NASA CR 152055

(NASA-CR-152055) TEST DATA REPORT, 1Ow N78-19049 SPEED WIND TUNNEL TESTS OF A FULL SCALE LIFT/CRUISE-FAN INLET, WITH ENGINE, AT HIGH ANGLES OF ATTACK (Boeing Commercial Airplane Unclas Co., Seattle) 165 p HC A08/MF A01 CSCL 01A G3/02 08587

TEST DATA REPORT, LOW SPEED WIND TUNNEL TESTS OF A FULL SCALE LIFT/CRUISE - FAN INLET,

WITH ENGINE, AT HIGH ANGLES OF ATTACK

By W. M. Shain

January 1978

Distribution of this report is provided in the interest of information exchange. Responsibility for the contents resides in the author or organization that prepared it.

Prepared under Contract No. NAS2-9640 by

The Boeing Commercial Airplane Company Seattle, Washington 98125

for

AMES RESEARCH CENTER R

NATIONAL AERONAUTICS AND SPACE ADMINISTRAT 0

https://ntrs.nasa.gov/search.jsp?R=19780011106 2018-02-15T16:12:02+00:00Z

2 Government Accession No 3 Recipient's Catalog No.1. Report No

NASA CR 152055 I 5 Report Date4. Title and Subtitle

Test Data Report, Low Speed Wind Tunnel Tests of Jan 1978 a Full Scale Lift/Cruise-Fan Inlet, with Engine, s. Performing OrganizationCode at High Angles of Attack

7 Author(s) 8 Performing Organization Report No

W. M. hainT6-6145 W. M Sh~n,10 Work Unit No

9, Performing Organization Name and Address

Boeing Commercial Airplane Company 11. Contract or Grant No NASZ_-9640P. 0. Box 3707

Seattle, Washington 98124 13. Type of Report and Period Covered 12 Sponsoring Agency Name and Address Contracto'r Report

14 SponoringAgencyCodeNational Aeronautical and Space Zip 20546Administration Washington D.C.

Notes15 Supplementary

16 Abstract

A low speed wind tunnel test of a fixed lip inlet with engine, was

performed at NASA-Ames Research Center. The inlet was close coupled

to a Hamilton Standard 1.4 meter, variable pitch'fan driven by a

Lycoming T55-L-llA engine. Tests were conducted with various

combinations of inlet angle-of-attack (0-120o), freestream velocities

(0-82 m/sec), and fan airflows (70-165 kg/sec-M2 ). Data were recorded

to define the inlet airflow separation boundaries, performance The report includes descrip-characteristics and fan blade stresses.

tions of the test model, installation, instrumentation, test, data reduction and final data.

17. Key Words i(Suggested by Author(s)) n 18 Distribution Statement

High Angle-of-Attack Low Speed Inlet Test Variable Pitch Fan High Bypass Ratio Fan Inlet

19. Security Classf. (of this report) 20. Security Classif. (of this Page) 21 No of Pages

Unclassified I UnclassifiedII For sale by the National Technical Information Service. Springfield. V;rgflna 22161

22 Price4

NASA-C-168 (Rev. 10-75)

xi"

CONTENTS

Page

1SUMMARY

2INTRODUCTION

2FACILITY

2INSTALLATION

MODEL 3

INSTRUMENTATION 4

DATA REDUCTION

14TEst-kOCDURE

15TEST RESULTS

TABLES 1-3 16

FIGURES 1-17 19

APPENDIX A Q-Fan Demonstrator, Lift Fan Technology 36

Program, Blade Stress Report

APPENDIX B Tape Record Logs 90

APPENDIX C Final Data 102

APPENDIX D lest Logs 104

1:i

LIST OF FIGURES

Page

1. Ames 40X80 Foot Wind Tunnel (12.2mx24.4m) 19

2. Wind Tunnel Installation Schematic 20

3. Model Installation 21

4. Nacelle Schematic 22

5. Cowl Static Pressure Instrumentation 23

6. Fan Face Instrumentation 24

7. Fan Face (Front View) 25

8. Fan Duct Instrumentation 26

9. Aft View Fan Nozzle Exit 27

10. Compressor Face Instrumentation 28 11. Core Engine Inlet Instrumentation 29

12. Core Engine Nozzle Instrumentation 30

13. Nacelle Operational Health Monitoring Locations 31

14. Block Diagram, Steady State Data Acquisition System 32

15. Lift/Cruise Fan Inlet Test Instrumentation System 33

16. Data Sheet Format 34

17. Nacelle Force Measurements - Nomenclature and Sign Convention. 35

http:12.2mx24.4m

LIST OF TABLES

Page

1. Magnetic Tape/Data Recording Channel Assignments 16

2. Lift/Cruise-Fan Inlet Test, Data Reduction Tables 17

3. Summary of Test Conditions 18

iv

REFERENCES

1. Syberg, J; Koncsek, J. L.; "Low Speed Tests of a Fixed Geometry Inlet for a Tilt Nacelle V/STOL Airplane" Boeing Document D180-20276-1 dated January 1977, (NASA CR-151922 dated January 1977).

