NASA Technical Memorandum 104247

Effects of Bleed Air Extraction on Levels of the F404-GE-400 Engine

Andrew J. Yuhas PRC Inc., Edwards, California

Ronald J. Ray NASA Dryden Flight Research Facility, Edwards, California

1992

N/kS/_ National Aeronautics and Space Administration Dryden Flight Research Facility Edwards, California 93523-0273

EFFECTS OF BLEED AIR EXTRACTION ON THRUST LEVELS OF THE F404-GE-400 TURBOFAN ENGINE

Andrew J. Yuhas* PRC Inc. P.O. Box 273 Edwards, California 93523-0273

Ronald J. Ray** NASA Dryden Flight Research Facility P.O. Box 273 Edwards, California93523-0273

Abstract EAFB Edwards Air Force Base, Edwards, CA EBAS A ground test was performed to determine the effects excess bleed air system of compressor bleed flow extraction on the perform- ECS environmental control system ance of F404-GE-400 afterburning turbofan engines. FFC forebody flow control The two engines were installed in the F/A-18 High GE General Electric Alpha Research Vehicle at the NASA Dryden Flight Research Facility. A specialized bleed ducting system HATP High Alpha Technology Program was installed onto the to control and measure HARV High Alpha Research Vehicle engine bleed airflow while the aircraft was tied down HPC to a thrust measuring stand. The test was conducted high-pressure compressor on each engine and at various power settings. The HPT high-pressure turbine bleed air extraction levels analyzed included flow rates LPT low-pressure turbine above the manufacturer's maximum specification limit. LPC The measured relationship between thrust and bleed low-pressure compressor flow extraction was shown to be essentially linear at M Mach number all power settings with an increase in bleed flow caus- N1 fan rotor speed, rpm or percent ing a corresponding decrease in thrust. A comparison N2 with the F404-GE-400 steady-state engine simulation high compressor rotor speed, rpm or showed the estimation to be within ±1 percent of mea- percent sured thrust losses for large increases in bleed flow rate. PAMB ambient pressure, psia PLA power lever angle, deg Nomenclature PS3 compressor discharge static pressure, psia PT56 turbine exhaust total pressure, psia ALT pressure altitude, ft TVCS thrust vectoring control system AOA angle-of-attack, deg T1 fan inlet temperature, °R A8 nozzle throat area, in 2 or percent T25 compressor inlet temperature, °R BV1 excess bleed air system control valve 1 T56 low rotor turbine exit temperature, °R BV2 excess bleed air system control valve 2 or °C) DTAMB ambient temperature increment from WFE fuelflow,(Ib/hr) standard day temperature, °R

