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AIAA 98-3713

Summary of Inlet Characteristics of the F/A-18A High Alpha Research Vehicle

Kevin WalshNASA Dryden Flight Research Center Edwards, California

William Steenken and John WilliamsGeneral Electric Aircraft EnginesEvendale, Ohio

34th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit

July 13–15, 1998 / Cleveland, OH

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SUMMARY OF INLET CHARACTERISTICS OF THEF/A-18A HIGH ALPHA RESEARCH VEHICLE

Kevin Walsh*

NASA Dryden Flight Research CenterEdwards, California

William Steenken† and John Williams‡

General Electric Aircraft EnginesEvendale, Ohio

Abstract

Effects of high-angle-of-attack flight on aircraft inletaerodynamic characteristics were investigated at NASADryden Flight Research Center as part of NASA’s HighAlpha Technology Program. The highly instrumentedF/A-18A High Alpha Research Vehicle was used for thisresearch. A newly designed inlet total-pressure rake wasinstalled in front of the right-hand F404-GE-400 engineto measure inlet recovery and distortion characteristics.Objectives included (1) determining the inlet total-pressure characteristics at steady high-angle-of-attackconditions, (2) assessing if inlet distortion issignificantly different between rapid angle-of-attackmaneuvers and corresponding steady aerodynamicconditions, (3) assessing inlet characteristics duringaircraft departures, (4) providing data for developingand verifying computational fluid dynamic codes, and(5) calculating engine airflow using four methods forcomparison with a reference method. This paperdescribes the results obtained from this investigation.These data and the associated database were rigorouslyvalidated to establish the foundation for understandinginlet characteristics at high angle of attack.

Nomenclature

AIP aerodynamic interface plane

ALF aft looking forward

AOA angle of attack, deg

AOSS angle of sideslip, deg

1American Institute of Aero

*Aerospace Engineer.†Aerospace Engineer.‡Aerospace Engineer.Copyright 1998 by the American Institute of Aeronautics and

Astronautics, Inc. No copyright is asserted in the United States underTitle 17, U.S. Code. The U.S. Government has a royalty-free licenseto exercise all rights under the copyright claimed herein for Govern-mental purposes. All other rights are reserved by the copyright owner.

CFD computational fluid dynamics

CROCP combined rate of change of aircraft motion, deg/sec

CVG compressor vane guide position, deg

D2 simple distortion intensity descriptor, e.g. (PTmax – PTmin)/PTavg

DFRC Dryden Flight Research Center, Edwards, California

DP/PC circumferential distortion descriptor

DP/PR radial distortion descriptor

DPRS predicted loss of stability pressure ratio

DPRSF predicted loss of stability pressure ratio for the fan

DPRSH predicted loss of stability pressure ratio for the compressor

FS fuselage station

GEAE General Electric Aircraft Engines, Evendale, Ohio

HARV High Alpha Research Vehicle

HATP High Alpha Technology Program

IGV inlet guide vane

INS inertial navigation system

L/DENG length per unit diameter of the engine

LEX leading-edge extension

MDA The Boeing Company (formerly McDonnell Douglas Corporation, St. Louis, Missouri)

PCM pulse code modulation

PSE low-response inlet static pressure, psid

PSK high-response static pressure, psia

PT total pressure, psia

nautics and Astronautics

PTavg average AIP total pressure, psia

PTmax maximum AIP total pressure, psia

PTmin minimum AIP total pressure, psia

PTE low-response total pressure, psid

PTK high-response total pressure, psia

PT0 free-stream total pressure, psia

PT2 fan inlet total pressure, psia

sps samples/sec

Introduction

Inlet pressure distortion affects on a propulsionsystem at high angle of attack (AOA) during steadyaerodynamic conditions, rapid aircraft maneuvers, andaircraft departures are not thoroughly understood. Ateam of NASA and industry researchers was formed aspart of the NASA High Alpha Technology Program(HATP) to investigate the inlet characteristics, inlet andengine compatibility, and prediction methodologies athigh-AOA conditions. This effort addressed questionsthat have arisen during past subsonic aircraftdevelopment programs. These questions included thefollowing subjects:

• How do the inlet total-pressure characteristics, suchas inlet recovery, circumferential and radialdistortion, planar wave, and turbulence, behave as afunction of AOA, angle of sideslip (AOSS) andMach number?

• How do inlet distortion levels during rapid,high-AOA maneuvers compare with distortionlevels at corresponding steady aerodynamic flightconditions?

• What are the characteristics of the inlet flow duringaircraft departures? What factors lead to enginestalls that have been experienced during aircraftdepartures? Are there any other significant factorsbeyond inlet-induced distortion which account forengine stalls?

• Can computational fluid dynamics (CFD)technology be used to accurately predict inletcharacteristics at high-AOA conditions?

• Are the engine airflow pumping characteristicsaffected at extreme aircraft maneuver conditions?What is the best method for measuring airflow inflight at these conditions?

To address each of these questions, flights to mapinlet characteristics as a function of AOA, AOSS, and

Mach number were conducted using the F/A-18A HighAlpha Research Vehicle (HARV). Inlet data obtained atsteady aerodynamic conditions formed the foundationfor this inlet research effort.

The HARV aircraft, flown at the NASA Dryden FlightResearch Center (DFRC), Edwards, California,provided the ideal platform for the controlledexploration of inlet characteristics related to highly agilefull-scale vehicles.1 The thrust-vectoring vane systemprovided the ability to maintain steady, high-AOAconditions. The aircraft was highly instrumented withemphasis on the region around, and in, the starboardinlet and engine. A newly developed, 40-probe, total-pressure, inlet rake was installed directly ahead of theengine.2 Surface static-pressure transducers wereinstalled at the inlet rake location and around the inletlip. High-frequency response instrumentation wasinstalled to monitor engine operation and behavior. Tomaintain the quality of the high-response data, a seriesof specific instrumentation calibrations was performedon the ground and in flight during the test program.

This paper summarizes the inlet total pressure data,such as recovery, turbulence, and distortion obtainedduring the HARV inlet program. Data are presented forthe specific aircraft maneuvers flown, such as steadyaerodynamic conditions,3 rapid changes in AOA andAOSS,4 and aircraft departures.5 The HARV flight testdata were also used to evaluate and improve a CFDapproach for predicting dynamic distortion6 and toestimate engine airflow with clean and distortedairflows.7

Use of trade names or names of manufacturers in thisdocument does not constitute an official endorsement ofsuch products or manufacturers, either expressed orimplied, by the National Aeronautics and SpaceAdministration.

HARV Inlet Program Objectives

The HARV inlet program objectives were developedfrom questions raised from previous inlet researchprograms. To understand these objectives, a historicalperspective of the traditional methods for obtaining inletdistortion and its subsequent influence on inlet–enginecompatibility is presented. Then, a brief summary ofeach part of the inlet research program is provided.