2. Syberg, J.; "Test Planning and Coordination Lift/Cruise Fan Inlet Test Program" Boeing Document D180-20725-l dated July 1977.

3. Syberg, J; "Low Speed Tests of a High Bypass Ratio Propulsion System with an Asymetric Inlet Designed for a Tilt Nacelle VSTOL Airplane" Boeing Document D180-22888-1, November 1977, (NASA CR-152072

dated January 1978).

SUMMARY

A low speed wind tunnel test of a fixed lip inlet operating at high

angles-of-attack was performed in the NASA-Ames Research Center (NASA-

ARC) 12.2 X 24.4 meter wind tunnel (40 X 80 WT). The purpose of the test was to demonstrate total pressure recovery and distortion levels which result in acceptable fan/core engine stall margins and blade stress levels. These characteristics would be demonstrated when operating at combinations of high angles-of-attack, freestream velocities and engine airflows. The inlet was designed and built by Boeing, was asymmetric, with a contraction ratio varying from 1.30 (leeward) to 1.76 (windward). It.was matched to a Hamilton Standard-1.4 meter (55") fan driven by a Lycoming T55-L-lIA gas turbine engine. The test model (nacelle) was mounted on its side on a turntable, approximately

3.8 meters (12.6') above the wind tunnel floor.

The nacelle contained approximately 225 parameters measured/recorded by a Boeing supplied data system: Measurements were also made of the nacelle lift, drag, and side forces and the pitch, yaw and rolling moments. Key operational and monitoring parameters were also recorded in real time on magnetic tape recording systems. The model performance

data were reduced on site into a semi-final form for test progress

analysis and the tape recorded data were available for, if needed, historical analysis.

Tests were first run to define the inlet airflow separation boundaries as a function of angle-of-attack. An airflow and tunnel velocity was set and the model slowly pitched until separation occurred within the inlet. Data were recorded at discrete points approaching the separated

conditions as defined by monitoring the dynamic pressure activity at the windward side of the fan face. Data were also recorded later defining the angle-of-attack and airflow hysteresis (attached/ separated/reattached) characteristics. Following definition of the operating

placard, tests were run at fixed angles of attack and velocity with varying airflow to define in detail the inlet performance characteristics.

The model occupied the wind tunnel for five weeks, three of which were required to repair the wind tunnel after a drive motor failure. Tests were done on a single shift basis and covered angle-of-attack ranges from 0 to 120 degrees, freestream velocities of 0 - 82 meters/ ec (160

kts) and fan airflows of 70 - 165 kg/sec-MC (14 - 34 lb/sec-ftC). Approximately 240 data conditions were recorded during 20 hours of engine operation. Included in the report are detailed descriptions of the model, installation, instrumentation and test procedures. Also included are the test, engine and instrumentation logs along with the final model performance and blade stress data.

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INTRODUCTION

During the first part Of 1976, The Boeing Aerospace Company, Military

Airplane Development was awarded, by NASA-Ames Res'earch 'Center, an inlet

design contract in support of a multi-mission VSTOL airplaiej contract

NAS2M215, The performance of a Boeing fixed lip inlet would 'be

demonstrated on a Hamilton-Standard Q-Fan @ demonstrator/T-55 engineOperating in a severe "angle-of-attack" environment provided by the 40 X 8b WT. Boeing was the prime contractor responsible for the inletand nabelle design, fabrication and assembly; installation; windtunnel test and data analysis. The responsibility for supplying the core engine/fan system along with its operation during test Was ubcontracted to Hamilton-Standard (HS).

The main objective would be to demonstrate that a fixed lip inlet ca provide adequate pressure recovery and distortion levels that result

in acceptable core engine/fan stall margins and fan blade stress

levels at combinations of large nacelle tilt angles, freestream velocities and engine airflow levels.

Initial testing in the wind tunnel began in July 1976. After several days of testingi a gear failure in the fan gearbox temporarily ended

t~sting. The test was resumed, after gearbox repairs, in September ot

.19-6. Again after several days of testing the same gear failed andseverly damaged the core engine. The test was again stopped without

acquiring enough data to meet the original objectives. Details of these tests were reported in Reference 1.

inJune l77, a new contract, NAS2-9640, was issued as a follow-oh to complete the initially planned test series. The HS gearbox was

analyzed, redesigned, remanufactured and verification tested to assure

structural integrity. Testing was be resumed in the fall of 1977.

FACILITY

Testing was performed in the NASA-ARC Wind Tunnel. The Wind Tunnel

has a closed 12.2 by 24.4 meter (40' X 80') test section with semicirculdr sides of 6.1 meters (20') radii, and a closed circuit airreturn passage. The general arrangement is.shown in Figure 1. Air isdriven in the wind tunnel by six 12.2 meter;(40') diameter fans which are powered by six, 6,000 horsepower electrfcmotors. The tunnel

Operates with a stagnation pressure equal to atmospheric. The stag.

nation temperature varies from ambient upwards, due to the entrained prdducts of combustion and the heat-from the tunnel drive system.