*Senior Propulsion Research Engineer. Introduction ** engineer. Copyright (_)1992 by the American Institute of Aeronau- The High Alpha Technology Program (HATP) was tics and Astronautics, Inc. No copyright is asserted in the initiated within NASA to develop and demonstrate United States under Title 17, U.S. Code. The U.S. Govern- ment has a royalty-free license to exercise all rights under concepts that provide favorable control and maneu- the copyright claimed herein for Governmental purposes. All verability for high-performance aircraft at high-angle other rights are reserved by the copyright owner. of attack. 1 Current fighters have degraded maneuver- Because the pneumatic FFC studies for the F/A-18 ability at high-angle of attack because of loss in aero- were encouraging, the HATP initiated a feasibility dynamic control surface effectiveness. One of the tech- study for a full-scale flight-test program to be con- nologies studied to enhance aerodynamic control is the ducted. To support this goal, the NASA Ames Re- use of pneumatic forebody flow control (FFC). The search Center at Moffett Field, CA, formed a research FFC injects air from a slot on the aircraft forebody team from Dryden and Moffett to evaluate engine to manipulate the forebody vortices that form at high- bleed air capability and aircraft environmental bleed angles of attack. For pneumatic FFC to be feasible, a air requirements on a full-scale aircraft. The F/A-18 significant supply of high-pressure airflow is required. High Alpha Research Vehicle (HARV) was equipped with a specially designed ducting system known as Experiments have demonstrated that the manipula- the excess bleed air system (EBAS) and tested on the tion of the asymmetric forebody vortices of a fighter at universal horizontal test stand at Edwards Air Force high-angle of attack through various blowing schemes Base (EAFB). This ducting system branches off from can have a strong effect on the yawing moment? -4 the aircraft environmental control system (ECS) and Forebody flow control feasibility studies into various is- allows engine bleed extraction to be controlled and sues have been conducted by various government agen- measured. Various ground test objectives were per- cies including NASA. The X-29 Advanced Technology formed using EBAS, including the determination of Demonstrator will test a simple pneumatic blowing sys- bleed air extraction effects on engine performance and tem to demonstrate FFC effectiveness. 5 The X-29 sys- operability. tem will use a compressed air bottle to supply the high- pressure air and is limited to a total of 20 sec of op- This paper addresses the measurement of thrust loss eration. The HATP supported research on pneumatic during increases in bleed air extraction for an F404-GE- FFC applications to an F/A-18, including various wind 400 engine installed on the F/A-18 HARV. A compar- tunnel tests. One feasibility study looked at various ison of the predicted thrust loss provided by an engine ways to obtain the estimated flow rates required. These computer simulation and the measured values will be included compressed or liquefied gas bottles, an aux- presented. The analysis offers a comparison of each en- iliary compressor, or engine bleed air. The results of gine at various power settings. A detailed description this study predicted that only engine bleed air would of EBAS and its function will also be presented. meet the mass flow requirements over an extended time period, e Bleed extraction however, has a negative ef- Description of Equipment and Model fect on engine performance. A past study on an F100- F-18 High Alpha Research Vehicle PW-220 engine predicts up to 2 percent of net thrust loss for each percent of airflow extracted for bleed. 7 The HARV (Fig. 1) is a one-of-a-kind research air- craft located at the NASA Dryden Flight Research

EC 91-201-9 Fig. 1. F/A-18 High Alpha Research Vehicle (HARV).

ORIGIr,,IAL P_E BLACK AND WHITE PHOiOGRAI..'r_ Facility.This pre-production,single-seatF/A-18 air- modifiedto provide a selectableturbine dischargetem- crafthas been uniquely modified to perform extensive perature biasforadditionalengine stallmargin at high flighttestingin the high-angle-of-attack(AOA) region. AOA. These modificationsinclude a thrust vectoringcontrol system (TVCS). The incorporationof the TVCS re- The controlsystem has two basicmodes of operation which aredependent on the throttlesetting(PLA). Be- quired modificationsto the aircraftavionics,flightcon- trols,hydraulics,cockpit,and engines.I low Militarypower, high compressor rotorspeed (N2) and exhaust nozzle area (A8) are scheduled. Throt- F404-GE-400 Turbofan Engines tle positionand compressor inlettemperature (T25) Two pre-production F404-GE-400 engines (General schedule N2. The throttlepositionschedules A8. At Electric,Lynn, Massachusetts) are installedin the Militaryand above, the controlsystem schedulesmax- HARV. The F404-GE-400 (Fig.2) is a two-spool,low imum fan speed (N1) and maximum low-pressuretur- bypass axial-flowturbofan with .The en- bine exittemperature (T56). Fan speed is controlled gine consistsof a 3-stage fan, 7-stage high-pressure as a function of engine inlettemperature (TI). The variableexhaust nozzle is modulated to maintain the compressor, annular combustor, l-stagehigh-pressure scheduled T56 which is a function of T1 and ambient turbine,and 1-stagelow-pressureturbine.Variablege- pressure (PAMB). ometry isincorporated into the fan and high-pressure compressor along with a convergent-divergentnozzle. Aircraft-Engine Bleed Air System

The F404 engines on the HARV have the divergent Bleed aircomes from an annular manifold at the exit portion of the nozzle and externalflapsremoved to in- of the high-pressurecompressor of each engine (Fig.3). corporate the TVCS. The engine controlsystem was A bleedairpressureregulatorand shutoffvalvein each