Historical Perspective

In developing an air-breathing propulsion system,designers seek a high degree of aerodynamiccompatibility between the inlet and engine. This

2American Institute of Aeronautics and Astronautics

compatibility is especially true for high-performancecombat aircraft with high-AOA-maneuveringrequirements. The current state-of-the-art process toachieve inlet–engine compatibility involves severalextensive inlet and engine test programs and closeinteraction between the airframe and engine companies.These tests separately determine inlet distortiongeneration and engine distortion tolerancecharacteristics. Understanding inlet distortion levels isneeded at the earliest possible stage of the designprocess to perform trade studies.

Inlet–engine compatibility is established bydetermining the worst-case level of inlet total-pressuredistortion to which an engine will be subjected. Theimpact of this distortion on the aerodynamic stability ofthe engine is then assessed. In general, this assessmentis accomplished by testing inlet-forebody scaled modelsin a wind tunnel at fixed attitudes covering a range ofinlet airflows, Mach numbers, AOA, and AOSS that areprojected to be encountered by the aircraft during itsservice life. The resulting time-variant, total-pressure,distortion patterns that cause the maximum loss incompression component stability margins are thendetermined. Compression components are designed tohave sufficient stability margin to cover both theexternal destabilizing influence of total-pressuredistortion determined from the wind-tunnel data and theinternal destabilizing influences, such as enginevariability, deterioration, control tolerances, and throttletransients. This technique successfully providescompatible inlet–engine systems on many commercialand military propulsion systems.

Inlet compatibility engineers have long beenconcerned that the maximum total-pressure distortionlevels determined during fixed-attitude testing in windtunnels may not represent the maximum total-pressuredistortion levels encountered during maneuvering flight.Two fundamental concerns existed. First, is themaximum level of distortion obtained from fixed-attitude testing adequate for predicting the maximumtotal-pressure distortion obtained at similaraerodynamic conditions during maneuvering flight?Second, is it possible that unsteady flow disturbancesassociated with the fuselage forebody, protuberances, orboth, would be missed between fixed-attitude testpoints? The HARV inlet research program was designedto provide the information that would help answer thesequestions by obtaining full-scale flight data at steadyaerodynamic conditions and during maneuvering flightat comparable conditions.

Inlet Distortion at Steady Aerodynamic Conditions

The inlet data obtained at steady aerodynamicconditions provided the foundation for comparing andanalyzing inlet characteristic data during rapid AOAmaneuvers, aircraft departures, and CFD predictions.The objective was to document inlet characteristics atMach 0.3 and Mach 0.4 and at various stabilized AOAand AOSS conditions.

Inlet Distortion During Rapid AOA Maneuvers

Rapid AOA maneuvers were conducted to determinewhether results obtained for steady aerodynamicconditions were adequate for describing the inlet-generated, total-pressure distortion levels which occurduring such maneuvers. The evaluation focused onwhether the constrained steady aerodynamic conditiontest matrix describes inlet trends in sufficient detail ascurrently practiced. If the results of the rapid AOAmaneuvers at any condition showed a significantincrease in peak level distortion when compared withthe steady aerodynamic condition results, the inlet datawould be analyzed to determine the source of theseincreased peak distortion levels. The affects of rapidAOA maneuvers on inlet distortion levels can best beassessed during flight tests.

Inlet Distortion During Aircraft Departures

An aircraft departure, or departed flight, occurs whenan aircraft is flown in regimes where the control surfacesno longer have sufficient effectiveness to fully controlthe flightpath of the aircraft. The aircraft enters a seriesof self-induced motions that represent a departure fromnormal aircraft trajectories. In general, this type of flightis only encountered by military pilots who push theiraircraft to extreme AOA and AOSS. However, inletcompatibility engineers have long been interested inknowing the manner in which the aircraft inlet flowbehaves during departures. If engine stalls should occurduring a departure, engineers want to know theparameters that contributed to the stall.

For these reasons, a series of aircraft departures wereconducted to try to induce engine stalls. The objectiveswere to quantify the performance of the inlet during adeparture in terms of inlet recovery and dynamicdistortion characteristics and to gain insight into thecauses of engine stalls.

Dynamic Distortion Prediction Using a Combined CFD and Distortion Synthesis Approach

Computational fluid dynamics has been undercontinuous and wide-spread development since the late


1960’s. The CDF offers distinct advantages that can beperformed at any stage of the airframe developmentprocess. Methods were developed during the 1970’s and1980’s to improve the ability to predict inletperformance. These methods, known as distortionsynthesis, use a random number process to synthesizethe fluctuating component of the instantaneous totalpressure from the statistical properties of the inletpressure data.8

Turbulence modeling remains one of the mostsignificant challenges in CFD. Notable advancementshave been made over the past several years, includingmany formulations of two-equation models. Asufficiently accurate CFD solution, using a two-equationturbulence model, can be used to obtain the requiredinputs for a distortion synthesis procedure. Thisprocedure would permit predicting peak dynamicdistortion before costly wind-tunnel or flight tests.

A procedure was developed and evaluated by TheBoeing Company (formerly McDonnell DouglasAircraft (MDA) Corporation, St. Louis, Missouri) usingflight data obtained from the HARV. This procedureinvolved a combined approach of (1) obtaining a CFDsteady-state prediction of the inlet diffuser flowfield, (2)obtaining the correlation of the computed data with root-mean-squared (RMS) turbulence, and (3) generatingestimates of the peak dynamic distortion using asynthesis approach. Reference 6 describes the procedurein detail.

Estimating Engine Airflow with Clean and Distorted Airflow

Engine airflow is an important parameter whenaddressing turbofan engine-powered aircraft issues. Forexample, inlet distortion is often correlated with engineairflow corrected to the aerodynamic interface plane(AIP). The AIP engine airflow is also an important partof the net thrust calculation.

The HARV provided a unique opportunity to addressthe difficult issue of providing reliable estimates ofin-flight-determined, total-engine airflow. The objectivewas to compare three methods of correlating andestimating airflow with one method having threevariants. Reference 7 describes the methods forobtaining in-flight airflow. Ground-test-cell flowcalibration of the engine and establishment of flowcorrelations for each method in flight at minimum inletflow distortion conditions were crucial to the success ofevaluating the airflow methods.

Airplane Description

Figure 1 shows the HARV, a single-place F/A-18Aaircraft (pre-production aircraft number 6) built byMDA and the Northrop-Grumman Corporation(Newbury Park, California). The F/A-18A HARV ispowered by two modified F404-GE-400 turbofanengines (General Electric Aircraft Engines (GEAE),Evendale, Ohio). Wing leading-edge extensions (LEX)are located on each side of the fuselage from the wingroots to just forward of the windshield. This aircraft wasmodified with extensive instrumentation and multiaxisthrust-vectoring paddles. Thrust vectoring provided theHARV with the ability to fly at sustained, aerodynamicattitudes beyond the capabilities of conventional high-performance aircraft.