INSTALLATION

schematic of the test model, installed in the wind tunnel, is shown inFigure 2. The main wind tunnel model-support struts were removed and

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the semi-span turntable installed for mounting the nacelle. The nacelle was bolted atop a Boeing designed pylon-strut which in turn was bolted to the turntable. This entire assembly was mounted "on balance" for measuring the model forces. A large fairing was designed and built to fit around the strut and turntable and mounted "off balance" to provide shielding from the Wind Tunnel air forces. The centerline of the nacelle was 3.84 meters (12.6') above the Wind Tunnel floor and located on the vertical centerplane of symmetry in the Wind TOnnel.

The center of the installation, or center of rotation, was at tunnel station 261.5. A model alignment check was done after installation and the model was determined to be 0.20 nose down at 00 angle of attack. No correction was made for this slight deviation from the horizontal.

The rotation of the semi-span turntable is in the horizontal plane (normal inlet yaw) and since the inlet was asymmetric the inlet was installed on its side, i.e., the 900 position on the inlet.lip was up inthe wind tunnel. This exposed the windward designed side of the inlet to the tunnel flow at angles of attack. Figure 3 shows the model installed in the wind tunnel.

The peripheral support equipment, other than the instrumentation tystems, were mainly hydraulic and lubrication supply systems for the variable pitch fan. One high pressure pump, reservoir and cooler were located on the first floor, with a gear box lubrication supply and scavenging pump (2)located on the second floor. The high pressure pump supplied fluid for the fan blade pitch change and control system

and the engine power lever position. The lube and-scavenge pumps supplied and scavenged the fan gear box of lubrication oil. Fuel for the engine (JP-5) was supplied from the facility system.

The onboard fire system consisted of manifolded nozzles within the core engine cowling attached to two high pressure nitrogen bottles. Inthe event of an external engine fire the cavity inside the cowling would be filled with inert gas (N2). More detailed information regarding the installation support equipment may be found in the Plan of Test, Reference 2.

MODEL

The test model, or nacelle, consisted of an inlet, a variable pitch -fan, a gas turbine core engine, and the appropriate fairings, nozzles etc.

The inlet has a 1.469 meter (57.826") highlight diameter, a 1.2 meter (47.236") throat diameter and a 1.397 meter (55") fan face diameter. The inlet contours are asymmetric with the windward side (180') having a higher cbntraction.ratio than the leeward side (0). The contraction

-3-lo OFPOOR .QU.ALIT-3-IAL PAGE IS

ratio varies from 1.76 at 1800 to 1.30 at 00. At a given inlet station, both the internal and external contours are circular in cross sectiof with offset centers. The inlet cowl was made of fiberglass.

The Hamilton-Standard Q-Fan @demonstrator is a 1.397 meter (55"), 13 bladed, variable pitch fan which utilizes a Lycoming T55-L-llA, 3750 hp gas turbine as the core engine. The fan has 17:1 bypass ratio and is driven through a 4.75:1 gear reduction to a maximum speed of 33Q5 rpm. The fan system used a 645 meter (25.4") diameter "semi-elliptical" nose dome fairing. The fan exit nozzle was a simple, round, constant area, aluminum nozzle with an exit area of 1.064 sq. meters (1649 square inches). The core engine exit nozzle supplied with the engine, had an exit area of .254 sq. meters (394 square inches). A schematic of the inlet and nacelle showing the major components and station designations isshown in Figure 4.

INSTRUMENTATION

The test model instrumentation wds divided into three groups: 1) Model performance, 2) Fan/engine operation and health and 3) Wind Tunnel condition. A brief description of each group follows:

Model Performance

INLET: The inlet contained 45 cowl surface static pressure ports, Figure 5, and seven fan face total pressure' rakes (70 total pressures, 7 statics and 3 flush mounted total pressure transducers,) Figure 6. These rakes are also shown in the photograph, Figure 7. One flush mounted Kulite was located near the inlet throat at 180' in the cowl wail.

FAN DUCT: The fan duct had two rakes located just ahead of the fan nozzle exit, Figure 8. The rakes each contained 10 total pressures, 3 total temperature probes, one Prandtl-type static probe (two flush surface statics were also on the core engine cowling) and are shown in Figure 9.

CORE ENGINE: The core engine compressor face was instrumented with ,eight,6 probe total pressure rakes (44 total pressures and 4 flush mounted transducers), and 8 static pressure ports, Figure 10. Three total pressures and one static were located at 00 on the core compressor inlet lip, Figure 11.

,Engine performance was determined 'by measuring NI, N2, and power shaft torque. The engine nozzle contained four static pressures, Figure 12.