EC91-201-9 Fig.2. F404-GE-400 turbofan engine. LPT discharge measuring plane Exhaust LPC HPC LPT Afterburner nozzle throat Inlet discharge discharge Inlet Exhaust Exhaust Free- Engine I HPC ICombustor nozzle nozzle stream Inlet discharge I 'ni°'I InletI I discharge I I 0 1 2 2.5 3 4 5 5586 7 8 9

manifold 920148

Fig. 3. F404 engine showing compressor bleed and measured engine parameters. engine compartment is controlled by the bleed air knob Excess Bleed Air System located in the cockpit. The cockpit control allows bleed The EBAS was a specially designed ducting system air to be selected from either or both engine sources, for measuring bleed air flow rates from the F/A-18 or to be turned off. The bleed air passes through a HAKV (Fig. 6). The EBAS hardware was designcd check valve which prevents reverse flow from one en- by NASA Ames Research Center at Moffett Field, CA gine to the other. The bleed air lines are then teed and is illustrated schematically in Fig. 7. The system together. The bleed air line is routed up the backbone consists of two principal sections, the ECS flow mea- of the aircraft over the fuel tanks and then down into surement section and the bleed airflow dump section. the ECS and heat exchanger ejectors (Fig. 4). Bleed The ECS flow measurement section required a section flow is regulated by the ECS. Maximum engine bleed is of ECS duct along the backbone of the aircraft to be re_ limited by the size of the bleed port on the engine but moved and replaced with modified ducting and a Ven- can exceed the manufacturer's maximum specification turi flowmeter (Fig. 4). This section of ECS ducting limit (Fig. 5). 1. Environmental control system (ECS) duct replaced for bleed airflow dump section

-2. (ECS) duct replaced for flow measurement section

920149

Fig. 4. Bleed air duct assemblies.

_,ALT < 30,000 ft 8 \ _-ALT • 50,000 ft 7 \ _ i_"_\ orM_I.0 Bleed 6 .._ \\_" ALT _<50,000 ft air flow, 15 \\_and M> 1.0 percent 5 of high 4 Z'ALT > 30,000 ft kk pressure 3 rGRND Idle l \ / FLT I"1- ,-MIN A/B Inlet air flow / _ u e MIL- _-MAX A/B-, compressor i f 10 30 50 70 90 110 130 PLA, deg r2mso

Fig. 5. Customer bleed air extraction limits. could alsobe replaced with a capped offduct section, Input into the steady-statemodel includes altitude, thereby allowing no airflowto the ECS from the en- Mach number (M), power levelangle (PLA), and am- gines.The bleed airflowdump sectionrequiredan ad- blenttemperature. The program allows for determin- ditionaiECS lineto be replacedwith a modified duct. ing the effectson engine operation of engine bleed air The duct teed the bleedflowout ofthe ECS line,away extractionfrom 0 to the manufacturer maximum spec- from the aircraft and was dumped to atmosphere. This ificationlimit(Fig.5). ducting had two butterfly valves installed, one (BV1), immediately after the teed duct (Fig. 7, item 1) and the Test Setup-Procedure other (BV2), before the end of the ducting. This duct- Thrust Stand Description ing was also equipped with a Venturi flowmeter. The entire system was instrumented with pressure and tem- The EAFB universalhorizontalthrust stand, oper- perature measurements along with the flowmeters. The ated by the Air Force, was used to measure installed ECS line modifications for EBAS were downstream of thrust. The thrust stand consists of four force mea- the bleed air pressure regulator and shutoff valves and suring tables arranged in a cross formation (Fig. 8). upstream of the aircraft ECS and ejectors. This layout accommodates a wide variety of aircraft. A pair of load cells is attached to each platform. The General Electric F404 Steady-State HARV was positioned on platform 3 and tied down Engine Model with chains. Special care was taken to ensure no sup- port equipment or EBAS structure was positioned on The manufacturer's specification model (GE Refer- the measuring platform. ence LP005S) is a full aerothermal steady-state perfor- mance program. This model was modified to reflect A control room bunker is located underground near the F/A-18 HARVconfiguration. The simulation pro- platform 1 and was used to monitor load data_ local vides values of a number of internal flow parameters weather information, and aircraft status during the for the engine as well as its overall performance. The test. Load and weather data are measured and stored model was derived from actual test data and repre- in computer memory and on hardcopy printouts in the sents the operation of an average F404-GE-400 engine. bunker. The thrust stand operator had radio contact

Fig. 6. Excess bleed air system ducting, bleed flow dump section.

