamic flight conditions

(airspeed, AOA, AOSS, and right engine military power setting) during this interval. These PTK

and PSK transducer calibrations consisted of comparing the time-averaged pressure levels of the

high-response measurements (PTK and PSK) with their corresponding low-response measurements (PTE, PSE). Because of flight constraints, the pilot typically performed this calibration in a

shallow climb.

Steady Aerodynamic Conditions

The criterion for a successful maneuver was that the flight condition tolerances be maintained.

The length of recorded data for each maneuver was established to be 6 sec; however, because of

aircraft maneuver limitations, shorter data records were obtained at the more extreme aerodynamic

conditions. To help assure the consistency of the data, the tolerance for Mach number during every

maneuver was _+0.01. The tolerance for AOA and AOSS was +1.0 °. For the +10 ° AOSS maneuvers, it was not possible for the airplane to maintain a 10 ° or -10 ° AOSS condition. As a result, the

maximum sustainable AOSS condition was used.

13
RESULTSAND DISCUSSION

This section presents a summary of inlet data obtained during steady aerodynamic flight conditions. Inlet data repeatability is discussed first by showing inlet recovery pressure patterns and

distortion levels at Mach 0.3, 30 ° AOA, and 0 ° AOSS. A summary of the effect of AOA, AOSS,

and Mach number on inlet recovery, turbulence levels, peak dynamic circumferential and radial

distortion is presented for Mach 0.3 and 0.4. A summary of the effects of AOA and AOSS on the

wall static pressures at the AIP at Mach 0.3 is shown. Lastly, a summary of the effects of AOA, at

0 ° AOSS, on the inlet cowl lip and duct surface pressures is presented.

All inlet pressure data presented here incorporate the in-flight calibrations and postflight instrumentation bias corrections. These calibrations and corrections were necessary for obtaining accurate and repeatable data.

Data Repeatability (Mach 0.3, 30 ° AOA, 0 ° AOSS)

To evaluate data repeatability, time-averaged total pressure patterns (or contours) were

obtained using high response (PTK) pressure measurements. Four of the eleven test points at

Mach 0.3, 30 ° AOA, and 0 ° AOSS were compared. The contour lines represent a deviation from

face average pressure. The four test points spanned 31 flights and 2 months.

Figure 8 shows the total pressure patterns for the four test points. All four pattern profiles are

nearly identical. Values of inlet recovery, turbulence, and maximum dynamic distortion are also

nearly identical. A low-pressure area is protruding from the lower, inboard wall. This area is indicated by the contour line labeled -6, which shows the percent decrease in pressure relative to the

face average pressure. The circumferential and radial distortion descriptors, DP/PC and DP/PR, are

nearly identical for the four steady aerodynamic conditions. Although not shown here, the distortion characteristics of the other Mach 0.3, 30 ° AOA, and 0 ° AOSS conditions are also very similar.

These data indicate that the high-response PTK pressures have excellent repeatability.

Inlet Performance at Mach 0.3 and 0.4

For figures 9-13, the data points represent time-averaged values for inlet recovery, turbulence,

and peak distortion descriptor values as a function of AOA and AOSS. A solid symbol indicates

that only one set of time-averaged data was obtained for the target AOA condition. The open symbols and the 'x' indicate that multiple data sets were obtained at the target AOA condition. All data

shown in figures 9(a), 10(a), 1 l(a), 12(a), and 13(a) were obtained at Mach 0.3 and are shown in

table 4. Figures 9(b), 10(b), 1 l(b), 12(b), and 13(b) show data obtained at Mach 0.4 and are shown

in table 5. Positive AOSS indicates that the aircraft is pointing nose-left. This orientation is windward for the right-hand inlet. Conversely, negative AOSS describes a nose-right condition. This

orientation is leeward for the right-hand inlet.

14
Table 4. Inlet performance at Mach 0.3.

Flight/

Test

Point

Hight Conditions

Mach AOA AOSS HPL PTO I"1"O

deg deg ft psia °R

215/08b 0.30 -8.6 -1.4 22924 6.3549 452.9

230/6b3 0.31 3.6 0.9 24034 6.0779 448.0

234/06b

216/14d2

211/4b3

230/12bl

216/14b

216/14c

0.31 11.0 -7.9 22537 6.4734 466.3

0.30 10.7 -5.6 22509 6.4765 458.0

0.31 9.9 0.8 22638 6.4494 463.4

0.32 9.4 0.8 21834 6.7122 458.0

0.31 9.4 6.2 21086 6.9073 461.4

0.30 9.3 8.3 22016 6.6051 458.0

Inlet data

Inlet Planar

Recovery Turbulence DP/PC max DP/PR max Wave

(PTE) (rms) (based on PTK + bias) (pk-pk)

0.9618 0.010 0.094 0.046 0.0137

0.9731 0.005 0.051 0.056 0.0104

0.9679 0.007 0.088 0.052 0.0088

0.9686 0.006 0.084 0.052 0.0098

0.9741 0.005 0.065 0.059 0.0091

0.9749 0.005 0.063 0.056 0.0072

0.9760 0.004 0.060 0.054 0.0094

0.9731 0.005 0.049 0.057 0.0100

212/08b 0.30 20.1 0.4 22926 6.3464 462.5 0.9765 0.005 0.053 0.046 0.0090

0.31 30.4 -8.3 25603 5.6851 452.6

0.28 30.8 -7.8 22828 6.3274 452.6

0.29 30.0 -3.4 23660 6.1366 449.3

0.31 30.5 -2.8 27095 5.2313 446.9

0.31 31.0 -0.3 25623 5.6686 441.8

0.30 29.8 -0.2 22132 6.5644 447.2

0.30 30.1 0.2 25916 5.5700 452.6

0.29 29.6 0.3 21584 6.7043 466.4

0.29 29.9 0.3 25052 5.7791 437.1

0.30 30.0 0.3 25405 5.7008 452.7

0.30 30.1 0.3 25118 5.7706 454.4

0.30 29.8 0.4 23176 6.2721 463.4

0.30 29.9 0.6 25782 5.6202 443.0

0.31 29.9 0.9 24828 5.8660 446.4

0.30 29.9 i .0 26301 5.4911 440.6

0.30 29.7 3.5 24226 6.0026 456.2

0.31 29.0 5.0 25972 5.5870 440.8

0.29 30.1 6.4 23818 6.0931 458.1

0.31 31.0 8.8 24490 5.9476 445.4

0.31 28.6 10.8 25294 5.7546 444.0

0.29 37.8 4.0 21533 6.7236 457.3

0.30 39.6 0.3 23645 6.1617 461.2

0.29 40.3 -3.5 21658 6.6758 459.9

211/25c

215/10c

215/10b

211/24c2

230/02b

238/26b

213/02b

234/02b

232/02b

211/02b

241/26b

212/02b

215/02b

216/02b

214/02b

211/24b

215/10d

211/25b

230/02c

215/10e

247/08b

212/09b

247/08d

232/03dl

232/3d2

242/13c

232/03c

240/12c

213/1062

231/3b2

232/03b

240/12b

240/13b

231/3c2

231/03d

0.29 49.7 -.4.4 24593 5.8809 431.3

0.29 48.9 -3.7 21426 6.7498 443.7

0.30 50.6 -316 23595 6.1583 453.8

0.29 50.3 -2.8 27064 5.2885 421.4

0.30 49.4 -1.6 21622 6.7139 452.1

0.30 49.4 -0.1 26601 5.4110 446. !