,Fan Engine Operation and Health Monitoring

'INLET: Inlet performance was monitored on-line through useof one X-Y -and two X-YY plotters. The X-Y plotter was used to monitor the fan face radial total pressure 'profile at 1800 (rake No. 4). The two X-YY

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plotters were used for monitoring RMS peak-to-peak pressure dynamics versus model angle-of-attack or airflow; one plotter for two of the fan face/inlet Kulites and the other for two of the compressor face Kulites. All eight of the Ktlites (PDFI-3; PDCl-4; PDS) were also visually monitored on RMS meters. Three of the seven fan face total pressure

rakes were strain gaged near the root to monitor stress levels during

testing.

FAN: Fan operation and health was monitored basically by RPM, blade angle, blade stress (5 blades were strain gaged) and vibration. In addition, the gearbox lube oil pressure, temperature and flow rate were displayed along with other key temperatures and vibrations. Several (6)

proximeter probes, located in the fan gear box, were used to monitor spacial displacement of the sun and ring gears.

CORE ENGINE: The engine system contained all the normal monitor and control parameters, NI, N2, TT7, oil pressures, level and temperatures (various locations), vibration pickups, power lever angle and fuel supply -pressures. Inaddition engine external structure and cowl cavity temperatures were displayed on panel meters.

The various temperature sensors, vibration pickups and proximeter probe

locations used inmonitoring the nacelle system operation are shown schematically in Figure 13.

Wind Tunnel

The Wind Tunnel instrumentation consisted of measuring the total pressure,

dynamic pressure, and the total temperature. In addition the model lift, drag and side forces, pitch, yaw and rolling moments were measured by the facility balance system.

Data Systems

The various test parameters were recorded on several data systems.

Model performance data on a Boeing steady state data acquisition system,

wind tunnel and force information on the facility data system, fan operation and health on a HS tape recording system and select, representative parameters for system diagnostic studies, should they be needed, on a Boeing tape recording system. Engine operational parameters were also hand recorded on the engine logs.

The steady-state data acquisition system, Figure 14, consisted of the various sensors, signal conditioners and a Hewlett-Packard (HP) 3052A automatic digital data system. The HP 3052A contained a 40 channel scanner, digital voltmeter, programmable calculator/controller, line printer and a paper tape punch. Data were recorded on magnetic tape (casettes), punched paper tape and a line printout.

The various wind tunnel condition parameters and force balance data were recorded on the facility data system. These values were continuously

displayed on-line and output on punched.paper tape for each test condition.

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The two tape recording systems were utilized for recording real-time data defining and/or monitoring critical operational/conditional parameters.

Table 1 lists the various parameters recorded and their recording location.

The HS system was used to record the fan blade strain gages (stress) and

enough peripheral information to define operating conditions. This data

was then used in developing the "Blade Stress Report" contained in

Appendix A. These tapes were retained and stored by HS. A tape log of

the contents of these tapes is also contained in the "Blade Stress Report".

The main purpose of the Boeing tape recording system was to provide a

real time record of the transient flow conditions within the fan and This data could then be used in analyzing any failure orcore engine.

The data, as of yet, hasanomalies which might occur during testing.

not been required for any analysis. The tapes remain stored at NASA

.ARC. Tape logs of their identification, content and calibrations are contained in Appendix B.

A complete schematic of the data acquisition and monitoring system is shown in Figure 15.

DATA REDUCTION

The punched paper tapes from the Boeing-steady-state data acquisition system and the wind tunnel data system were input to the NASA IBM 360 computer system. The data were calibrated, combined and then calculated for on-site, "short-turn-around" data. The original data tapes were then returned to Seattle for final editing, reprocessing, conversion to Standard International (S.I.) units and plotting. This data has been reproduced and is contained on microfiche as Appendix C.

The data were processed inSeattle with program PN026. The following pages list the basic equations and constants. Item numbers refer to the prinitout location shown in the "Data Sheet Format", Figure 16.

1. Test Number - Boeing Test Identification 2593. NASA-ARC Test 513

2. W/T.inputs VO - Wind Tunnel Velocity - KTS; Alpha -

Nacelle angle-of-attack-Degrees; Q-Dynamic pressure, PSF; PTO-Tunnel Total pressure psia; TTO - Tunnel Total Temperature -

OR; Tunnel Static pressure, psia PO = PTO QPSF/144

3. Barometric Pressure - Reference pressure, psia: All recorded pressures were referenced to barometric PACTUAL = PBAR + S/V reading.

4. Inlet corrected Airflow, lbs/sec

WK1 = WKlA * 12.98

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WKIA = 46.315*CDT/PTFA, lb/sec-ft2

4a.