6

OR!_ l,',t._k P,_ _ BLACK AND WHITE PHOTOGRAPH with the aircraft and ground support personnel as well Excess Bleed Air System Configuration as the NASA Dryden control room during the testing. The setup of EBAS for steady-state thrust measure- Weather data recorded at the thrust stand included ment required the following configuration modifications ambient pressure and temperature as well as wind from the prior hardware description: (1) The ECS flow speed and direction. Wind limits were maintained dur- measurement section was replaced with a duct section ing the test 'to prevent data biasing caused by wind containing a cap. An external ground supply provided forces, variable inlet recovery, and tailpipe gas recircu- aircraft ECS cooling flow. This insured that all engine lation: maximum headwinds, 15 kts.; maximum cross- bleed flow was sent down the bleed airflow dump sec- winds, 5 kts; and maximum tailwinds, 0 kts. tion. (2) The BV1 valve was full open, and (3) the

To atmosphere BV2

4 4

Venturl BV1 flowmeter ECS duct capped off, allowing no flow to ECS during test.

1. ECS duct replaced for EBAS bleed airflow dump section (see Fig. 4) 2. ECS duct replaced for flow measurement section or capped duct (see Fig. 4) 3. Engine bleed air pressure regulating and shutoff valves I OnHARVF/A-18 4. Engine bleed air check valves 5. Secondary engine bleed air pressure regulating and shutoff valve To ECS

92O152

Fig. 7. Excess bleed air system for F/A-18 HARV. Platform 3 15 ft 0 in. x 42 ft 6 In.

J EBAS

Platform 2 Platform 1 27 ft 0 In. x 15 ft 0 in. 27 ft 0 In. x 15 It 0 In. I/L-

1

V

Noah TO bunker "-

Legend

_] Load cell

B_mm_

Platform 4

15 ft 0 In.x 42 ft 6 in. G20153

Fig. 8. Thrcst stand p]atforms. BV2 valve was used to controlbleed flowthrough the combustor fuel flow (WFE), and T56 were used to con- measurement ducting. firm stabilization. At Military and above, A8 and T56 were used. The EBAS flow measurements were also The operationof controlvalvesand recordingofmea- monitored to confirm stabilization. After a stabilized surements on the EBAS was located in a data van lo- condition was reached, measurements were recorded in cated near the thrust stand. All EBAS parameters the EBAS data van and thrust stand. Aircraft and (Table 1) were recorded on a portable data acquisi- engine measurements were continually recorded in the tion system. The system enabled real-timeparameter NASA data facilities. After measurement recording readouts includingbleed mass flow rates. The EBAS was complete, the BV2 valve was opened in a 5- or 10°- increment. Engine and EBAS measurements were al- Table 1. Excessive bleed air system parameters. lowed to stabilize and data was again recorded. This incremental opening of the BV2 valve continued un- Name Description Units til an engine limitation was reached or bleed flow no P1 Pressure before branch psig longer increased. P2 Pressure on 3-in. Venturi psig P3 Pressure on 6-in. Venturi psig Results and Discussion P4 Pressure before BV2 psig P5 Pressure after BV2 psig A comparison of the estimated and measured effects of bleed flow extraction on thrust levels of the F404- DP2 Differential pressure on 3-in. Venturi in. of water GE-400 engine at ground level is presented. Pxedic- DP3 Differential pressure on tions using the manufacturer's engine model indicate 6-in. Venturi in. of water a significant increase in bleed flow extraction from the T1 Temperature before branch °F F404 compressor will cause a substantial decrease in T2 Temperature on 3-in. Venturi °F thrust. All power settings tested had predictions which T3 Temperature on 6-in. Venturi °F showed a thrust loss with increasing bleed flow demand. T4 Temperature after BV2 °F Differences between measured and predicted thrust BV1 Butterfly valve 1 position degree (0-90) values were caused by engine-to-engine variations, en- BV2 Butterfly valve 2 position degree (0-90) gine degradation, and aircraft installation variations. The manufacturer's model assumes a nominal engine. Predictions were made using the measured bleed flow, data van had radio contact with the aircraft and ground power settings, atmospheric pressure and temperature support personnel, thrust stand operator, and the conditions. Because the EBAS valves would not seal NASA control room during the testing. completely in the closed position, about 0.7 lb/sec of Instrumentation bleed flow was measured at the minimum setting. 92-percent N2 - Left and Right Engine Along with the thrust stand and EBAS measure- ments, the F/A-18 HARV aircraft and F404 engines The comparison of estimated and measured thrust had extensive instrumentation installed. The engine loss at 92-percent N2 showed good agreement. The left parameters measured are shown in Fig. 3. All en- and right engine comparisons are presented in Figs. 9 gine and aircraft measurements were telemetered to and 10. The thrust loss with bleed flow extraction at the NASA control room and recorded. The engine op- part power shows a gradual thrust decrease with in- eration during the test was monitored in the NASA creasing bleed flow rate. At part power, the relation- control room by way of CRT displays and stripchart ship between bleed flow and thrust is nearly linear. For recordings. analysis comparison a least-squares curve fit was per- formed on the data. The least-squares curve fit of the Test Procedure estimated and measured data is given in Table 2. The Testing was conducted on each F404 engine at three estimated loss of 89 to 97 lb of thrust per lb/sec of power settings: part power (92-percent N2), Military, bleed flow is less than the 100 to 125 lb/lb/sec value and maximum afterburning. For each engine-power measured. The simulation at part power underpredicts setting condition, the following procedure was used. the thrust loss due to bleed flow extraction.