EC91 495-15

Figure 1. NASA F/A-18A HARV aircraft(preproduction aircraft number 6) with multiaxis thrust-vectoring paddles.

Propulsion System Description

The F/A-18A aircraft inlets are fixed-geometry,external compression inlets with 5° compression rampsmounted on the sides of the aircraft fuselage. Inlets arelocated approximately 25 ft aft of the aircraft nose underthe LEX of the wing (fig. 1). Figure 2 shows a schematicof the air induction system, including key inletdimensions. These inlets are located approximately 5 in.from the sides of the fuselage to avoid ingestion of thefuselage boundary layer. The subsonic diffuser makes agradual transition from the inlet throat to the engine facewhich is centered approximately 12 in., or 0.4 ductdiameter, above and 12 in. inboard of the inlet centroid.The diffuser is approximately 12 ft long with a resultinglength per diameter of the engine (L/DENG) of


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

Aircraft centerline

Engine gas and fuel system pressures

42 duct static pressures

40 high- and 40 low-response total pressures

Vortex generators

Porous bleed plate

Rake face Flow

FS379FS411FS550.9

Top view (not to scale)

Side view 980360

23.3

2°

5°

FS425(throat)

Engine face

FS500FS560

approximately 5.3. A pair of vortex generators is locatedon the lower surface of the diffuser to prevent local flowseparation in this area.

The external and internal geometry of the inlet cowllip was optimized for maneuvering in the subsonic,high-AOA region of the flight envelope. The lower andlower-inboard portions of the inlet lip were cut back andthickened to decrease the compressor face distortion atextreme aircraft attitudes. References 9, 10, and 11provide further details of the F/A-18A air inductionsystem.

Figure 3 shows the F404-GE-400 engine, a lowbypass, twin-spool, and axial-flow turbofan withafterburner. Two single-stage turbines independentlydrive the three-stage fan and seven-stage high-pressurecompressor. In the fan, the inlet-guide vanes (IGV) and

the stators of the first stage are variable. In the high-pressure compressor, the IGV and the first two statorstages are variable. These variable IGV direct the inletair at the optimum angle for efficient and stable engineoperation. The through-flow annular combustor usesatomizing fuel nozzles, and the mixed-flow afterburneruses air from the bypass and from the high-pressurecore. The F404-GE-400 engine control is an integratedsystem, using both hydromechanical and electroniccontrol components. The uninstalled sea-level-staticmilitary thrust of each engine is approximately10,700 lbf, and the maximum afterburner thrust isapproximately 16,000 lbf. Maximum corrected airflowthrough the engine is approximately 144 lbm/sec.

Three thrust-vectoring vanes were mounted on theaircraft and positioned approximately 120° apart aboutthe periphery of each engine behind the nozzle exits.


Figure 3. Right engine instrumentation supporting HARV inlet research program.

Engine inletFan inlet Compressor

inlet

Compressor discharge

1 2 2.5 3 4 5 558 6 7 8

Combustor discharge

Turbine discharge Turbine discharge

measuring plane

Afterburner inlet

Exhaust nozzle inlet

Exhaust nozzle throat

Turbine dischargepressure

Inlettemperature

Combustorfuel flow

Nozzlearea

980361

Fan rotor speed andvane guide position

Combustorfuel flow andtemperature

Combustorstatic pressures

Productionengine40 sps

Enginediagnostics

800 sps

Turbine exhaustgas temperature

Combustorstatic pressure

Additional engine (flight test)

40 sps

Enginestation

Afterburner fuel flows and temperature

Compressor rotor speed and vane guide position

Throttleangle

Fan dischargestatic pressures

Inlet and outlet main fuel control pressure

The corners of each vane were clipped to avoidinterference with adjacent vanes at full deflection of25°. These engines were modified to accommodate thethrust-vectoring vane installation by removing thedivergent section of the nozzle. Reference 1 providesfurther details of the HARV thrust-vectoring vanesystem.

Instrumentation

The HARV inlet research objectives required accuratemeasurement of specific inlet, engine, and airdataparameters. These measurements were recorded duringsteady aerodynamic flight conditions up to 60° AOA,rapid AOA maneuvers, and aircraft departures withengine stalls.

Aircraft and Engine

Aircraft instrumentation included accelerometers, rategyroscopes, surface position measurements, and airdata.A flight research airdata system consisting of swivelingpitot probes with conventional AOA and AOSS vaneswas mounted on each wingtip. These probes swiveled

freely for local AOA between –15° and 72° and for localAOSS between ±40°.12

Both engines had basic instrumentation formonitoring engine operation and were equipped with areal-time thrust measurement system.13 Figure 3 showsthe instrumentation that was added to the productionengine instrumentation which included guide vanepositions, fan and compressor speeds, fuel flows, anddiagnostic pressures.

Inlet Rake

An innovative inlet total-pressure distortionmeasurement rake was developed for the F/A-18A/B/C/D aircraft inlet by a team of NASA DFRC andGEAE personnel. The inlet rake was installed at the AIP,which is defined as the measurement plane between theinlet and engine. Inlet distortion and performance aredetermined in this measurement plane. For the F/A-18A,the AIP is located 4 in. in front of the bullet nose of theengine. The inlet rake consisted of a streamlinedcenterbody and eight aerodynamic rake legs. Each rakeleg consisted of five dual probes located at the centroids


of five equal areas. The 40 dual probes measured high-and low-frequency response total pressures. Figure 4shows the rake orientation and the nomenclatureassigned to each pressure port. References 2, 3 and 14describe the inlet rake. Orientation of the rake wassimilar to that used in previous F/A-18A inlet tests.9

The inlet rake probe was designed to be insensitive toairflow angularity, an important criterion whenmeasuring total pressure in distorted flows. Thisconfiguration allowed the sensors to read true pressurelevels of local airflow at yaw angles from ±25° and atpitch angles from 15° to –25° with positive angles beingin the direction of the engine centerline. Blockage oftotal airflow at the AIP caused by the inlet rake probeswas 0.4 percent, and the maximum blockage of the flowarea caused by the rake structure was less than8 percent. The rake structure was located 1.5 in.downstream of the AIP.

Inlet Low-Response Pressures

The low-frequency response static and total pressures(PSE and PTE) were measured using differentialpressure transducers. To obtain the absolute pressures,an accurately measured reference pressure was added tothe differential pressures. The right-hand aft lookingforward (ALF) engine bay was selected as the referencepressure location.

Each differential transducer unit provided themeasurement of 32 pressures. These units were

thermally stabilized to minimize zero drift associatedwith temperature variations and were capable of in-flight calibrations.