0.30 49.3 0.7 25109 5.7816 442.5

0.30 49.7 1.5 24676 5.8916 432.6

0.30 48.9 2.9 24086 6.0485 452.7

0.31 49.2 3.1 24162 6.0532 450.5

0.30 49. i 3.4 23448 6.2207 447.7

0.30 48.5 5.2 19730 7.2636 462.4

0.9605 0.010 0.098 0.046 0.0142

0.9619 0.009 0.095 0.044 0.0123

0.9678 0.009 0.090 0.046 0.0131

0.9683 0.009 0.097 0.046 0.0162

0.9701 0.008 0.090 0.048 0.0128

0.9709 0.007 0.083 0.045 0.0121

0.9695 0.007 0.082 0.047 0.0119

0.9715 0.007 0.076 0.048 0.0105

0.9697 0.007 0.085 0.046 0.0135

0.9712 0.007 0.083 0.045 0.0132

0.9698 0.007 0.082 0.048 0.0119

0.9699 0.007 0.080 0.047 0.0116

0.9702 0.007 0.084 0.046 0.0129

0.9705 0.007 0.081 0.046 0.0119

0.9702 0.007 0.084 0.046 0.0111

0.9708 0.006 0.065 0.047 0.0110

0.9719 0.006 0.059 0.051 0.0118

0.9701 0.007 0.066 0.050 0.0129

0.9683 0.007 0.074 0.056 0.0120

0.9697 0.007 0.070 0.054 0.0124

0.9610 0.009 0.102 0.059 0.0120

0.9599 0.010 0.109 0.051 0.0140

0.9567 0.010 0.113 0.053 0.0148

0.9293 0.018 0.152 0.073 0.0293

0.9407 0.014 0.128 0.062 0.0174

0.9250 0.019 0.154 0.081 0.0264

0.9297 0.018 0.147 0.075 0.0228

0.9394 0.016 0.137 0.066 0.0249

0.9380 0.016 0.126 0.067 0.0217

0.9402 0.015 0.129 0.067 0.0279

0.9445 0.013 0.123 0.063 0.0197

0.9456 0.014 0.110 0.062 0.0202

0.9450 0.015 0.130 0.068 0.0252

0.9495 0.012 0.113 0.064 0.0173

0.9503 0.012 0.110 0.059 0.0155

24716b2 0.30 56.6 -2.2 24439 5.9432 448.1 0.9118 0.019 0.163 0.081 0.0227

0.32 59.1 -2.2 27126 5.3336 439.7

0.32 59.1 -0.3 26904 5.3825 437.0

0.31 60.0 1.3 31865 4.2867 416.9

0.30 58.9 1.9 26855 5.3488 421.9

0.30 59.1 1.9 28658 4.9404 433.9

0.30 59.0 2.6 22229 6.5327 460.1

0.30 59.0 2.7 25670 5.6384 448.4

0.30 59.0 3.0 22501 6.4788 461.4

0.31 59.2 3.5 26076 5.5514 435.4

243/29c

247/6bl

247/6b3

232/05b

242/17bl

242/16b

212d05b

243/16b

231/04b

0.8999 0.023 0.188 0.094 0.0270

0.9040 0.022 0.164 0.091 0.0305

0.9040 0.024 0.167 0.102 0.0287

0.9140 0.021 0.152 0.087 0.0231

0.9093 0.022 0.160 0.087 0.0272

0.9147 0.019 0.141 0.078 0.0249

0.9139 0.021 0.149 0.081 0.0284

0.9144 0.019 0.144 0.075 0.0218

0.9145 0.021 0.162 0.082 0.0231

15
Table 5. Inlet performance at Mach 0.4.

Hight/

Test

Point

Flight Conditions Inlet data

Mach AOA AOSS HPL PTO TTO

Inlet Planar

Recovery Turbulence DP/PC max DP/PR max Wave

deg deg (PTE) (rms) (based on PTK + bias) (pk-pk) ft psia OR

215/09b 0.41 -10.3 0.2 24363 6.2847 453.7 0.9590 0.010 0.111 0.028 0.0159

230/07b 0.40 3.7 0.4 22653 6.7337 458.1 0.9776 0.005 0.043 0.055 0.0085

234/07c 0.41 9.3 6.8 22133 6.9360 468.9 0.9822 0.004 0.051 0.055 0.0066

234/07b 0.40 9.4 4.9 22099 6.8981 468.9 0.9815 0.004 0.047 0.055 0.0070

213/13b 0.39 9.9 0.6 22282 6.8125 468.9 0.9770 0.004 0.071 0.054 0.0078

234/07d 0.40 10.4 --4.7 21559 7.0655 474.3 0.9715 0.005 0.081 0.055 0.0083

234/07e 0.41 11.0 -5.9 22964 6.6691 468.9 0.9677 0.006 0.094 0.050 0.0088

213114b 0.40 19.7 0.4 32134 4.4217 429.9 0.9786 0.005 0.055 0.051 0.0084

216/I Ic 0.42 27.2 7.5 28986 5.1683 429.1 0.9780 0.005 0.056 0.050 0.0108

0.9560 0.012 0.121 0.056 0.0245

0.9673 0.009 0.100 0.055 0.0142

0.9740 0.007 0.067 0.051 0.0113

0.9734 0.007 0.070 0.051 0.0113

0.9747 0.007 0.066 0.051 0.0123

0.9761 0.006 0.065 0.049 0.0121

231/08b

216/Iid2

231/19b

231/05b

213115b

216/llb

0.40 31.3 -7.6 25843 5.8663 442.8

0.41 29.5 -3.8 28470 5.2565 430.6

0.40 29.4 -0.l 28182 5.2883 437,I

0.41 29.0 0.0 32536 4.3580 417.4

0.41 29.9 0.2 35287 3.8205 414,6

0.39 29.3 3.5 32229 4.3654 414.0

243/25a I 0.40 38.3 4.I 30634 4.7248 432.I 0.9654 0.010 O.109 0.057 0.0177

241/lOb 0.40 37.7 4.5 29065 5.0855 439.1 0.9681 0.010 0.096 0.057 0.0156

0.9529 0.013 0.132 0.061 0.0237

0.9556 0.013 0.137 0.058 0.0179

0.9620 0.012 0.I12 0.058 0.0177

0.9618 0.011 O.112 0.059 0.0163

0.9666 0.011 O.103 0.058 0.0164

0.9665 O.Oll 0,I01 0.057 0.0155

0.40 40.I ---4.5 34068 4.0367 414.2

0.39 39.0 --4.0 28141 5.2741 441.6

0.40 39.2 --0.3 36178 3.6499 403.2

0.40 39.8 0.0 32800 4.2887 413.0

0.40 38.9 1.2 33550 4.1332 415.0

0.39 39.2 3.5 31165 4.6021 424.3

243/25c I

241/10c

214/2ab3

233/04b

23ad08b

235/08c

Inlet Recovery

Figures 9(a) and 9(b) show the effect of AOA and AOSS on inlet pressure recovery. An inlet

recovery value of 1.0 indicates that the total pressure at the AIP is equal to the free-stream total

pressure. This is the ideal value and indicates 100-percent pressure recovery at the AIP. Negative

AOSS have an obvious detrimental effect on inlet recovery for all AOA at Mach 0.3 and 0.4. As

AOA increases, the detrimental effect of AOSS becomes increasingly pronounced. Pressure recovery remains constant at approximately 97 percent from 4 ° to 30 ° AOA. At increases above 30 °, the

loss in pressure recovery becomes more sensitive to AOA. At Mach 0.3, 60 ° AOA, 0 ° AOSS, the

pressure recovery is less than 91 percent (fig. 9(a)). At Math 0.4, -10 ° AOA, pressure recovery

levels are less than the recovery at 4 ° to 40 ° AOA and 0 ° AOSS (fig. 9(b)). Inlet recovery levels

are highest in the 4 ° to 20 ° AOA range.

Turbulence

Figure 10 shows the effect of AOA and AOSS on turbulence levels. The detrimental effects of

AOA and AOSS on turbulence are similar to those shown for inlet recovery. Comparing figure 9

with figure 10 shows that inlet recovery decreases as the flow turbulence increases. Going from

nose-left to nose-right (positive to negative AOSS) increases turbulence. This AOSS trend is

slightly more pronounced at Mach 0.4 than at Mach 0.3. The turbulence level at Mach 0.3 is

approximately 0.005 from 4 ° to 20 ° AOA. The turbulence level begins to increase above 20 ° AOA.

At 60 ° AOA, turbulence has increased to 0.022. At Mach 0.4, the turbulence level at -10 ° AOA is

close to that at 40 ° AOA. Turbulence levels are lowest in the 4 ° to 20 ° AOA range.

16
PeakDynamicCircumferentialDistortion

Figure 11showstheaffectof AOA andAOSSonthepeakdynamiccircumferentialdistortion.