At idle power, the cockpit bleed air switch was se- Military Power - Left and Right Engine lected to the required engine. The BV2 valve was The estimated and measured thrust loss at Military set to full closed (0 ° position). The required power power with bleed flow demand was compared for each level was selected and the engine was allowed to stabi- engine (Figs. 11 and 12). The thrust loss at Military lize. The control room monitored engine operation to power has an essentially linear relationship with bleed determine a stabilized condition. At part power, N2, Table 2. Least-squares curve fit of predicted and measured thrust loss due to bleed flow extraction. Thrust = bleed flow rate (slope) -4-Y-intercept Slope: Percentage Power Data lb thrust/ Y-intercept:* thrust loss setting Engine source lb/sec bleed lb thrust bleed = 1 lb/sec 92 percent N2 Left Predicted -97 5,875 1.65 Measured - 125 6,311 1.98 Right Predicted -89 5,503 1.62 Measured -100 6,079 1.64

Military Left Predicted -245 9,194 2.66 Measured -177 8,774 2.02 Right Predicted -246 9,086 2.7 Measured -189 8,690 2.17 Maximum Left Predicted -421 13,241 3.18 Measured -371 12,873 2.88 Right Predicted -428 12,983 3.29 Measured -344 12,033 2.85 *Ideal zero bleed thrust value

ALT - 2084 ft, M = 0.0, ALT = 2093 It, M = 0.0, 6400 DTAMB = 34 °R DTAMB = 11 °R [] 6200 9200 • Simulation prediction 9000 ._ Measured data 6O00 8800 _ Maximum Net 5800 thrust, Ib 5600 th_st, WOO 5400 8200 5200 : , Net 8600 5000 0 1 2 3 4 5 8000 I I _ i I 78000 1 2 3 4 5 Bleed flow rate, Ib/sec 920154 Bleed flow rate, Ib/sec 920156 Fig. 9. Thrust loss due to bleed air extraction 92- Fig. 11. Thrust loss due to bleed air extraction; Mili- percent N2, left engine. tary power, left engine.

ALT = 2143 ft, M = 0.0, ALT = 2093 It, M = 0.0, DTAMB = 45 °R DTAMB = 18 °R • Simulation prediction 9000 o Simulation prediction [] Measured data 6200 8800 . Measured data Maximum xlmum 6000 Net 8600 5800 thrust, 8400 Net Ib thrust, 5600 8200 Ib pec limit 5400 8000 I I 5200 7800 0 1 2 3 4 5 5000 I I I I n Bleed flow rate, Ib/sec 0 1 2 3 4 5 920_57

Bleed flow rate, Ib/sec 920155 Fig. 12. Thrust loss due to bleed air extraction; Mili- Fig. 10. Thrust loss due to bleed air extraction 92- tary power, right engine. percent N2, right engine.