Inlet High-Response Pressures

The high-frequency response static and total pressures(PSK and PTK) were measured using individualtemperature-compensated pressure transducers. Eachtransducer was mounted at a rake measurement port inclose proximity to a low-frequency response pressureprobe. Calibration procedures for the high-frequencyresponse transducers allowed accurate measurement oftotal and static pressures during rapid aircraft maneuversand departures. The sample rate of these transducers wasconfigured to 2143 samples/sec.

Inlet wall static pressure measurements at the inletrake and around the inlet entrance were also obtainedduring the HARV inlet program. Results from thesemeasurements are not included in this paper.Reference 3 provides further details of these total andstatic pressure instrumentation and calibrationmeasurements.

Data Acquisition and Analysis

Data acquisition for the HARV inlet program usedthree pulse code modulation (PCM) systems. Two PCMsystems telemetered aircraft and engine data to ground-based computers. The third recorded inlet specific dataonboard the aircraft. To synchronize the three PCM


Figure 4. HARV inlet rake and instrumentation locations.

• Rake total pressures - PTK and PTE

980362

34°

Inboard

Rake 12

3

4

5

6

7

8

AB

CDE

Forward looking aft

ALF, not to scale

systems, an embedded time code was inserted into eachdata stream. These telemetered data were synchronizedwith the onboard recorded data to the fastest sample ratedata which was 2143 samples/sec. Data obtained foreach test condition were processed postflight, using dataquality and analysis software to assure high-qualityresults.

Inlet Parameters

The analysis calculations included inletcharacteristics, such as inlet recovery, circumferentialand radial total-pressure distortion levels as well asplanar wave and turbulence values. The softwareprovided time-averaged and peak-pattern screening ofthe inlet characteristics. The inlet parameters werecalculated based on industry standard guidelines15, 16

and established GEAE methodology.17, 18 Reference 3provides additional details of the data acquisition andanalysis methods.

Steady Aerodynamic Estimation Model

The inlet characteristics from the discrete steadyaerodynamic conditions were transferred to data look-up tables to allow comparison with the time-variantconditions of the rapid aircraft maneuvers. These datatables and other inlet calculations for each inletcomparison parameter were estimated as a function ofAOA, AOSS, and Mach number. A software applicationused these tables in combination with other inletcalculations to compute the comparison of the inletparameters between steady aerodynamic estimationsand the rapid aircraft maneuvers at equivalent AOA,AOSS, and Mach number conditions.

Trajectory Reconstruction

The AOA and AOSS vanes were calibrated to giveaccurate AOA and AOSS measurements during steadyaerodynamic maneuvers. The calibration removes suchaerodynamic affects on the indicated vane positions aslocal upwash, sidewash, and aircraft angular rates.These corrections are not valid during rapid aircraftmaneuvers. In addition, AOA and AOSS measurementhysteresis during aircraft maneuvers make thesemeasurements unreliable. An alternate source for thecalculation of AOA and AOSS during aircraftmaneuvers is the inertial navigation system (INS)onboard the aircraft. A procedure called trajectoryreconstruction provides more accurate AOA and AOSSresults than one using the data from the vanes. Duringaircraft maneuvers, INS, calibrated AOA and AOSSvanes, and winds aloft data are integrated. Trajectory

reconstruction required 2 sec at steady aerodynamicconditions before a rapid aircraft maneuver.

Flight Test Description

A matrix of flight maneuvers was flown to evaluatethe inlet research objectives. These maneuvers consistedof steady attitudes of AOA and AOSS, positive andnegative sweeps of AOA, and aircraft departures. Alltesting was performed between military (maximum non-afterburning) and maximum afterburning power settingsto maintain a constant corrected engine airflow ofapproximately 144 lb/sec. When at the required flightcondition, the power setting was held steady whileresearch data were recorded.

In-Flight Calibrations

During flight, calibrations of the inlet rake and duct,low- and high-response pressure transducers (PTE,PTK, PSE, and PSK) were performed just before aseries of maneuvers. The calibration required constantaltitude and airspeed for 15 sec. Reference 3 providesadditional details of the calibration process.

Steady Aerodynamic Conditions

Seventy-nine flight maneuvers at steady aerodynamicconditions were conducted at Mach 0.3 and Mach 0.4and from an altitude of 20,000 to 35,000 ft. Followingpressure transducer calibration data collection, thelength of recorded data for all maneuvers wasestablished at 6 sec. However because of aircraftmaneuver limitations, shorter data records wereobtained at the extreme AOA and AOSS conditions. Asuccessful maneuver required maintaining the flightcondition tolerances. Tolerance for Mach number duringall maneuvers was ±0.01. Tolerance for AOA and AOSSwas ±1.0°. For the ±10° AOSS maneuvers, it was notpossible for the airplane to maintain a 10° or –10° AOSScondition. As a result, the maximum sustainable AOSScondition was used. Reference 3 provides additionaldetails of the steady aerodynamic flight conditions.

Rapid AOA Maneuvers

Forty-six aircraft maneuvers with rapidly changingAOA were performed by holding the initial Machnumber, AOA, AOSS, and engine airflow for 2 to 3 sec.Table 1 shows the initial AOA and AOSS conditions.

After the steady conditions were held for 2 to 3 sec, amaximum rate AOA sweep was performed. Suchsweeps were performed at negative and positive ratesdependent on the initial AOA condition. The overall


maximum AOA rates achieved were –37 and 46 deg/sec.The end point of a maneuver was established when theMach number exceeded 0.02 of the initial Machnumber. Limiting the variation in Mach number ensuredvalid comparisons between the steady aerodynamicconditions and the rapid AOA maneuvers. The rapidAOA maneuver required a reversal in the AOA sweeponce the pilot exceeded a maximum AOA limit (60° atMach 0.3 and 40° at Mach 0.4) or minimum AOA limit(10° at both Mach 0.3 and Mach 0.4).

.

Aircraft Departures

Twelve high power departed flight maneuvers wereperformed between Mach 0.3 and Mach 0.4 at an entrycondition of 35 kft. These high yaw rate departures weredivided evenly between nose-left and nose-rightmaneuvers with increasing rates of change in yaw atentry. Throttles for both engines were set at militarypower at the start of the maneuvers and were reduced toidle power during the aircraft recovery phase.

Flight departure maneuvers were initiated by startingfrom a 1-g flight condition. The pilot increased AOA tobetween 50° and 60°. Then, either a left or right yawrate was induced by moving the control stick to aposition that produced the desired entry yaw rate. Afterthe desired rate was achieved, the pilot initiatedrecovery of the aircraft from the departed flightcondition by releasing the stick and allowing it to returnto the neutral position. At the same time, the pilotmoved the engine throttles to the idle power position.