As a function of AOA, thesetrendsaresimilar to thoseof turbulence.Thepronouncedaffect of

AOSSis evidentfor all AOA at Mach0.3and0.4.As theaircraftmovesmorenose-left(positive

AOSS),thepeakdynamiccircumferentialdistortiondecreases.From 4° to 30° AOA, this distortionincreasesslightly; above30° AOA, it increasesrapidly.For Mach0.3,this distortiondoubles

from approximately0.08at 30° AOA to 0.17at60° AOA. At Mach0.4and0° AOSS,decreasing

AOA from 4° to -10° increasesthe peakdynamiccircumferentialdistortion from 0.04 to 0.11.

Peakdynamiccircumferentialdistortionlevelsarelowestin the4° to 20° AOA range.

PeakDynamicRadialDistortion

Figure 12showsthatAOA andAOSShavelittle affectonpeakdynamicradialdistortionupto

40° AOA. At Mach 0.3, the minimum valueof peakdynamicradialdistortion wasobtainedat

30° AOA (fig. 12(a)).IncreasingAOA from40° to 60° causesthepeakradialdistortionto increase

rapidly.IncreasingAOA to50°and60° alsoincreasesthesensitivitytoAOSS,similarto theAOSS

effectpreviouslydiscussedin this section.At Mach0.4 andfrom 4° to 30° AOA, thepeakradial

distortion level is nearlyconstantat approximately0.055 (fig. 12(b)).At -10° AOA, the peak

radialdistortiondecreasesto approximately0.03.Thistrendistheoppositeof theoneobtainedfor

thecircumferentialdistortiondata.

Peak-to-PeakPlanarWave

Figure 13showsthepeak-to-peakvarianceof theplanarwavedescriptor.Effectsof AOA and

AOSSonthe peak-to-peakvaluesof aplanarwavearesimilar to thetrendsdescribedfor turbulenceanddistortiondescriptordata.AsAOA increases,variabilityincreasesfor agivenAOSS.For

instanceat Mach0.4, 30° AOA, -7.5° AOSS,a higherpeak-to-peakvalueof nearly0.024 was

obtained,comparedwith 0.011at theother30° AOA points(fig. 13(b)).Thepeak-to-peakplanar

wavelevelsarelowestin the4° to 20° AOA range.Theplanarwavevaluesat thesehigh-power

conditionswereextremelylow.

AIPTotalPressureCharacteristicsatMach0.3

Figures14to 16showtimehistoriesof themaneuverflight conditions,isocontoursof thetimeaveragedPTE pressures,andtime historiesandpowerspectraldensityfunction(s)(PSD)of the

PTKinletrecoveriesat locations2A, 2D,6A, and6D.Theselocationswerechosento showacross

sectionof thevariationininletrecovery,turbulencelevels,andPSD'sat theAIP. Thecontoursare

viewedfrom anALF perspective.

10° AOA and0° AOSS

Figure 14showstheAIP totalpressurecharacteristicsandmaneuverdescriptionfor 10° AOA,

0° AOSS,andMach0.3.Timehistoriesof Machnumber,AOA, AOSS,andaltitudeareshownin

figure 14(a).Figure 14(b)showsisocontoursof time-averagedPTEtotalpressures.Themajority

17
of thepatternisuniform with a smalllow-pressureregiondevelopingin theupperinboardregion.

Theprobelocations2A, 2D,6D, and6A arehighlightedasareferencefor thePTK inletrecovery

time histories(fig. 14(c))andPSD's (fig. 14(d)).Probelocations2D and 6D arein the highpressureregion(2percent)and2A and6A arein alow-pressureregion(-4 percent).This canalso

be seenin thePTK inlet recoverytime histories(fig. 14(c)).The 2D and 6D locationsarenear

100percentrecovery;whereas,locations2A and6A arebetween90-percentand93-percentrecovery.Thevery low activity of thetimehistoriesatlocations2Dand6Dexhibitsignificantlyreduced

turbulencein comparisonwith locations2A and 6A. The turbulencevalues(TU) areshownin

figure 14(d).

Figure14(d)showsPSD'scorrespondingto thefigure 14(c)timehistories.At locations2D and

6D, a frequencypeakcanbe seenat approximately215Hz. This frequencyis equivalentto the

1-per-revolutionbladepassagefrequencyof thefan.Peakscanbeseenattheselocationsagainst

thelow-turbulentairflow background.The 1-per-revolutionfrequencyis notseenatlocations2A

and6A becauseof increasedbackgroundairflow turbulencelevels.

30° AOA and 0 ° AOSS

Figure 15 shows the effect of increasing AOA to 30 ° on the AIP total pressure characteristics.

Time histories of the maneuver aerodynamic conditions are shown in figure 15(a). The isocontours

of the time-averaged PTE pressures are shown in figure 15(b). Note the change in pattern from

10 ° AOA. The lowest pressure region is now exhibited in the lower inboard region. The highpressure region is predominately in the upper and outboard regions of the inlet. Note the increase

in the circumferential distortion descriptor DP/PC (0.061 at 30 ° AOA versus 0.045 at 10 ° AOA)

and the decrease in the radial distortion descriptor DP/PR (0.037 at 30 ° AOA versus 0.048 at

10 ° AOA). The probe locations 2A, 2D, 6D, and 6A are highlighted as described in the "10 ° AOA

and 0 ° AOSS" section.

Figure 15(c) shows time histories of the PTK inlet recoveries for highlighted probe locations.

Location 6D now shows lower inlet recovery and higher turbulence when compared with

10 ° AOA. This corresponds to figure 15(b) where location 6D is in the low-pressure (-2 percent)

region. Location 2D remains in the high-pressure region where inlet recovery is nearly 100 percent,

and turbulence is low in comparison with the other locations. Turbulence at location 2D is slightly

greater at 30 ° AOA than at 10 ° AOA.

Figure 15(d) shows PSD's of the figure 15(c) time histories. A 1-per-revolution frequency

(215 Hz) can best be seen at location 2D. Turbulence at location 2A has increased enough to mask

the 1-per-revolution frequency. At location 2D, and to a lesser extent location 2A, another frequency peak appears at approximately 55 Hz. The cause of this peak is currently unknown.

60 ° AOA and 0 ° AOSS

Figure 16 shows the effect of increasing AOA to 60 ° on the AIP total pressure characteristics.

Time histories of figure 16(a) show increased variation in the aerodynamic conditions while trying

to maintain 60 ° AOA. The flight condition tolerances could only be met for 4 sec. Isocontours, of

the time-averaged PTE pressures show a pronounced pressure gradient across the face of the inlet

18
(fig. 16(b)).The low-pressureregionoccupiestheinboardhalfof theinlet faceandextendsalong

thetop andbottomoutboardwalls. Themaximumlow-pressureregion(-10 percent)is alongthe

mid-inboardwall; whereas,themaximumpositivepressureregionis anislandaboutlocations3C

and3D.Thecircumferentialdistortiondescriptor,DP/PC,increasedto 0.103,andtheradialdistortion descriptor,DP/PR,increasedto 0.057.

Timehistoriesof theinletrecoveriesof locations2A, 2D,6A and6Dareshownin figure 16(c).

All timehistoriesshowtheincreaseinturbulencelevelat60° AOA. Theinletrecoveryrangesfrom

approximately80percentat location6A to approximately96percentat location2D.

The PSD'sfrom the PTK-recoverytime historieshaveelevatedenergylevelsacrossthe frequencyrange(fig. 16(d)).Theseelevatedlevelsmaskoutthefrequencypeaksthatwereobtained

atAOA conditionsof 30° andbelow.

Matrix ofAIP TotalPressureContours

The effectsof high-AOA conditionson the inlet arevividly shownby the seriesof totalpressurecontourpatternsin figure 17.Thesepatternsarebasedontime-averagedinlet rakemeasurementsandsummarizetheAIPtotalpressurepatternsobtainedfor arangeof aerodynamicconditions.Thepatternsarearrangedin ascendingorderofAOA from -10° atthebottomto 60° atthe

top of thepage.TheAOSSsequencegoesfrom negative(nose-right)to positive (nose-left)and

from left-to-right respectivelyin the figure.The AOSSrangeis approximately-8.3° to 8.8° at

30° AOA. IncreasingtheAOA from 10° to 60° resultsin anincreasein sizeof the localpressure

deficit in the lower inboardregionandexhibitsa morepronouncedpressuregradientacrossthe

enginefacefor all AOSS.Also at 30° AOA, thepositiveAOSS(nose-left)patternsaresimilar to

the0° AOSSpatterns.Thepatternsfor negativeAOSSshowtheeffectof theflow field off theleewardsideof thevehicle.This AOSSeffectis mostobviousat 10° and30° AOA wherethechange

in AOSSis greatest.Thelow-pressureregionencroachesfurtheracrossthefaceof theengine,and

thehigh-pressurecenterareadecreases.Above30° AOA, AOSShasdecreasedaffectonthecontourpatterns.This decreaseresultsfrom thefuselageforebodyhavingprogressivelylessinfluence

ontheflow enteringtheinlet.Thecharacteristicsof thecontourpatternforthenegativeAOA case

aresignificantlydifferentfrom thoseatpositiveAOA. This differenceis mostlikely causedby the

airflow off theLEX andfuselageforebodyenteringtheinlet.