CDT was obtained from model scale data where the inlet airflow was calibrated as a function of the average inlet throat static pressure PAV

PAV = (PC13 + PC33 + PC37 + PC38)/4PTO

PAV vs CDT - Table 2

PTFA = item 6

5. Inlet Airflow, lb/sec

W1 = WKl * 6TI T- ORIGINAL PAGE IS OF POOR QUALITYwhere

6T TTQ/518.688

6T PTO/14.696

6. Fan face average total pressure recovery, area weighted.

PTFA = PTFAV/PTO 10

6a. PTFAV 1/1869.12 P PTFRn (A ) psia

n=l 7

PTFR'= 1/7 E PTFXn n=l

6b. PTFXn = Individual fan face total pressures, number x,

An = Area Weighting factor

n A-in 2 PTF-X

1 93.462 1, 11, 21, 31, 41, 51, 61 2 93.462 2, 12, 22, 32, 42, 52, 62 3 280.362 3, 13, 23, 33, 43, 53, 63 "4 280.362 4, 14, 24, 34, 44, 54, 64 5 280.362 5, 15, 25, 35, 45, 55, 65 6 280.362 6, 16, 26, 36, 46, 56, 66 7 280,362 7, 17, 27, 37, 47, 57, 67 8 93.462 8, 18, 28, 38, 48, 58, 68 9 93.462 9, 19, 29, 39, 49, 59, 69

10 93,462 10, 20, 30, 40, 50, 60, 70

7. Fan face total pressure distortion

DISF = (PTFXMAX - PTFXMIN)/PTFAV

-7

http:1/1869.12

8. Fan duct airflow, lbs/sec

A01i PTMn PM y PMn - -~

1 /2. TRY,:M ~ ) t' I n) ~ ] _ _P _20 _ _

Y = 1.4015 2

g = 32.1741 FIft/sec

R = 53.35 ft-lb/lb0R

8a. PTM = Fan duct nozzle total pressure

8b, PM Fan duct nozzle static pressure: 28.73R-18.61

for 1< n < 10 = (PM2-PM) (Rn10.65 ) + PM?-Rn 18.91

Rn - 18.61

for 11 < n < 20 = (PM4-PM3) Rn .6 ) + PM410.65

A. = f(PTMn)

in? -n - RadiusAi

1,11 117.1 28.73 2,12 110.5 27.4 3,13 99.8 26.16 4,14 90.7 24.98 5,15 82.5 23.86 6,16 75. 22.79 7,17 68.4 21.77 8,18 63.5 20.78 9,19 58.5 19.83

.10,20 60.2 18.91

6 8c. F- (6 L TTMn)/518.688n=l

TTM = fan duct nozzle total temperature

-8

9. Fan duct corrected airflow, 1bs/sec

WKF = WF * VF/ 6 F

where 6F = PTMAV/14.696

9a. PTMAV = .992

[6 9

1649

_ -0 (PTMn

n=l

+2piPTMn+OA

n 2 n+1O)A ] psia A = two times the values shown in A. (8b) table, for fl 1

values 1 < N < 10

10. Core engine compressor face airflow, lb/sec

8 WE = E WER

n=l

16.96 * PTCARn J7 F-

2.TT LYP n 2 1/2

R 1+

PTCAR = The individual core engine rake average pressure.

10a. lOb.

11.

TTC = Core engine compressor face total temperature, OR PSC n = Core engine compressor face static pressure aligned

with each core engine rake arm.

Core engine compressor face corrected airflow.

WKE = WE *-(E/ 6 E

SE= TTC/518.688

6 E= PTCAV/14.696

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11. Continued

PTCAV = Area weighted average total pressure at the compressor face, psia

48 PTCn

= E 48n=l

lla. PTC Individual total pressures (48) at the core engine compressor face

12. Area weighted average total pressure recovery, core engine compressor face.

PTCA = PTCAV/PTO

13. Core engine compressor face total pressure distortion

DISC = (PTCMAX - PTCMIN)/PTCAV h PA E Is OF poRI Q TL

14. Core engine horsepower

EP = 2 iT* ET * N2/33000

14a. where ET = Core engine torque (Table 2)

N2 = Power turbine speed

15. Corrected core engine nozzle thrust - table look-up (Table 2), lbf FKN vs. corrected engine horsepower (CP)

CP = EP/C VE * E

16. Core engine nozzle thrust, lbf

FN = FKN *6 E

17. Core engine nozzle exit velocity, ft/sec

'VN = FN * (g/WE) + FF

where FF is a table look-up (Table 2) FF vs. KNI KNI Corrected compressor speed, Nl.

1.8. Calculated inlet airflow, lb/sec

E 2PPA,) ]PPA 1/270 A*PTFn 2

=l V -PT

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18. Continued

A = An/7 (from the An table used in the calculation of PTFAV)

PPA = Interpolated average static pressure for each fan face total pressure probe.

27.228

= (PCz - PPy) * (RA - 13.797) RA t

13.703 +.PPy 13.237

18a. PC = Inlet cowl compressor face static pressure

18b. PP = Fan face rake prandtl static pressure

N z -y

l n l0 39 1 11 n 20 40 2 21 n 30 41 3 31 n 40 42 4 41 ! n ! 50 43 5 51 n 6Q 44 6 61 5 n 70 45 7

19. Calculated inlet corrected airflow, lb/sec

WK2 = W2* -(-T/6 T

20. Total fan face airflow Ccalculated), lb/sec

W3 = WF + WE

21. Fan pressure ratio

FPR = PTMAV/PTFAV

22. Fan duct exit velocity, ft/sec

Y-l1 1/2 * TAY PMAV v

VM 2gR 1 -PH~ ' TTMAV Average fan duct nozzle total temperature

1 6

TTMAV 6- TTMn

PMAV = Average fan duct nozzle static pressure.