10 flow demand. A least-squares curve fit of the estimated The effect of power setting on thrust loss due to bleed and measured data is given in Table 2. A slope com- extraction is summarized in Fig. 15. This comparison parison shows the predicted loss of 245 lb of thrust per of percentage thrust loss for 1 lb/sec of bleed extraction lb/sec of bleed flow is greater than the measured value shows a definite sensitivity to power setting. For a of 177 to 189 lb/lb/sec for both engines. The estimated given bleed flow value, the percent thrust loss increases thrust loss at Military power is overpredicted compared with increasing power settings. The model tends to to the measured results. slightly over predict this tendency when compared to the measured values. Left to right engine comparisons Maximum Power - Left and Right Engine were very good at Military power and above and only The maximum power comparison of the estimated off by 0.4 percent at 92-percent N2. The estimated and measured thrust loss on the left and right engine is and measured losses were within +0.5 percent of each shown in Figs. 13 and 14. The maximum power thrust other. Overall, the steady-state simulation deck was loss due to bleed flow shows the same linear relationship able to provide good estimates for thrust loss due to as Military power. A least-squares curve fit study of the bleed flow extraction. These results encourage the use estimated and measured data is presented in Table 2. of the manufacturer's model to predict thrust loss due A slope comparison shows the estimated thrust loss of to bleed extraction at other flight conditions. 421 to 428 lb of thrust per lb/sec of bleed flow is greater than the measured value of 344 to 371 lb/lb/sec. The maximum power thrust loss estimation over pre- Predicted left dicts thrust loss when compared to the measured data. I Measured left Predicted right I Measured right ALT = 2082 ft, M = 0.0, 3.5 DTAMB = 29 OR o Simulation prediction 3.0 " Measured data 13,000 2.5

12,500 Thrust 2.0 lossl Net 12,000 percent 1.5 thrust, 11,500 1.0 Ib _- Maximum 11,000 spec limit .5 10,500 i i I i i 0 10,000(_ 1 2 3 4 5 92-percent Military Maximum N2 Bleed flow rate, Ib/sec 920158 Power setting _,olso

Fig. 13. Thrust loss due to bleed air extraction; max- Fig. 15. The effect of power setting on the percent of imum power, left engine. thrust loss for 1 lb of bleed extraction.

ALT = 2094 ft, M = 0.0, The maximum bleed flow rate obtained during testing DTAMB = 34 °R was limited by choked conditions in the aircraft-EBAS 13,000 o Simulation prediction ducting. Bleed flow rates above the manufacturer's 12,500 maximum specification limit were reached during the Measured data military and maximum afterburning power test con- 12,000 Net .,- Maximum ditions as shown in Figs. 11 through 14. Thrust loss thrust, 11,500 showed the same linear characteristic as predicted be- Ib low the maximum specification limit. Operation above 11,000 the maximum specification limit was limited during spec limit 10,500 testing because of turbine life concerns.