AIP Wall StaticPressureCharacteristicsatMach0.3

To betterunderstandtheflow characteristicsattheinlet entranceandsubsequentdownstream

effectsatthe AIP, the inlet lip staticpressureswereevaluated.The effectof AOA andAOSSon

the AIP wall staticpressuresis shownin figure 18.ThePSEpressuresweretime averaged,referencedto thefree-streamtotalpressure,andobtainedatMach0.3.Figure18(a)showsthewall static

pressurelocationswith referenceto theinletrake.Theangularpositionof 0° referencesthetopof

theinlet duct.Theorientationof figure 18(a)is ALF.

Theeffect of AOA onPSEsat 0° AOSSis shownin figure 18(b).The angularpositionfrom

12° to 323° correspondsto staticpressurelocations8A to 8H. IncreasingAOA from 10° to 60°

19
causesthe minimum pressureto rotatefrom the 189° position(locationE) to the 234° position

(locationF).Thiscorrelateswell with therotationof therakelow-pressureregionasAOA increases in figure 17. Figure 18(19) shows that by increasing AOA from 10 ° to 30 °, a small decrease in

pressure occurs at locations D, E, and F. Increasing AOA to 50 ° the pressure decreases markedly

at all locations. As AOA increases to 60 °, the pressure decreases between 3 to 5 percent for all

locations compared with the 50 ° AOA data.

Figure 18(c) shows the affect of AOSS on the PSE pressures at 30 ° AOA. Increasing AOSS

from 0 ° AOSS to a nose-left position of 8.8 ° shows little change in static pressures. As the airplane

moves nose-right to an AOSS of-8.3 °, the decrease in pressure is small with the minimum pressure

at locations D and E. This minimum pressure correlates well with figure 17. The lowest pressure

region of the rake total pressures, at 30 ° AOA and -8.3 ° AOSS, is near the base of rake leg 4. Leg 4

is between wall locations D and E.

Inlet Lip Surface Pressure Characteristics at Mach 0.3

Figure 19 shows the affect of AOA on inlet entrance pressure coefficients for 0 ° AOSS and

Mach 0.3. Pressure coefficients at 10% 30% 50% and 60 ° AOA are shown. These data are from the

same maneuvers as those presented in figure 18. Figures 19(a)-I 9(d) show four rows (circumferential rows A, C, E, and F) of pressure coefficients as a function of the distance from the inlet tip

in inches. Figure 6 shows close-ups of the inlet lip transducer locations. The vertical line on each

of the figures at 0.0 in. represents the inlet tip. A positive distance from the inlet tip denotes

transducer locations further downstream along the inner wall of the inlet duct. Negative distance

denotes downstream locations along the outside wall of the inlet duct.

A pressure coefficient of 1.00 represents the stagnation point of the flow as it impinges on the

inlet lip. The horizontal line across each of the data figures represents the pressure coefficient

corresponding to the local speed of sound. At free-stream Mach 0.3, the pressure coefficient is

-6.947. At this flight condition a pressure coefficient more negative than -6.947 indicates local

supersonic flow.

Figure 19(a) shows pressure coefficients at the top of the inlet duct. At -0.75 in. on the outer

skin the pressure coefficient is near 1.0 for all AOA. This location is in close proximity to the stagnation point of the airflow entering the inlet. At 0.938 in. inside of the inlet tip, the local flow

becomes supersonic at all AOA's except 60 °. At 10 ° AOA and 0.938 in. inside of the inlet tip, the

minimum pressure coefficient of -9.255 was obtained for the top of the inlet duct.

At the outboard side of the inlet duct, or at the 90 ° position (fig. 19(b)), the pressure coefficients

exhibit trends similar to what is seen in figure 19(a) as AOA increases. Comparing the side of the

duct with the top of the duct, the pressure coefficients at the inlet tip when at 50 ° and 60 ° AOA are

slightly more negative than at the top of the duct for the same conditions. Only the 10 ° AOA point

0.688 in. inside from the inlet tip is clearly supersonic. At this AOA and location, the minimum

pressure coefficient of-7.950 was obtained. The other AOA conditions were near to sonic conditions at this location.

20
Figure 19(c)showspressurecoefficientsat thebottomof the duct.Additional rowsof static

pressuremeasurementsarepresentedfor the 7.75in. and 11.75in. locationsdownstreamof the

inlet tip. The 11.75in. locationis neartheinlet throat.Thevariationin pressurecoefficientsas

AOA increasesfor all rowsis moresignificantatthebottomof theductthanatthetop or outboard

sideof theinletentrance.Onlythe 10°AOA conditionshowsstagnatedflow 0.75in. outsideof the

inlet tip. At higherthan10° AOA, thestagnationpoint movesmoredownstreamon theoutersurfaceof the inlet. From theinlet tip inwardto the 1.75in. row, the pressurecoefficientsindicate

supersonicflow for allAOA conditionsexcept10°.At the7.75in. and11.75in. locations,thepressurecoefficientsareincreasing,indicatingimprovedpressurerecoveryastheairflow movesdownstreamin theinlet duct.

Figure 19(d) shows the pressure coefficients at the wing and cowl junction. The variation in

pressure coefficients as AOA increases (from the inlet tip to the 2.188 in. location) is the greatest

at the wing and cowl junction. Only the 10 ° AOA condition shows stagnated flow 0.75 in. outside

of the inlet tip. From the inlet tip inward to the 4.25 in. location, the pressure coefficients indicate

supersonic flow for all AOA conditions except 10 ° and from 30 ° at the inlet tip. At 60 ° AOA, the

pressure coefficients at the inlet tip to the 4.25 in. location show a fiat profile which signifies separated flow. This is the condition where separated flow is most apparent.

CONCLUSIONS

Flight tests of the F/A- 18 HARV were conducted to study effects of steady-state and dynamic

pressure phenomena on inlet performance during steady aerodynamic conditions. Aerodynamic

interface plane (AIP) total pressures, inlet entrance, duct, and AIP wall static pressures were measured for aircraft Mach 0.3 and 0.4 at various angle of attack (AOA) and angle of sideslip (AOSS)

conditions. Analysis of this database has yielded the following conclusions:

. Pressure measurements obtained at 30 ° AOA, 0 ° AOSS showed consistent results over the

course of the flight test program. The inlet total pressure contours and inlet distortion

descriptors were repeatable. Consistent results during steady aerodynamic conditions are

essential for providing confidence in the analysis of inlet distortion during rapid aircraft

maneuvers. In-flight pressure calibrations and postflight instrumentation bias corrections

were essential for obtaining accurate and repeatable results.

. At 0 ° AOSS, increasing AOA from 0 ° to 20 ° had little effect on time averaged inlet recovery. As AOA increased to 30 °, inlet recovery began to decrease. As AOA was increased

from 30 ° to 60 ° , the loss in inlet recovery becomes more pronounced. The trend of the loss

in inlet recovery at increasing AOA was similar to the trend of higher levels of turbulence

and circumferential distortion at increasing AOA. The highest levels of inlet recovery

occurred in the AOA range from 4 ° to 20 ° and coincided with the lowest turbulence and

distortion levels.

. At positive AOSS (nose-left aircraft position), the right-hand inlet was most efficient at all

AOA. At 5 ° AOSS, the airflow has a more direct, streamlined path through the inlet with

little fuselage interactions as compared to nose-right. By rotating the aircraft nose-right

21
o

°

(increasingly negative AOSS) while maintaining AOA, the inlet recovery, turbulence, and

circumferential distortion descriptors were adversely affected. This adverse affect becomes

more evident as AOA increases. The airflow becomes increasingly turbulent as a result of

its interactions with the fuselage as evidenced by the separation of the inlet lip pressures.