4 1/4 L PMn

n=l

.-II.

23. Corrected compressor speed, rpm

KNI NI/ f- E

24. Corrected power turbine speed, rpm

KN2 =N2/f"E

25. Fan inlet cowl surface Mach number

Yl1 1/2

MCn PTO 1 n ci thru 45 [;1F\Cn/

26. Force Balance (reference Figure 17)

FX = Nacelle drag forcelbf QX Nacelle roll-moment,ft. lbf

FY = Nacelle lift forcelbf QY = Nacelle yaw moment,ft. lbf

FZ = Nacelle sideforcelbf - QZ = Nacelle pitching foment,ft.

FXY " Resultant nacelle axial force = (FX*FX+FY*FY)l/2, lbf

27, Fan blade angle b - degrees 43-170

28. Core engine power lever angle - degrees 0-100*

29. Fan and compressor face. Kulite pressures - RMS PSI/PTO

30. Steady state data acquisition system measured tunnel total and static pressure and total temperature (for use as a backup)

Note: Individual pressure measurements tabulated on data sheets (items 6b, 8a, 8b, lOb, lla, 18a, 18b, 25a,) are ratioed to PTO.

31. Core engine exhaust gas temperature OR.

Ip addition to the Max minus Min-over-average distortion for the core engine, an "Allison Distortion Index" was calculated from the compressor fae total pressure data. This data were output on a separate page. The index is calculated using the following procedure. First the rake average total pressure for each of the eight rake arms are calculated. Ne t the rake average pressures for 12 imaginary arms are calculated. These pressure have the following relationship to the actual recorded Rtessures. -AF,

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Imaginary = 5/6 of the average of Actual + 1/6 of the average of Actual Rake No. (A) Rake No. (B) Rake No. (C)

Rake No. (A) Rake No. (B) Rake No. (C)

l (30 deg) 8 1 3 (90 deg.) 7 6 4 (120 deg) 6 7 6 (180 deg) 5 4 7 (210 deg) 4 5 9 (270 deg) 3 2

10 (300 deg) 2 3 12 (360 deg) 1 8

Imaginary = 1/2 [Average of Actual + Average of Actual1 Rake No. (A') Rake No. (B') Rake No. (C')

Rake No. (A') Rake No. (B') Rake No. (C')

2 C60 deg) 7 8 5 (150 deg) 5 6 8 (240 deg) 3 4

11 (330 deg) 1 2

From these rake average pressures, the twelve contiguous 120 degree sector average pressures are calculated. The minimum 120 degree sector isthen located from this array.

The average compressor face total pressure (PTCAV) is calculated as the arithmetic average of the original 48 recorded pressures. Since ring

number 3 lies close to the radius that separates the outer 40 and inner 60 percent of the total compressor face area the outer 40 percent average

total pressure (PTCRPA) iscalculated as the arithmetic average of rings

1-3 and inner 60 percent total pressure (PTCRFA) as the arithmetic average

of rings 3-6. Radial total pressure distortion is then calculated as

KR = PTCRPA - PTCRFA/PTCAV

The circumferential total pressure distortion iscalculated as

KTHETA = PTCAV240 - PTCAV120/PTCAV360

The composite distortion index is calculatei as

KCOMP = (KR + KTHETA 2)1/2

All of the data were calculated using the formula and definitions in the preceeding pages. All-constants, etc. were in English units. The final

conversion to SI units was done by a multiplication factor applied to the computed or calibrated English value. These conversion factors are listed inTable 2.

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24

TEST PROCEDURE

Testing was done on second shift, with first shift being utilized for troubleshooting and model maintenance. For safety considerations, due to the previous experiences with the test model, engine running was not permitted before 5:00 p.m. At the beginning of each day's testing and/or run startup the data systems were all check calibrated and

-adjusted for the proper barometric readings. A wind-off zero was taken on the force measuring system. The following general sequence was then followed:

1. Start Wind Tunnel into the synchronizing mode.

Start fan gearbox scavenging pump.

3. Start fan gearbox lubrication pump.

4. Start fan blade and PLA control hydraulic pressure pump.

5. A final inspection of the facility and model test systems was completed.

6. The tape recorders were turned on.

7. The engine was started.

8. - The Wind Tunnel access door was closed.

9..: The. WJind- Tunnel was- then -brought -on-line.

10. The engine was left at idle power-until the Wind Tunnel speed was within-approximately 60% of the end value. The engine/fan system was set at an intermediate-to-high power setting (usually around N = 12000 rpm) and the nacelle was yawed to the desired angle-ofatack. The engine and wind tunnel were then trimmed to the required setting.