Reference 5 provides further details of departure flightmaneuvers.

Engine Airflow Measurements

In-flight engine airflow measurements were obtainedin two phases. First, inlet total pressure airflow datawere obtained for 1-g level flight, at Mach 0.6, and at analtitude of 20,000 ft. These flight conditions correspondto minimum inlet airflow distortion levels. Second,airflow data were obtained during various steadyaerodynamic conditions to evaluate airflow during lowto high distortion levels and to determine what affectdistortion has on the various airflow correlationmethods.

Results and Discussion

This section summarizes the inlet rake data obtainedduring the HARV inlet program. Consistent results for11 test points over the course of the flight test programwere obtained at Mach 0.3, 30° AOA, and 0.0° AOSS.3

Inlet performance data obtained at steady aerodynamicconditions show the affects of AOA, AOSS, and Machnumber on inlet recovery, turbulence, and peak dynamicdistortion. During rapid, high-AOA maneuvers, inletdistortion levels are compared with levels atcorresponding steady aerodynamic flight conditions.Flight departure data are presented to show the cause ofthe engine stalls. A comparison of CFD inlet distortionprediction with flight data is then presented. Resultsfrom four airflow estimation methods are provided.

Inlet Performance at Mach 0.3

Figures 5 and 6 show the trends for inlet recovery,turbulence, and peak circumferential and radialdistortion values as a function of AOA and AOSS atMach 0.3. These data points represent time-averagedvalues. A solid symbol indicates that only one set oftime-averaged data was obtained for the target AOAcondition. The open symbols and the x indicate thatmultiple datasets were obtained at the target AOAcondition. Positive AOSS indicate that the aircraft waspointing nose-left, windward for the right-hand inlet.

Inlet Recovery

Figure 5(a) shows the affect of AOA and AOSS oninlet pressure recovery. An inlet recovery value of 1.0 isthe ideal value and indicates 100-percent pressurerecovery at the AIP when referenced to the free-streamtotal pressure. Negative AOSS have an obviousdetrimental affect on inlet recovery for all AOA,

Table 1. Initial aerodynamic conditions fordynamic maneuvers.

Mach 0.3 Mach 0.4

AOA, deg

AOSS, deg

AOA, deg

AOSS, deg

10 –5 10 –5

10 0 10 0

10 5 10 5

30 –5 13 –5

30 0 13 0

30 5 13 5

60 –2 40 –4

60 0 40 0-1

60 2-3 40 4


showing the impact of flow washing over the fuselage.As AOA increases, the detrimental affect of AOSSbecomes increasingly pronounced. Pressure recoveryremains constant at approximately 97 percent for 4° to30° AOA. As AOA increases above 30°, the loss inpressure recovery becomes increasingly sensitive to

AOA. At 60° AOA and 0° AOSS, the pressure recoveryhas dropped below 91 percent.

Turbulence

Figure 5(b) shows the affect of AOA and AOSS onturbulence levels at Mach 0.3. The detrimental affects of


(a) Inlet recovery.

(b) Inlet turbulence.

Figure 5. Affect of AOA and AOSS on inlet recovery and turbulence during steady aerodynamic conditions at Mach 0.3.

1.00

.98

.96

.94

.92

.90

.88– 15 – 10 – 5 0 5 10

10°

15

Inletrecovery

AOSS, deg

AOA

– 9°4°

10°20°30°40°50°57°60°

970459

60°

7% recoveryloss from 10°to 60° AOA

.028

.024

.016

.020

.012

.008

.004

0– 15 – 10 – 5 0 5 10

60°

10°

15

Turbulence

AOSS, deg

AOA

– 9°4°

10°20°30°40°50°57°60°

970461

AOA and AOSS on turbulence are similar to those shownfor inlet recovery. Comparing figure 5(a) with figure 5(b)clearly shows that as the flow turbulence increases, inletrecovery decreases. Going from nose-left to nose-right(positive to negative AOSS) increases turbulence. From4° to 20° AOA, the turbulence level is approximately0.005. The turbulence level begins to increase atapproximately 30° AOA. At 60° AOA, turbulence is upto 0.022. At Mach 0.4, the turbulence level at –10° AOAincreases almost to the level at 40° AOA.3

Peak Dynamic Circumferential Distortion

Figure 6(a) shows the affect of AOA and AOSS on thepeak dynamic circumferential distortion. Detrimentalaffects of AOA and AOSS on circumferential distortionare similar to those for inlet recovery. Comparingfigure 6(a) with figure 5(a) shows that circumferentialdistortion increases as inlet recovery decreases, ananticipated result. The affect of AOSS is pronounced forall AOA at Mach 0.3. As the aircraft moves increasinglynose-left (positive AOSS), the peak dynamic distortiondecreases. As the aircraft moves increasingly nose-right(negative AOSS), the peak dynamic distortion increases.From 4° to 30° AOA, the peak dynamic distortionincreases slightly. Above 30° AOA, the peak dynamicdistortion increases rapidly. This distortionapproximately doubles from 0.08 at 30° AOA to 0.17 at60° AOA.

Peak Dynamic Radial Distortion

Figure 6(b) shows AOA and AOSS having little affecton peak dynamic radial distortion up to 40° AOA.Increasing AOA from 40° to 60° causes the peak radialdistortion to increase rapidly. Increasing AOA to 50°and 60° increases the sensitivity to AOSS, similar to theAOSS affect on circumferential distortion. Reference 3provides additional details.

Inlet Distortion for Rapid AOA Maneuvers

To determine whether the results obtained duringsteady aerodynamic conditions were adequate fordescribing the inlet-generated, total-pressure distortionlevels which occur during rapid AOA maneuvers, aseries of positive and negative AOA sweeps wereperformed. Maximum rates of 46 deg/sec and

were attained. Inlet recovery and peakdistortion values were then compared with an estimatedvalue based on equivalent steady aerodynamicconditions.

Figure 7 shows inlet recovery and circumferentialdistortion levels during a low- to high-AOA sweep atMach 0.3 with AOSS near 0° compared with the steadyaerodynamic estimation. Figure 7(a) shows the AOAsweep from 10° to 62° with corresponding AOSS.Figure 7(b) shows good agreement between the

–37 deg/sec


(a) Maximum peak dynamic circumferential distortion.

Figure 6. Affect of AOA and AOSS on peak dynamic circumferential and radial distortion during steady aerodynamicconditions at Mach 0.3.

.20

.16

.12

.08

.04

0– 15 – 10 – 5 0 5 10

10°

60°

15

Peak dynamiccircumferential

distortion

AOSS, deg

AOA

– 9°4°

10°20°30°40°50°57°60°

970465

12

American Institute of Aeronautics and Astronautics

(b) Maximum peak dynamic radial distortion.

Figure 6. Concluded.