The total pressure contour patterns show that increasing the AOA from 10 ° to 60 ° results in

an increase in the size of the local pressure deficit in the lower inboard region and exhibits

a more pronounced pressure gradient across the engine face for all AOSS. The positive

AOSS (nose-left) patterns are similar to the 0 ° AOSS patterns. The patterns for negative

AOSS show an adverse effect on the flow field off the leeward side of the vehicle. The

characteristics of the contour pattern for negative AOA cases differs significantly from

those of positive AOA. This difference is most likely caused by the airflow off the LEX and

fuselage forebody entering the inlet.

At the top of the inlet duct, pressure coefficients obtained at all AOA conditions indicate

stagnated flow near 0.75 in. outside of the inlet tip. Stagnated flow is also indicated at the

side of the duct for all AOA except the 60 ° condition. For the bottom of the duct and the

wing and cowl junction, only the 10 ° AOA condition shows stagnated flow 0.75 in. outside

of the inlet tip. Supersonic flow is most evident at the bottom of the inlet duct and at the

wing and cowl junction. Pressure coefficients from the inlet tip (0.0 in. location) to the

4.25 in. location indicate supersonic flow at all AOA's except for 10 ° AOA. As the airflow

travels downstream, pressure recovery is apparent. At the wing and cowl junction, the fiat

pressure coefficient profile at 60 ° AOA indicates separated flow.

22
APPENDIXA

INLET FLOW DESCRIPTORDEFINITIONS

For a given patternof 40 total pressuremeasurements(8 rakes,5 rings), the total pressure

distortiondescriptorsaredefinedasfollows:

MaximumCircumferentialDistortionDescriptor- DP/PCmax

This circumferentialring pressuredistortionindexis basedonthecalculationof DP/PCringi,

wherethe subscript"ring i" refersto anyof thefive pressurerings.

Theringsarecountedin ascendingorder,outermostto innermost.

DP/PCringi ---- { [PAVring i - PMINring i] / PAV} for rings i = 1,5

where

PAVring i is the area-averaged total pressure of ring i

PMINring i is the minimum total pressure in the largest low-pressure region in ring i

PAV is the area-averaged total pressure at the engine face

DP/PC max is defined as the largest of

0.5 * [DP/PCring i -t-DP/PCring i+l ] for i = 1,4

Maximum Radial Distortion Descriptor - DP/PR max

Similar to DP/PC, DP/PR is calculated for each of the five rings as follows:

DP/PCring i = [PAV - PAVring i ] / PAV for rings i = 1,5

where

PAV is the area-averaged total pressure over the complete face

PAVring i is the area-averaged total pressure of ring i

23
If aDP/PRring valueis negative,it is assumedto bezero.DP/PRmaxis definedasthelargerof

DP/PRringI or DP/PRring5"If DP/PRmaxislocatedinring I, thedistortionishubradial.If DP/PR

maxis locatedin ring 5, thedistortionistip radial.

PlanarWavePeak-to-Peak

Thesteadinessof theinlet recoverywith time is quantifiedby theplanarwavepeak-to-peak

parameter.First, the AIP average total pressure is computed at each time slice (PAV) and then

averaged over a period of interest (PAV):

PAV =

n

1/n _ (PTKi)

i=1

where PTK i represents individual PTK pressure measurements and n is the number of totalpressure probes; therefore,

PAV =

m

1/m ,_, (PAVj)

j=l

where m is the number of data samples.

Then, the maximum value minus the minimum value in this period divided by the average AIP

total pressure is defined as the planar wave peak-to-peak value. It can be expressed in equation

form as follows:

PW(pk-pk) = [max(PAY) - min(PAV)] / PAV

Turbulence

The inlet turbulence parameter is the average of the root-mean-square of the total pressure

fluctuations at each PTK divided by the AIP time-averaged total pressure.

TU- 1/n 1/m [ (PT Kj - P-'A-'V)/ (_"-V

i=l k j=!

where n is the number of total pressure measurements included in the average, and m denotes the

number of data samples.

24
Recovery

Inlet recoveryis describedasthe averagetotal pressureat the AIP normalizedto the freestreamtotal pressure.For the HARV Inlet Program,the inlet recoveryvalue (REC) addsthe

bias-correctionsdescribedin the"In-Flight Calibration"section,andis time averagedfor a given

datasegment.

REC -

rn

1/m _ (PAV/PT0) i

i=l

where m denotes the number of data samples, and PT0 is the free-stream total pressure, psia.

Inlet Lip Pressure Coefficients

Pressure coefficients of the static pressure measurements about the inlet lip are calculated by

taking the difference of a measured PSE static pressure and free-stream static pressure and dividing

the result by the free-stream dynamic pressure. All pressure coefficients presented are time averaged for each data segment.

m

(PSE-PS0_

Cp = 1/m _ \ Q'B-'A'-R" )i

i=l

where m is the number of data scans, PS0 is free-stream static pressure, and QBAR is the freestream dynamic pressure.

To determine whether a local static pressure measures supersonic conditions, the pressure

coefficient corresponding to local speed of sound, Cp*, is calculated and is time averaged for each

data segment.

Cp* =- mf( lI( 1t

1/m _ 2 ()'- 1)M2+ 2 3.5

i=1 (_+i) -1

25
REFERENCES

_Regenie, Victoria, Donald Gatlin, Robert Kempel, and Neil Matheny, The F-18 High Alpha

Research Vehicle: A High-Angle-of-Attack Testbed Aircraft, NASA TM-104253, Sept. 1992.

2Yuhas, Andrew J., Ronald J. Ray, Richard R. Burley, William G. Steenken, Leon Lechtenberg,

and Don Thornton, Design and Development of an F/A-18 Inlet Distortion Rake: A Cost and Time

Saving Solution, NASA TM-4722, Oct. 1995.

3yuhas, Andrew J., William G. Steenken, and John G. Williams, "F/A-18A Inlet Flow Characteristics During Maneuvers with Rapidly Changing Angle of Attack," presented at 5th HighAngle-of Attack Technology Conference, Langley Research Center, Hampton, Virginia,

Sept. 17-19, 1996.

4Steenken, William G., John G. Williams, and Andrew J. Yuhas, "An Inlet Distortion Assessment During Aircraft Departures at High Angle of Attack for an F/A- 18A Aircraft" presented at

5th High-Angle-of-Attack Technology Conference, Langley Research Center, Hampton, Virginia,

Sept. 17-19, 1996.

5Williams, John G., William G. Steenken, and Andrew J. Yuhas, Estimating Engine Airflow in

Gas-Turbine Powered Aircraft with Clean and Distorted Flows, NASA CR-198052, Sept. 1996.

6Norby, William E and Andrew J. Yuhas, Dynamic Distortion Prediction Using a Combined

CFD/Distortion Synthesis Approach, NASA CR-198053, Sept. 1996.

7"Gas Turbine Engine Inlet Flow Distortion Guidelines," Society of Automotive Engineers

Aerospace Recommended Practice, ARP 1420, Mar. 1978.

8"Inlet Total-Pressure-Distortion Considerations for Gas-Turbine Engines," Society of Automotive Engineers Aerospace Information Report, AIR 1419, May 1983.

9GE Distortion Analysis Program User's Manual, Version III, General Electric Aircraft

Engines, Evendale, Ohio, TM88-352, Sept. 1988.

l°Model Specification for F404-GE-400 Turbofan Engine, Specification No. CP45K0006, General Electric Aircraft Engines, Evendale, Ohio, Nov. 1975, Reprinted Feb. 1983.

llAmin, N.E, C.J. Richards, E.G. de la Vega, and M.A. Dhanidina, F/A-18A Engine Inlet

Survey Report, vol. I of HI, Northrop Corporation, Aircraft Division, Hawthorne, California,

NOR 81-316, Nov. 1981.

12Morse, D.B., N.E Amin, EW. Marxen, J.A. McGuire, E.G. de la Vega, and M. Yamada,

Propulsion System Functional and Performance Analysis Report, Northrop Corporation, Aircraft

Division, Hawthorne, California, NOR 77-364, July 1978.