11. After 20-40 seconds stabilization time data were recorded on both

the wind tunnel and steady-state data systems. (The tape recorders'

were left on-continuously during testing and each run/condition number was announced on the tape.)

12. The model was then reset to a new angle of-attack and/or the air/flow was changed for the next condition number.

Detailed logs were kept during testing defining the wind tunnel, fan-engine, and data conditions. These logs are included as Appendix D.and include "Test Logs'i, 'tEngine Logs", and "Tape Record Logs'1. Detailed information noting the pertinent parameters and ,any erroneous or peculiar data or conditions are recorded.

-14

TEST RESULTS

The test nacelle was assembled and instrumented inSeattle in July, August, 1977. A functional check of the data systems was also perfbrmed at this time. The model was shipped'from Seattle, completely ready to install inthe wind tunnel. Wind tunneloccupancy began on September 12, 1977, with testing beginning on September 16. After a "static conditions" airflow sweep (full power to idle power) testing were done to define the low airflow inlet separation boundaries with varying angle-of-attack. The airflow (fan rpm, Blade angle) was set a constant value and the nacelle was rotated (increasing alpha) until airflow separation occured within the inlet. This was determined by on-line monitoring of the windward side fan face total pressure profile and the dynamic pressure fluctuations. This was done at nominal tunnel velocities of 20, 38, 46, 54, 64, 72 and 82 meters/sec (40, 75, 90, 105, 125, 140 and 160 kts). These tests were then followed by more definative testing at a fixed angle-ofattack and velocity with varying airflow rates.

On September 23, 1977, testing was interrupted by a failure in one of the tunnel main drive motor/generator sets. The repair took approximately three weeks and testing resumed on October 13, 1977. The test was finished on October 15, 1977. During the latter portion of testing data were recorded with the airflow separated to the inlet lip providing none of the prescribed system limits or stresses were exceded during the time-oncondition. Prior to this, time data had been recorded only up-to-near lip

separation. The hysteries characteristics of these separated airflow conditions were evaluated both by increasing/decreasing angle-ofattack and decreasing/increasing rpm (separated/attached). A detailed analysis of the data, inlet performance characteristics and specific

test results may be found in Reference 3. Table 3 summarizes the conditions, velocity and angles-of-attack tested and the applicable test run numbers.

-15

TABLE 1 MAGNETIC TAPE/DATA RECORDING CHANNEL ASSIGNMENTS

PARAMETER TAPE TRACK

Boeing Tape Recording System

Fan Face Dynamic Pressure PDFl 1 Fan Face Dynamic Pressure PDF2 2 Fan Face Dynamic Pressure PDF3 3 Cowl Static Dynamic Pressure PDS 4 Compressor Face Dynamic Pressure PDCI 5 Compressor face dynamics pressure PDC2 6 (Reels 1-6) Model Angle of Attack a 6 (Reels 7-11) Compressor face dynamic pressure PDC3 7 Compressor face dynamic pressure PDC4 8 Fan Blade Strain Gage 9 Fan Blade Strain Gage 10 Power Turbine Speed N2 11 Fan Blade Angle a 12 Fan Face Pressure Rake Stress 13 IRIG "B"Time Code 14

VOICE EDGE TRACK

Hamilton Standard Tape Recording System

Fan Blade Stress #1 Bending I Fan Blade Stress #6 Bending 2 Fan Blade Stress #7 Bending 3 Fan Blade Stres-s #2 Vee 4 Fan Blade Stress #9 Vee 5 Fan gear vibration fore &aft Ot. T AG IS 6 Fan gear vibration lat. OR1GIA] ,A1pTY 7

0N2/Sta. MPX #1 O . 1 L 8

Sta. MPX #2 9 Torque 10 Fan Blade Angle 11 Fan Speed - IP 12 Voice

Proximity Probe #1 - Ring gear 1 #4 - Ring gear 2 #7 - Retaining Nut 3 Multiplexed #6 - Sun gear 4

#10 - Retaining Nut 5 Gearbox vibration vertical 7 Gearbox vibration vertical 9 Inlet vibration vertical 10 Inlet vibration horizontal 11 Proximity probe #9 Sun gear 12

-16

TABLE 2

LIFT/CRUISE - FAN INLET TEST, DATA REDUCTION TABLES

Inlet Corr. S.I. Conversion Airflow Vs. Calibration Pressure

PAV CDT

Engine Torque Vs. Torquemeter Indication

Indication' ET % of

-

Corrected Fuel Flow vs. Corr. Engine Speed

r77N/ V-9- WF/6Vo. .

Corrected Thrust Vs.