(a) AOA and AOSS, deg.

Figure 7. Time histories of inlet recovery and circumferential distortion compared with the steady estimation modelduring a low- to high-AOA sweep at Mach 0.3.

.20

.16

.12

.08

.04

0– 15 – 10 – 5 0 5 10

10°

60°

15

Peakdynamic

radialdistortion

AOSS, deg

AOA

– 9°4°

10°20°30°40°50°57°60°

970463

70

AOA andAOSS, deg

Time, sec10 2 3 4 5

980363

60

50

40

30

20

10

0

– 10

AOATJ

AOSSTJ

(b) Inlet recovery comparison.

(c) Circumferential distortion comparison.


.99DynamicSteady

Inletrecovery

0Time, sec

1 2 3 4 5

980364

.98

.97

.96

.95

.94

.93

.92

.91

.90

.89

0

.20

Time, sec

Circumferentialdistortion

1 2 3 4 5

980365

.16

.12

.08

.04

recovery level for the rapid AOA maneuver and thesteady estimation model. A discontinuity in recovery atapproximately 55° occurred. This discontinuity isbelieved to result from increased inlet flow separation.Figure 7(c) shows a comparison of the circumferentialdistortion levels. As expected, the level of distortionincreases as high AOA is reached. The peaks in thecircumferential distortion levels for the transientportions of the rapid AOA maneuver are less than thelevels of the steady aerodynamic estimation model,which represents the maximum peaks at equivalentAOA. Reduced distortion levels were obtained for allAOA transients between 30° and 62°. The low distortionlevels during rapid AOA transients confirm the currentpractice of fixed-attitude testing for obtaining peakdistortion values. However, flight conditions wereencountered where unexpected inlet recovery variationswere obtained.

Figure 8 shows inlet recovery and circumferentialdistortion levels during a mid- to low-AOA sweep atMach 0.3 compared with levels obtained from thesteady aerodynamic estimation model of the maximumdistortion levels. Figure 8(a) shows the AOA sweepfrom 28° to 0° and AOSS from –6° to 6°. Figure 8(b)shows good comparison for inlet recovery between theAOA sweep and the steady estimation model.Figure 8(c) shows a comparison of the circumferentialdistortion levels. Except between 0° and 2° AOA, thepeaks in the circumferential distortion levels for thisAOA maneuver were less than the levels obtained fromthe steady aerodynamic estimation model.

An examination of the other rapid AOA maneuversverified that at a given flight condition, inlet distortionlevels are less than the levels at an equivalent steadyaerodynamic condition. The sudden change in the


14



(b) Inlet recovery comparison.

(c) Circumferential distortion comparison.

Figure 8. Time histories of inlet recovery and circumferential distortion compared with the steady estimation modelduring a mid- to low-AOA sweep at Mach 0.3.

AOATJ

40

30

20

10

0

– 10

AOA andAOSS, deg

0 2 4 6Time, sec

980366

AOSSTJ

DynamicSteady

Time, sec20 4 6

980367

.986

Inletrecovery

.982

.978

.974

.970

.966

.962

.958

LEX-generated flow disturbance

0

.20

Time, sec

Circumferentialdistortion

2 4 6

980368

.16

.12

.08

.04

distortion level near 0° AOA and as the aircraft movedfrom nose-right to nose-left is believed to result from anLEX-generated flow disturbance. Note also that the inletrecovery drops slightly near 0° AOA. The discretesteady aerodynamic conditions test matrix clearly didnot provide a sufficiently detailed description of the inletbehavior during this maneuver. No steady aerodynamicdata were obtained at 0° AOA. This result was observedin a number of low-AOA sweeps at Mach 0.3 and Mach0.4. The maximum difference in recovery seen for theentire database was 0.005.

Inlet Distortion During Aircraft Departures

Twelve aircraft departures were achieved betweenMach 0.3 and Mach 0.4 at 35 kft by pulling AOA toapproximately 60° and inducing left or right yaw ratesof 40 to 91 deg/sec. Engine stalls were obtained during4 of the 12 departures. Figures 9–13 show data from aright-hand departure performed to a maximum yaw rateof 91 deg/sec. This departure induced two right-handengine stalls. These stalls occur at approximately 10.8and 14.6 sec and are indicated on these figures by eventmarkers.

Figure 9 shows time histories of the aircraftaerodynamic positions (AOA and AOSS) and motionsduring a departure. The AOA and AOSS were obtainedfrom a NASA trajectory reconstruction analysisprocedure to correct for winds aloft. Figure 9(a) showsthat AOA ranged from approximately 20° to 90°, andAOSS ranged from approximately –30° to 15°.

Figure 9(b) shows the aircraft motion as described byrate of change of pitch, roll, and heading. These datawere obtained from the INS and onboard rategyroscopic measurements. For this departure, aircraftrates ranged from approximately –30 to 45 deg/sec forpitch rate, –5 to 70 deg/sec for roll rate, and 0 to90 deg/sec for yaw rate.

Examining the pressure waveforms obtained fromhigh-response instrumentation at the fan discharge andcompressor discharge permits identification of the stall-initiating engine component. Figure 10 shows timehistories of the discharge pressures and the engine inlettotal pressure. The relative phasing and the direction ofthe pressure pertubations gives the sequence andpropagation of the instability. An increase in pressurefrom the prestall level indicates that the pressure waveassociated with stall initiation is propagating forwardfrom high pressure located downstream of themeasurement location. Analysis of the stall events whichoccurred during departed flight revealed that these stallswere initiated in the compressor. This result wasunexpected because, based on F404-GE-400 distortionmethodology, engine stalls caused by inlet maneuverdistortion affects should be initiated in the fan. All ofthese engine stalls recovered without pilot action.

Figure 11 shows time histories indicating thevariations of the inlet recovery and distortionparameters. Predicted loss of stability pressure ratios(DPRS) is also shown. The top of figure 11(a) shows



Figure 9. Time histories of aircraft aerodynamic positions and rates during a departure.

0 5 10Time, sec

980369

15

AOA andAOSS, deg

100

80

60

40

20

0

– 20

– 40

Stall 1 Stall 2

Departure

AOATJ

AOSSTJ

16


(b) Aircraft rates, deg/sec.


Figure 10. Time histories of measured inlet–engine entry and engine internal pressures.

60

40

20

0

– 20

– 40

60

80

40

20

0

– 20100

80

60

40

20

0 5 10Time, sec

980370

15

Pitchrate,deg/sec

Rollrate,deg/sec

Yawrate,deg/sec

Stall 1 Stall 2

10

Engineinlet,psia

Fandischarge,

psia

Compressordischarge,

psia

5

020

10

0100

50

0 2 4 6 8Time, sec

10 12

980371

14 16

Stall 1 Stall 2

(a) Inlet recovery and distortion descriptors.