13Moes, Timothy R. and Stephen A. Whitmore, A Preliminary Look at Techniques Used to Obtain Airdata From Flight at High Angles of Attack, NASA TM-101729, Dec. 1990.

14Ray, R.J., J.W. Hicks, and R.I. Alexander, Development of a Real-Time Aeroperformance

Analysis Technique for the X-29A Advanced Technology Demonstrator, NASA TM-100432,

May 1988.

26
lSThomton, D.A. and William G. Steenken, Summary of lnlet Distortion Rake Activities." Flight

Clearance Through High-Temperature Rake Delivery, General Electric Aircraft Engines, Cincinnati, Ohio, NAS 3-25951, Oct. 1995.

27
EC91495-15

Figure 1.NASA F/A-18A HARV aircraft(preproductionaircraftnumber6) with multiaxisthrustvectoringpaddles.

Engine gas and _ \

fuel system

press

40 high- and //J/>'_'..;(. J/"_x_"'__J J _

low-response JJ_ __j'_

total

Alrcralt Centerline

"-l--_.... 23.3 I/-vn,

L _ "l"..... _./ generators

- :'",L- _Lit IT...... _,

_" 1_3-- -- _ - "..... ' ,_i-,

-'--L-_..... i ....... rT--_so I

....... T---_:_.... ",1 I

, I_,

FS560 FS550.9 FS500 FS425 FS411 FS379

I I (throat) _L_ Two position

I i _ _ ramp bleed

i j I i "--"_ exitdoors

---; ............ ; I

Engine _,'l L "t .......

SJdILYklW face -_ L., Rake ....... "='=:-_:!t /Flow

_, I face Porous _:._] I

--J--I- ......... bleed plate -/1_i_i!] I

Figure 2. F/A-18A air induction system.

28
LPT discharge

measuring plane

LPC HPC LPT I

B

Exhaust

Inlet discharge discharge I nozzlel throat

Engine HPC Combustor I Afterburner Exhaust I

Engine inletl inletl discharte I Inletl nozzle Iinlet I

.,.,on__, _ _ " "r '_-_

! °

I,J. -'°q '"' L_

I_ SFUELG' liP" 212

m

PSI 1(311

Production

engine

40 sps

Additional

engine

25-40 sps

Engine

diagnostics

800 sps

Real-time

thrust method

25 sps

Figure 3. Right engine instrumentation used for obtaining HARV inlet research data.

29
EC94 42514-11

Figure 4. NASA and GEAE inlet pressure rake mounted in the HARV right inlet with engine

installed (forward-looking-aft orientation).

• Rake total pressures - PT

o Wall static pressures - PS

H 1_ewe

Duct static

position A

I

/

2

B

G

Inboard

3

7 C

F

6

D

/ s

E

Figure 5. HARV inlet rake and duct instrumentation locations (ALF orientation).

30
Transducer locations

0 Low response

• High response

.......... 2A •

EC93 42237-8

(a) Section A, 0% top of inlet duct entrance.

EC93 42237-5

(b) Section C, 90 °, outboard side of inlet

duct entrance.

EC93 42237-4

(c) Section E, 180 °, bottom of inlet duct

entrance.

EC93 42237-3

(d) Section F, 225 °, inlet and ramp junction

of inlet duct entrance.

Figure 6. Inlet lip transducer locations.

31
AOA,

deg

70 : i

55 data sets

60 ....................................... i...................................... _................... o............

L i o

i

s0- i i

i

i i ^o 9

30 oo

...... °'.",=° .......... L................................. T.....................................

' i

"°"°-'_°- _ ........................

o ii

, !

4o o i °

L _

0

20

10

0

-10

-15

..........................................................i..................................... ".................................. i.................................i..........................

...L.............. .o.. ............... _._i ............................................................................ ° ............. o...... i.....................................

! l o i i

-10 -5 0 5 10 15

AOSS, deg 970457

(a) Mach 0.3.

AOA,

deg

7O

60

5O

4O

30

20

10

0

-10

-15

24 data sets

o

..................................... i ............................. .Q .....

-10

v

i !

!

I

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

0 000 i

0

! o

0 .................................. 0 .............................. _ ........... "6...........................................................

L

o i

-5 0 5 10 15

AOSS, deg 97_58

(b) Mach 0.4.

Figure 7. Inlet data obtained during fixed-attitude maneuvers, altitude ranged from approximately

20,000 to 35,000 ft, corrected engine airflow was held constant at 144 lb/sec.

32
Inboard -

Low

pressure

region

I

i

AOA = 30.0 ° REC = 0.971 AOA = 30.10 REC = 0.969

AOSS = 0.125 ° TU = 0.007 AOSS = 0.002 ° TU = 0.007

MACH = 0.299 DP/PC = 0.061 MACH = 0.297 DP/PC = 0.062

HPL = 25405 ft DP/PR = 0.037 HPL = 25916 ft DP/PR = 0.037

(a) Flight 211. (b) Flight 213.

Outboard

I I

I I

AOA = 29.6 ° REC = 0.972 AOA = 30.1 ° REC = 0.970

AOSS = 0.158 ° TU = 0.007 AOSS = 0.171 ° TU = 0.007

MACH = 0.293 DP/PC = 0.058 MACH = 0.298 DP/PC = 0.060

HPL = 21584 ft DP/PR = 0.036 HPL = 25118 ft DP/PR = 0.037

(c) Flight 234. (d) Flight 241.

Figure 8. Contours of time-averaged high response total pressures showing data repeatability at

Mach 0.3, AOA 30 °, and AOSS 0 °. ALF orientation.

33
Inlet

recovery

1.00

.98

.94

.92

.50

.88

-15

_/_/_ 10°

n IB

,,_ 7_f, recovery Ic_is from 10 °

t_ 60 ° AOA

A

6( I°

-10 -5 0 5 10 15

AOSS, deg

970459

(a) Inlet recovery at Mach 0.3.

AOA

• _ 9 0

• 4 °

o 10 o

• 200

o 300

[] 40 °

A 500

• 570

X 600

1.00

.98

.96

Inlet .94

recovery

.92

.90

/L

i +

..--O

.88

-15 -10 -5 0 10 15

Leeward I AOSS, deg / Windward ___

97046o

AOA

• - 100

• 4°

o 10°

• 200

• 270

o 30 °

+ 38 °

[] 400

(b) Inlet recovery at Mach 0.4.

Figure 9. Effect of AOA and AOSS on inlet recovery during steady aerodynamic conditions.

34
Turbulence

.028

.024

.020

.016

.012

.OO8

.OO4

0

-15

i

i i A ,

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

i

-10 -5 0 5 10 15

AOSS, deg 97o461

AOA

• _ 9 °

• 4 °

o 10 °

• 20 °

o 30 °

,_ 40 °

A 50 °

• 57 °

x 60 °

(a) Turbulence levels at Mach 0.3.

Turbulence

.028

.02+ i +

| ! i

°o..................................... +..................................... ............................. +........................................................................

.016 i t

.012 ..................................................... _ ,_ j

-,+ t

.OO8

.OO4

I

,.,,+_.+.+................. i

AOA

• - 10 o

• 4 °

• 10 °

• 20 °

• 27 °

o 30 °

+ 38 °

,., 40 °

0 !

-15 -10 -5 0 5 10 15

Leeward -- AOSS, deg -- Windward

970462

(b) Turbulence levels at Mach 0.4.

Figure 10. Effect of AOA and AOSS on turbulence levels during steady aerodynamic conditions.

35
Peak dynamic

circumferential

distortion

.2O

.16

.12

.08

.04

x

AA

A A

[

$6°°I

°___

-15 -10 -5 0 5 10 15

AOSS, deg 97o*_

(a) Maximum dynamic circumferential distortion at Mach 0.3.

AOA

• - 9 °

• 4 0

o 10 °

• 20 °

¢ 300

u 40 °

A 500

• 57 °

x 60 °

.20 AOA

• - 10 °

* 4 °

o 10 °

• 20 °

• 27 °

¢ 30 °

+ 38 °

o 40 °

•" .................................................... i............................................................... i...................................................

i 4o° i i

Peak dynamic .12

circumferential

distortion

.08

illOO

............................i........................ i..................... T...................'.............

0

-15 -10 -5 0 5 10 15

Leeward _ AOSS, deg _ Windward

970466

(b) Maximum dynamic circumferential distortion at Mach 0.4.