I I BI I I 10,0000 1 2 3 4 5 Concluding Remarks Bleed flow rate, Ib/sec 92O159 A ground test at the NASA Dryden Flight Re- Fig. 14. Thrust loss due to bleed air extraction; max- search Facility was performed on the F/A-18 High Al- imum power, right engine. pha Research Vehicle to determine the bleed air avail- ability of the F404-GE-400 engines for a pneumatic 11 forebodyflowcontrolsystem.Oneobjective was to Flight Using Vortex Manipulation," AIAA 89-2192, measure the thrust loss due to bleed air extraction and AIAA 7th Applied Aerodynamics Conference, Seattle, to confirm the predicted relationship by the General WA, July-Aug. 1989. Electric steady-state simulation deck. The comparison aWilliams, D. and J. Bernhardt, "Proportional Con- of estimated and measured thrust loss showed good trol of Asymmetric Forebody Vortices with the Un- agreement. For a 1 lb]sec increase in bleed flow de- steady Bleed Technique," AIAA 90-1629, AIAA 21st mand, comparisons of the percent thrust loss between estimated and measured characteristics were within Fluid Dynamics, Plasma Dynamics and Lasers Confer- ence, Seattle, WA, June 1990. +0.5 percent of each other for all power settings tested. The percent thrust loss for a given bleed setting in- 4Ng, T.T. and G.N. Malcolm, "Aerodynamic Con- creased with higher power settings. At military power, trol Using Forebody Blowing and Suction," AIAA 91- a 1-percent increase in bleed airflow caused a 2-percent 0619, 29th Aerospace Sciences Meeting, Reno, NV, loss in thrust. The measured thrust loss above the Jan. 1991. manufacturer's maximum specification bleed flow limit 5Guyton, R.W. and G. Maerki, "X-29 Forebody showed the same linear relationship as predicted below Jet Blowing," AIAA 92-0017, 30th Aerospace Science the limit. The results encourage the use of the man- Meeting and Exhibit, Reno, NV, Jan. 1992. ufacturer's model to predict in-flight performance loss due to bleed extraction. 6Kroo, I. and L. Roberts, Forebody Tangential Blow- ing for Control at High Angles of Attack - Feasibility References Study Final Report, JIAA TR-101, Joint Institute for Aeronautics and Acoustics, Stanford University, Stan- 1Schneider, Edward T. and Robert R. Meyer, Jr., ford, CA, June, 1991. "F-18 High Alpha Research Vehicle Description, Re- sults and Plans," ISSN#0742-3705, presented at the _'Evans, Alison B., The Effects of Compressor 33rd SETP Symposium, Sept. 1989. Seventh-Stage Bleed Air Extraction on Performance of the FIOO-PW-_O Aflerburning Turbofan Engine, 2Malcolm, Gerald N., T. Terry Ng, Liane C. Lewis, NASA CR-179447, 1991. and Daniel G. Murri, "Development of Nonconven- tional Control Methods For High Angle of Attack

12

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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE | 3. REPORT TYPE AND DATES COVERED July 1992 I Technical Memorandum 4. TITLE AND SUBTITLE S. FUNDING NUMBERS Effects of Bleed Air Extraction on Thrust Levels of the F404-GE-400 Turbofan_ RTOP-505-68-30 !6. AUTHOR(S)

Andrew L Yuhas and Ronald J. Ray

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) 8. PERFORMING ORGANIZATION REPORT NUMBER NASA Dryden Flight Research Facility P.O. Box 273 H-1806 Edwards, California 93523-0273

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(EB) 10. SPONSORING/MONITORING AGENCY REPORT NUMBER

National Aeronautics and Space Administration Washington, DC 20546-3191 NASA TM-104247

11. SUPPLEMENTARY NOTES

asAIAA 92-3092 at theAIAA 28th JoL,g P_on Omfcrence, Nashville,"IN, July6-8, 1992.

12=. DISTRIBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION COOE

Unclassified _ Unlimited Subject Category 07

13. ABSTRACT (Maximum 200 words)

A ground test was performed to determine the effects of compressor bleed flow extraction on the performance of F404-GE-400 afterburning turbofan engines. The two engines were installed in the F/A- 18 High Alpha Research Vehicle at the NASADryden Hight Research Facility. A specialized bleed ducting system was installed onto the aircraft to control and measure engine bleed airflow while the aircraft was tied down to a thrust measuring stand. The test was conducted on each engine and at various power settings. The bleed air extraction levels analyzed included flow rates above the manufacturer's maximum specification limit. The measured relationship between thrust and bleed flow extraction was shown to be essentially linear at all power settings with an increase in bleed flow causing a correspond- ing decrease in thrust. A comparison with the F404-GE-400 steady-state engine simulation showed the estimation to be within +1 percent of measured thrust losses for large increases in bleed flow rate.

14. SUBJECT TERMS 15. NUMBER OF PAGES )5 Bleed air extraction, Engine simulation, FFC, HARV, Pneumatic flight control, 16. PRICE CODE Thrust loss, 'l_d_fan engine A02

17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSiFiCATION 20. LIMITATION OF ABSTRACT OF REPORT OF THIS PAGE OF ABSTRACT Unclassified Unclassified Unclassified Unlimited i NSN 7540-01-280.5500 Standard Form 298 (Rev. 2-8B) _v ..aM at,,. zao-I o 298-102