Figure 11. Time histories of inlet recovery, distortion descriptors, and predicted loss of stability pressure ratio for thefan and the compressor.

1.1

Inlet recovery

DP/PR

DP/PC

Stall 1 Stall 2

1.0

.9

.8.25

.20

.15

.10

.05

0.15

.10

.05

0 2 4 6 8Time, sec

10 12

980372

14 16

that the increase in AOA coupled with the decrease inAOSS caused a reduction in recovery and an increase inlarge, dynamic variations in the distortion descriptors.An examination of the departed flight maneuversrevealed that the magnitude of the peak time-variant,total-pressure distortion levels were well beyond thoseencountered in the normal engine operating envelope.The peak time-variant circumferential total-pressuredistortion levels exceeded the limits used as guidancefor the F404-GE-400 engine installation by as much as25 percent.

Figure 11(b) shows the predicted losses of stabilitypressure ratio of the fan and the compressor resultingfrom the measured levels of time-varying spatial inlettotal-pressure distortion. High levels of DPRS for thecompressor (DPRSH) occur immediately before the

stalls. However, there were many instances whereequally high DPRSH levels occurred which did notinitiate a stall. The occurrence of the stall events did notcorrelate with the magnitude of the distortion or theDPRSH levels alone. Other factors were involved.

Reviewing the stall data revealed that a parameterdescribing the combined rate of change of aircraft pitch,roll, and yaw would be an improved correlating factor ofthe stall events. Rates of change of aircraft motionproduce gyroscopic moments. These moments andforces change the clearances of the rotor airfoils andthereby affect the stability limits of the compressioncomponents. For these reasons, a combined rate ofchange of aircraft motion parameter (CROCP) wasdefined by taking the root-sum-square of the pitch, roll,and yaw rates.


(b) Predicted loss of stability pressure ratio for the fan and compressor.


DPRSF

DPRSH

Stall 1 Stall 2.15

.15

.10

.05

0 2 4 6 8Time, sec

10 12

980373

14 16

.10

.05

0

Figure 12 shows the stalls superimposed on the timehistory of the combined rate of change of aircraftmotion. The stalls occur at local maximum values. Otherdeparture-induced stalls exhibited the same behavior.However, maximum values of combined rate of aircraftmotion alone were not an indicator of a stall.

By combining the affects of inlet distortion andaircraft motion, figure 13 shows the loss of compressorstability ratio as a function of the combined rate ofchange values throughout all of the departures at stalland at poststall. These stall data fall into two classes:isolated and non-isolated. The isolated data arecharacterized by no stall occurring during the previous2.4 sec or more. The non-isolated stalls may haveoccurred because the engine did not have time tore-establish the thermal equilibrium characteristics thatare associated with normal operation.

Of the 14 right-hand engine stalls, 7 were isolatedstalls and 5 were non-isolated stalls. Data were not

available for the two remaining stalls. These isolatedstall data all tend to group at simultaneous high valuesof predicted DPRSH and CROCP as expected;meanwhile, the non-isolated stalls occurred at lowcombined values. Therefore, the isolated stalls whichoccurred during departed flight appear attributable to thecombined affect of inlet total-pressure distortion on thecompressor and compressor clearance changes. Changesin compressor clearance may result from eccentricity ofthe compressor rotor, distortion of the compressorcasing caused by high rates of aircraft motion, or both.

CFD Inlet Distortion Prediction Comparison with Flight Data

A procedure was developed to estimate inlet dynamicdistortion using CFD-computed data. Figure 14 showsthat the prediction of the average total-pressure recoveryat the AIP at low AOA were within 1 percent of the flightdata. This result was typical for eight simulated flightconditions. At high AOA, the predicted recovery patterns


19


Figure 12. Time history of the combined rate of change of aircraft motion.

Figure 13. Affects of aircraft motion on time-variant inlet distortion on compressor stall line.

Combinedrate-of-change

of aircraft motion,deg/sec

150

100

50

0 4 8 12Time, sec

980374

16 20

Stall 1

Stall 2

.16

.12

.04

0 20

.08

Effect of inletdistortion oncompressorstall line –

DPRSH

40 60Combined rate-of-change of aircraft motion, deg/sec

80 100 120

980375

Stall-free data"Isolated" stall data"Non-isolated" stall data

Figure 14. Steady-state total-pressure recovery comparisons at 3.6° AOA, Mach 0.4, and ALF.

Flight measured steady state pattern CFD predicted steady state pattern

980376

0.9740.0410.0590.116

– .16– .14– .12– .10– .08– .06– .04– .02 0

RecoveryDP/PCDP/PRD2

0.9780.0270.0500.097


– 2

2

– 4

– 4

– 6

0

were significantly different (fig. 15). The CFD-predictedminimum total pressures at the AIP were consistentlyless than the flight test data, which caused thepredictions of dynamic distortion to be high in all cases.

Several suspected causes for the differences betweenpredicted and measured engine face recovery patternsexist. An accurate computation of the vortices generatedfrom the vortex generators is essential in computingaccurate recovery patterns, in computing turbulencelevels, and in synthesizing dynamic distortion. Smallerrors in the trajectories of these vortices can cause alarge error in the interpolated value of a total pressureand erroneously dominate the synthesized patterns. Thetrajectory error is thought to be a strong function of theturbulence model. Another possible cause is modelingthe engine effect. A simulation of a porous screenrepresenting the engine face helped the predictions.Another cause for the CFD model deficiency is thatmass flow is currently corrected using full CFD exitrecovery, not a 40-probe average at the AIP. Thisdifference in recovery can be as large as 0.8 percent. Inaddition, the inlet flow is assumed to be completelyturbulent. An inlet boundary-layer transition region mayexist which significantly affects the subsequentboundary-layer separation and vortex formation.

Engine Airflow Estimation

Impact of inlet-generated, total-pressure distortion onestimated levels of engine airflow was studied. Four

airflow estimation methods were used. The referencemethod was a fan-corrected airflow to fan-correctedspeed calibration from an uninstalled engine test.In-flight airflow estimation methods used the average, orindividual, inlet duct static- to total-pressure ratios andthe average fan-discharge static pressure to average inlettotal-pressure ratio. Correlations were established at lowdistortion conditions for each method relative to thereference method. A range of distorted inlet flowconditions was obtained from –10° to 60° AOA and –7°to 11° AOSS.

Figure 16 shows a comparison of two of the fourairflow estimation methods as a function of thereference method. The solid symbols represent datapoints where steady-state AIP total pressure patternswere generated and compared (fig. 30, ref 7).Figure 16(a) shows that the individual inlet probepressure ratio method correlation resulted in a±1.15-percent airflow spread for all distorted flow levelswith a bias error of –0.7 percent. Figure 16(b) showsthat the fan-discharge pressure-ratio method correlationresulted in a ±0.3-percent airflow spread withessentially no bias error. Inlet-generated total-pressuredistortion and turbulence had no significant impact onengine airflow. As a result, a speed-flow relationshipmay provide the optimum airflow estimate for a specificengine under all flight conditions.