Figure 11. Effect of AOA and AOSS on maximum peak dynamic circumferential distortion during

steady aerodynamic conditions.

36
Peak

dynamic

radial

distortion

.20

.16

.12

.O8

.04

0

-15

i

i

i

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

I

A

_L_

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

i

-10 -5 0

AOSS, deg

10

(a) Maximum dynamic radial distortion at Mach 0.3.

AOA

• _ 9 °

• 4o

o 10 °

• 20 °

o 30 °

a 40 °

• . 50 °

• 57 °

x 60 °

Peak

dynamic

radial

distortion

.2O

.16 ................................................................. i................................

.12 ...............................................................................................

.08 .............................................................................................

.04 ........................................................................ -..................................

0 _

................................ ,L.............................. . .....................................

-15 -10 -5 0 5 10 15

Leeward _ AOSS, deg _ Windward _-_

970464

AOA

• - 10 °

• 4°

o 10 °

• 20 °

• 27 °

¢ 300

+ 38 °

a 40 °

(b) Maximum dynamic radial distortion at Mach 0.4.

Figure 12. Effect of AOA and AOSS on maximum peak dynamic radial distortion during steady

aerodynamic conditions.

37
,036

,O32

wave

l_ak-topeak ,016

,012

,008

,004

0

-16

A X

.024 ...............................................................................

Planar ,020 .............................................................................................................

A

....................................................................................... 0 ............... •

o

................................................... _._

0

°

10 °

F 60°

,_...X........)C ...........................................................................................

A

0

-10 -5 0 5 10 15

AOSS, deg 9ro_7

(a) Planar wave peak-to-peak fluctuations at Mach 0.3.

AOA

• _9 °

• 4 o

o 100

• 20 o

o 30 °

o 40 °

A 50 o

• 57 °

x 60 °

................................... i...................................... • ....................................

.036 !

.032 ................................................................................................................

.028 ................................................................................................................

....................................................... 9.................. iD ..............................

.024

.020

.016

.012

.008 .....................................................................

AOA

• - 100

• 40

o 10°

• 200

• 270

o 30 °

+ 38 °

i i___._ o 40 °

Planar .............................................. i...................................... _...................................

wave +i _

peak-to- _ri-q;i ...................................... _.................................... peak "

0

..................................... '...................................... i .................................................. _& _"*"'_ ...................................... ........................................................... i_.....................................

................................... i..................................... +.....................................

0

-15 -10 -5 0 5 10 15

AOSS, deg 97o4e8

(b) Planar wave peak-to-peak fluctuations at Mach 0.4.

Figure 13. Effect of AOA and AOSS on planar wave fluctuations during steady aerodynamic

conditions.

38
0.26 1 .....

0 1 2 3 4 5

14 •

S2 ....

0 I 2 3 4 5

0 1 2 3 4 5

xlO3

24

0 1 2 3 4 $

Tlme,

(a) Time histories of aerodynamic conditions.

+ Denotes probe location

I

2A

/ / / "_ +

• .-,o '%' _,i_\

t / , + ' '_Li -I÷// . 0 ' " , iil-

/.i,_ ÷, ÷ ' Ill/

\ , I" q. i I, ,,'y, _ . , • ,,SK/

\ 4,,, .+ " i I

AOA= 9.92° REC = 0.974

AOSSB 0.S9° TU = 0.005

MACH,t 0.308 DP/PC n 0.045

HPL= 22638 ft DP/PR z 0.048

Outboard

(b) Isocontours of time-averaged PTE total pressures normalized to the face average pressure.

_ o.:_r-, ..... ..............- .... ,--7_,.........

7m"-"* -71

0,80_ .... 0.75 .... ..... :

o 1 2 3 4 5

.6 .......

4

-I0

-12

0 50 100 1_,0 200

019_ .........

o.90 i..........

0.80 _ .........

ons] , ....

0 I 2 3 4 5

-4. "

-12

0 SO I00 150 200

0.95

09o

0.85

0.80

0.75 1

0 1 2 3 4 5

-4

-6.... i ..... i_

-tO

-12

0 50 100 150 20D

1.00 : .......... i i

o 1 2 3 4 5

Tlmo,_c

(c) Time histories of PTK inlet recoveries

for rake locations 2A, 2D, 6D, and 6A.

.6•

-I0

-12

0 50 I00 150 2130

Frequency, Itz

(d) PTK PSD's for rake locations 2A, 2D,

6D, and 6A.

Figure 14. AIP total pressure characteristics at AOA 10 °, AOSS 0% and Mach 0.3.

39
i0=0.28 l

°-_'1 f 2A

71....

_ 32

2

0 I 2 3 4 5

T'm_e, sec

(a) Time histories of aerodynamic conditions.

AOA = 30.0 ° REC = 0.071

AOSS • 0.35 ° TU • 0.007

MACH • 0.299 DP/PC • 0.061

HPL = 25405 tt DP/PR • 0.037

-- Outboard

(b) Isocontours of time-averaged PTE total pressures normalized to the face average pressure.

1.00 .......

0.90 1

o._ 7

O.IK)

0.7_ : :

0 I 2 3 4 5

-4

,oI

0 50 100 150 200

0.75 ]

0 I 2 3 4 $

4 t TU = 0.0035

-12 I .... ; .... i

0 50 100 150 2O0

o..I .......

°:_"_ 1

A

'°t TU'0_05 "l

0 1 2 3 4 5 0 50 lOCi 150 200

n i lli_. iiIL,iJ.l.li*,u_L,IJ_h_ll, ti,,

0.80 ............. -ill"10t TU = 0,0137 "-' "v"_t"_tr

o

0 1 2 3 4 5 0 50 I,_l'l 150 2(10

Time, le¢ Frequency, Hz

(c) Time histories of PTK inlet recoveries

for rake locations 2A, 2D, 6D, and 6A.

(d) PTK PSD's for rake locations 2A, 2D,

6D, and 6A.

Figure 15. AIP total pressure characteristics at AOA 30 °, AOSS 0% and Mach 0.3.

40
O34o]2_ + Denotes probe location

z

030

_ 0.28

0.26

S8

56 -- ', _, I t I 8 _' " '

\ \ ,,_--_- 5o.- +_.H

- o X \ ",",_V'J_+ - X_/

o l 2 , + J "'..."-...." .... +---Z-....

_1o3

28

2"/_ AOA : 59.12 °

t,

26 AOSS : -0.31 °

1 MACH : 0.319

I

24 HPL = 26904 ft

0 ! 2 3 4 5 6

T_. s_

REC = 0.904

TU : 0.022

DP/PC : 0.103

DPJPR : 0.057

(a) Time histories of aerodynanfic conditions.

-- Outboard

(b) Isocontours of time-averaged PTE tot_ pressures normalized to the face average pressure

0 1 2 3

o / il2/

0 50 100 150 200

|

°-95_l]llt'llr'T_ff_'lvt'"l_ir_TIw!Irrlrrll'i _'* w"rrlprlm' Ivi,"r,rq'_

0.90 _ -- '_ _ "

0.85 ......

080 ......... o175 : .....

0 1 2 3

:::1,u.o.o, 1

0 Z_O 100 1_0 200

1.00 ............... ; ............

i .gs ....... i

_ 0.90 0J5

0.80

0.75

0 ! 2 3

I r -12_ r

0 JO I_ 150 200

1.00 .....

0.950 ....... . .............

0,75 t _ _ " " -- _ _ "

0 L 2 3

Th11_, NC

(c) Time histories of PTK inlet recoveries

for rake locations 2A, 2D, 6D, and 6A.

_ "10] TU = 0.0352 -1

-12

0 _0 i_ 150

FrlilUlm_, lllz

(d) PTK PSD's for rake locations 2A, 2D,

6D, and 6A.

Figure 16. AIP total pressure characteristics at AOA 60 °, AOSS 0 °, and Mach 0.3.

41
i

AOA 59.1 °; AOSS -2.2 °

n

AOA 59.1°; AOSS -0.3 °

b

AOA 59.1°; AOSS 1.9 °

i

i

AOA 49,7°; AOSS -4.40 AOA 49.40; AOSS -0.1 °

I

, '", + _) " . . .4','_

,:*., + • '

q

AOA 48.5°; AOSS 5.2 °

i

I

AOA 30.4o; AOSS -8.3 °

I

" * ° /'i _

AOA 30.1°; AOSS 0.2 °

I

I

AOA 31.0°; AOSS 8.8 °

AOA 11.0°; AOSS -7.9°

q

AOA 9.90; AOSS 0.8 °

p

__i,.:' • ." 0 ." ' ' '':'_,,,ili -

I

AOA 9.30; AOSS 8.3 °

Nose-right J

i

u

AOA -10.3°; AOSS 0.2 °

I Nose-left J

Figure 17. Contours of AIP total pressures for a range of AOA and AOSS.

42
0 °

8Ho r">'SA(12 °)

(32_

9 °)

.o(,83°) 2

• (98°)

8F _

81)

8E (144° )

(189 °)

(a) AIP wall-static pressure locations.

.84

.82

.80

.78 PSE8/

PTO .76

.74

S

AOA

• 10°

• 300

• 500

.72 -- x 600

.70 I

12 59

I I I I I I

98 144 189 234 283 323

Angular position, deg 970469

(b) Effect of AOA on ESP wall pressures at the AIP, AOSS 0%

.84 --

PSE8/

PTO

.82

.80

.78

.76 --

.74 --

.72

.70

12

AOSS

* 8.8 °

• 0.0 °

• - 8.3 °

I

59

I I I I I I

98 144 189 234 283 323

Angular position, deg 970470

(c) Effect of AOSS on ESP wall pressures at the AIE AOA 30 °.

Figure 18. Effect of AOA and AOSS on averaged wall static pressures at the AIP, Mach 0.3.

43
Cp

I I I

4 6 8

Distance from inlet tip, in.

(a) A-rows: 0 degrees; top of duct.

AOA

---o--- 10 °

--o-- 30 °

.... _..... 50 °

--_-. 60 °

I I I

10 12 14

970471

Cp

-13

-11

-9

-7

-5

-3

-1

Outer

-- well

Inner

wall

AOA

10 °

w--[_--. 300

.... ¢,.... 500

---x---- 60 °

Cp*

1

14

-- I!

I,Ji

J #

i. I

-- h,'

i.i ,

e I

0

, I

-2 2

I I I I I I

4 6 8 10 12 14

Dlatance from Inlet tlp, In.

970472

(b) C-rows: 90 degrees; side of duct.

Figure 19. Comparison of inlet lip surface pressures at AOA 10 °, 30 °, 50 °, and 60°; AOSS 0 °, and

Mach 0.3

44
Cp

-13

-11

-9

-7

-5

-3

-1

Outer Inner

-- wall wall AOA

_9[-._.--]D,.'- ---o--- 10 °

--_-- 30 °

-- i.... _--_ .... _,.... 50°

/"_'.. _-_. __><.._. 60 °

:_.,I _ ........ _.,.,

Cp, iij//_!i x---,x-........... ............ _..:....,:._x_&.:_: .........

-//L/

i<lI/ -- ._I

i(

-IA

, / I

-2 0 2

I I I

4 6 8

Distance from inlet tip, in.

Throat

location ---_

I Ii

10 12 14

970473

(c) E-rows: 180 degrees; bottom of duct.

Cp

-13

-11

-9

-7

-S

-3

-1

Outer

wall

-- !

Co*i/

_- 'l

i//

-- _l t

At

__ /

a/

1 n

-2

Inner

wall

.-ZI- -.

,-_B_'"'4

"-A...

/_" " ,, . _ - .-o- _ _'_,-_..._..:... :,.._...:,...,.,,,

/ ""X- ........ -)( --X-__,

¢[

I I I I I

2 4 6 8 10

Distance from Inlet tip, In.

AOA

100

--D-- 30 °

.... _ .... 50 °

--)<-- 60 °

Throat

I°_ti°_7

tl

12

I

14

970474

(d) F-rows: 225 degrees; wing/cowl junction.

Figure 19. Concluded

5 . _ _:.._
REPORT DOCUMENTATION PAGE FormAl)proved

OMB No. 0704-0188

PulSe reporting buraen for ha= co#ectmn of informalmn is eotJmite0 to average 1 hour per response, including the time (Or revmwmg m/tructoni, aearch*ng malting data Iourcei. gethermg anti

maintaining the _ltll needed, and ¢_mp_tmg lim:l mg the ¢olklclion of informlitbon. Send ¢ommlints fe_lrding this burOen utimlite of any other alC_lct of this collection of mformlitlon

including sugge=t_rlll for mduc_ I_s burden, to Wlmhinglon Headquartarl Services, Direclorate for Intormltictl C_eratJons and P,etx)rts, 1215 Jef_mmn Dov_ Higttway, Suita 1204, Arlington,

VA 2220_..430Q. Ind to the Office of Management lind Budget, Paperwork Reductmn Project (0704-0188), Washington, DC 20503.

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

April 1997 Technical Memorandum

4.TITLE AND SUb lit ,r S. FUNDING NUMBERS

Inlet Distortion for an F/A-18A Aircraft During Steady Aerodynamic

Conditions up to 60 ° Angle of Attack

5.AUTHOR(S)

Kevin R. Walsh, Andrew J. Yuhas, John G. Williams, and William G.

Steenken

7. F',-HI-OFIMING O_GANIZATION NAME(S) AND ADDRESS(ES)

Analytical Services & Materials, Inc.

Edwards, California

General Electric Aircraft Engines

Cincinnati, Ohio

I). SPONSORING/MONITORING AGENCY NAME(S) AND AD_,_ESS(ES)

NASA Dryden Flight Research Center

EO. Box 273

Edwards, California 93523-0273

529 31 04 00 37 00 F-18

5. PERFORMING ORGANIZATION

REPORT NUMBER

H-2173

10.8PONSORINGR4ONITORING

AGENCY REPORT NUMBER

NASA TM- 104329

NASA Contract NAS 3-26617

17. SECURITY CLA_h_iFiCATION 18. SECbm+_' Ct.ASi_'iCATK)N 19. SECb_¢ C_RCATION

oFRr_Om OFTHIS_ OFASmWCT

Unclassified Unclassified Unclassified

NS N 7540-01-28¢- =:,_0 Av:i,_ble,_,_-_theNASACw-,;6iforAoroSpacelrdon.natlon, 8OOEikridgeLandi_R_d,

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

15. NUMBER OF PAGES

50

15. PRICE CODE

AO3

20. uMrrATION OF ABSTRACT

Unlimited

i

Stanaa_ Form 298 (Rev. 2-89)

Pmombod by _._1_ _dZ_H e

298-102

_4.SU_ECT_.-,,,As

FA-18A aircraft, High angle of attack, Inlet distortion, Pressure distributions,

Steady aerodynamic conditions

The effects of high-angle-of-attack flight on aircraft inlet aerodynamic characteristics were

investigated at NASA Dryden Flight Research Center, Edwards, California, as part of NASA's High

Alpha Technology Program. The highly instrumented F/A-18A High Alpha Research Vehicle was

used for this research. A newly designed inlet total-pressure rake was installed in front of the starboard

F404-GE-400 engine to measure inlet recovery and distortion characteristics. One objective was to

determine inlet total-pressure characteristics at steady high-angle-of-attack conditions. Other

objectives include assessing whether significant differences exist in inlet distortion between rapid

angle-of-attack maneuvers and corresponding steady aerodynamic conditions, assessing inlet

characteristics during aircraft departures, providing data for developing and verifying computational

fluid dynamic codes, and calculating engine airflow using five methods. This paper addresses the first

objective by summarizing results of 79 flight maneuvers at steady aerodynamic conditions, ranging

from -10 ° to 60 ° angle of attack and from -8 ° to 11 ° angle of sideslip at Mach 0.3 and 0.4. These data

and the associated database have been rigorously validated to establish a foundation for understanding

inlet characteristics at high angl e of attack.

13. ABSTRACT (',=_.:,..;;-;;. 200 _,,,_,.,4)

UnclassifiedmUnlimited

Subject Category 07

12b. DISTRIBUTION CODE

11. SUPPLEMENTARY NOTES

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

Virginia, Sept. 17-19, 1996. This report is available on-line on the NASA Dryden Technical Report Server,

http://www.dfrc.nasa._ov/c_i-bin/DTRS/dtrs.pl

1211._, niUTION/AVAiLABlUTY STATEMENT