21


Figure 15. Steady-state total-pressure recovery comparisons at 39° AOA, Mach 0.4, and ALF.

(a) Method 1.

Figure 16. Corrected airflow relative to reference method as a function of circumferential distortion level (adaptedfrom ref. 7).

Flight measured steady state pattern CFD predicted steady state pattern

980377

0.9610.1050.0440.167

– .16– .14– .12– .10– .08– .06– .04– .02 0

0.9620.0730.0420.130



– 2

2

– 4

– 4

– 8

– 6

0

1

Correctedairflow relative to reference method,

percent

0

– 1

– 2

– 30

4

13

7

2

5

6

Mach0.30.4

8

9

10

111213

.02 .04 .06Maximum DP/PC

.08 .10 .12

980378

(b) Method 2.


2

1

Correctedairflow relative to reference method,

percent

0

– 1

– 20

4 1

3

7 2

56

Mach0.30.4

8910

111213

.02 .04 .06Maximum DP/PC

.08 .10 .12

980379

Conclusions

An extensive inlet aerodynamic database wasobtained for the NASA F/A-18A High Alpha ResearchVehicle (HARV). High quality HARV inlet data allowedsignificant insights and increases in the understandingof inlet performance at high AOA and during rapidaircraft maneuvers. Inlet data obtained at Mach 0.3, atan angle of attack (AOA) of 30°, and at an angle ofsideslip (AOSS) of 0° showed excellent repeatability.During steady aerodynamic conditions, inlet recoveryand peak circumferential distortion increased as AOAincreased and as AOSS became increasingly negative(nose-right). During rapid AOA maneuvers, it wasverified that at a given flight condition, the maximuminlet distortion levels are less than at an equivalentsteady aerodynamic condition. However, rapidmaneuvers did identify aerodynamic conditions whendiscrete changes in inlet behavior occurred.

Twelve intentional aircraft departures were conductedto obtain inlet data leading to and during engine stalls.Results revealed that the magnitudes of the peak time-variant total-pressure distortion levels were well beyondthose encountered in the normal engine-operatingenvelope. An unexpected result was that all engine stallswere initiated in the compressor, not in the fan aspredicted.

The computational fluid dynamics predicted inletdistortion comparison with flight data shows that the

prediction of the average total-pressure recovery at thelow-AOA conditions evaluated was within 1 percent ofthe flight test data. At high AOA, the predicted recoverypatterns were significantly different. Based on the studyof four airflow estimation methods, the speed-flowrelationship provides the best airflow estimate for aspecific engine under all flight conditions.

References

1Regenie, Victoria, Donald Gatlin, Robert Kempel,and Neil Matheny, The F-18 High Alpha ResearchVehicle: A High-Angle-of-Attack Testbed Aircraft,NASA TM-104253, 1992.

2Yuhas, Andrew J., Ronald J. Ray, Richard R. Burley,William G. Steenken, Leon Lechtenberg, and DonThornton, Design and Development of an F/A-18 InletDistortion Rake: A Cost and Time Saving Solution,NASA TM-4722, 1995.

3Walsh, Kevin R., Andrew J. Yuhas, John G.Williams, and William G. Steenken, Inlet Distortion foran F/A-18A Aircraft During Steady AerodynamicConditions up to 60° Angle of Attack, NASATM-104329, 1997.

4Yuhas, Andrew J., William G. Steenken, and John G.Williams, F/A-18A Inlet Flow Characteristics DuringManeuvers with Rapidly Changing Angle of Attack,NASA TM-104327, 1997.


5Steenken, William G., John G. Williams, andAndrew J. Yuhas, An Inlet Distortion Assessment DuringAircraft Departures at High Angle of Attack for anF/A-18A Aircraft, NASA TM-104328, 1997.

6Norby, W. P. and J. A. Ladd, A. J. Yuhas, DynamicInlet Distortion Prediction with a CombinedComputational Fluid Dynamics and DistortionSynthesis Approach, NASA CR-198053, 1996.

7Williams, J. G., W. G. Steenken, and A. J. Yuhas,Estimating Engine Airflow in Gas-Turbine PoweredAircraft with Clean and Distorted Flows, NASACR-198052, 1996.

8Sedlock, D., “Improved Statistical Analysis Methodfor Prediction of Maximum Inlet Distortion,”AIAA-84-1274, June 1984.

9Amin, N. F., C. J. Richards, E. G. de la Vega, andM. A. Dhanidina, F/A-18A Engine Inlet Survey Report,vol. 1 of 3, NOR 81-316, Northrop Corporation, AircraftDivision, Hawthorne, California, Northrop FSCM76823, November 1981.

10Morse, D. B., N. F Amin, F. W. Marxen, J. A.McGuire, E. G. de la Vega, and M. Yamadar, PropulsionSystem Functional and Performance Analysis Report,NOR 77-364, Northrop Corporation, Aircraft Division,Hawthorne, California, July 1978.

11Amin, N. F. and D. J. Hollweger, “F/A-18AInlet/Engine Compatibility Flight Test Results,”AIAA-81-1393, July 1981.

12Moes, Timothy R. and Stephen A. Whitmore, APreliminary Look at Techniques Used to Obtain AirdataFrom Flight at High Angles of Attack, NASATM-101729, 1990.

13Ray, R. J., J. W. Hicks, and R. I. Alexander,Development of a Real-Time Aeroperformance AnalysisTechnique for the X-29A Advanced TechnologyDemonstrator, NASA TM-100432, 1988.

14Thornton, D. A. and W. G. Steenken, Summary ofInlet Distortion Rake Activities Flight ClearanceThrough High-Temperature Rake Delivery, GeneralElectric Aircraft Engines, Cincinnati, Ohio, October1995.

15Society of Automotive Engineers, Inc., GasTurbine Engine Inlet Flow Distortion Guidelines,Aerospace Recommended Practice 1420, Warrendale,Pennsylvania, March 1978.

16Society of Automotive Engineers. Inc., Inlet Total-Pressure-Distortion Considerations for Gas-TurbineEngines, Aerospace Information Report 1419,Warrendale, Pennsylvania, May 1983.

17P. M. Shumate, GE Distortion Analysis ProgramUsers’ Manual, version 3, GEAE TM-88-352, GeneralElectric Aircraft Engines, Cincinnati, Ohio, September1988.

18General Electric Aircraft Engines, ModelSpecification for F404-GE-400 Turbofan Engine,Specification no. CP45K0006, Cincinnati, Ohio,November 1975, Reprinted February 1983.


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