Corr. Power

EPK FKN

From

KTS

Constants

To

M/SEC

Multiply By

.5144 1300 lb-ft inch-lb: %18720 RPM lb/hr hp lb FPS M/SEC .3048

.55 .961 10 15

1550 2350

' 55% 60

340 390

0 200

0 30

lb/ft2

lb/in 2 BAR BAR

.0004788

.068948 .60 .937 20 3150 65 450 400 53 lb/sec Kg/sec .45359 .65 .904 25 3900 70 520 600 73 OR OK .55556 .70 .860 - 30 4600 75 620 800 91 ft-lbf Newton.75 .806 35 5300 80 750 1000 1 Meters 1.356 .80 .739 40 5900 85 940 1200 123 lbf Newton 4.448 .85 .'654. 45 6550 90 1220 1400 138 HP, 2 KW 2 .7457 .90 .546 50 7150 95 1620 1600 152 lB7sec-ft Kg/Sec-M 4.8826 " .925 .478 55 7750 100 2120 1800 165 .950 .390 60 8350 -2000 178 .975 .250 65 9000 2200 191

1.0 0 70 9650 2400 203 75 10350 2600 215 80 11150 2800. 226 85 12000 3000 238 90 13000 95 14150

100 15450

TABLE 3 SUMMARY OF TEST CONDITIONS

Velocity Angle-of-Attack Run Number M/Sec (KTS) Degrees No. of Airflows

3 4

0 20.6

(0) (40) 90

0 - 120

6* Const. (1)**

29 90 3 30 120 9 23 30.9 (60) 75 5 24 90 7 25 105 7 5 38.6 (75) 70 - 94 Const. (3) 9 0 5

38 20 3 10 " 60 5 I 75 6 13 90 12 31 2,

46.3 (90) 70 - 81 60

Const. (1) 8

27 75 9 28 90 5 6

12 54.1

" (105) 50 - 95

0 Const. (4)

5 37" 20 3 14 60 7

15,-l 75 6 32 64.4 (125) 48 - 84 Const. (5) 17 " 60 8 7 72,. (1401 40 - 68 Const. (5)

8-19. " 0, 9 35 " 20 3, 20, " 45 17 36 60 5 33 34- 84.4 (160)

70 39

- 71 - 63

Const. (1)Const. (4)

* Airflow was u ually varied betwe n nominal vail,ues 2of 73.2 kg/sec M?

(),5, lbb/sec ft) and 1:66 Kg/sec M - (34 lb/sec ft ),

** Constant airflow conditions with varying angle of attack ),= Number of airflows at which an alpha sweep was done

IS-AGE

OF pOOR QUAITY *ORIGINAL

-18

TOP VIEW

__.69m

Lo

(23')R 12.2m (40') DIA TURNING VANES NVANESCHO D

1.83mn (6') CHORD

Io

E

.4

zo

E

..............

264.6m (868')

TEST SECTION

12.2x24.4m (40'xBO')

SIDE VIEW

20.9m (68'-7")

Figure I Ames 40x80 Foot Wind Tunnel (12.2mx24.4m)

PRIMARY NOZZLE

00 180

FAN YOKE FAIRING BALANCE-MOUNTEDPYLON

PYLON FAIRING(OFF-BALANCE)

Figure 2. Wind Tunnel Installation Schematic

C=

FRONT VIEW a--30

a-0AFTVIEW

a= 12 o

a=6 o

Figure 3. Model Installation

OF POOR QUA

n1398 64.991

ASRAKES1270,11 *O&SCI2

FACE

RAKES(1H

(7(M2 1-i

C-F.

STA FAN DUCT

/VTIM

STAONDSINT

WINDWARD SIDE LIP 11F Of Rotation and Moment Center

Figure 4L IA"ca Schematic

LEEWARD (00) COWL STATIC PRESSURES

WINDWARD (1800) COWL STATIC PRESSURES SIDE COWL STATIC PRESSURES

NO. X/RFAN RL/R,AN S/RFAN NO. X/RFAN RIR,,, S/RA, N / RLRFAN,

1 .1242 1.0650 -.1628 19 .1242 1.2602 -.1932 37 .4921 .8588 ,00

2 .0466 1.0300 -.0775 20 .0466 1.2227 -.107 38 .4982 .588 2700 3 .0137 1.0036 -.0352 21 .0137 1.1849 -.0563 4 .0021 .9844 -.0127 22 .0021 1.1 527 -.0220 *STEADY STATE & DYNAMIC PICKUP 5 0 .9719 0 23 0 1.1308 0 (FLUSH MOUNTED KULITE, PDS) 6 .0007 .9640 .0079 24 .0007 1.1183 .0125 7 .0042 .9519 .0205 25 .0042 1.0999 .0313 FAN FACE COWL STATIC PRESSURES 8 .0109 .9399 .0343 26 .0109 1.0811 .0513 A 9

10 .0209 .0500

.9280

.9057 .0499 .0865

27 28

.0209

.0500 1.0619 1.0244

.0729

.1204 PC

NO X/R R R FAN

/R L FAN _ )

11 12

.1048

.1817 .8810 .8620

.1468

.2261 29 30

.1048

.1817 .9781 .9342

.1923

.2809 3911.5931 a1.0 15.70

14 .4445 .8573 .4895 * 2 .4445 I .8619 .5545 al11. 15 .5954 .8709 .6410 33 .5954 .8524 .7058 42 170,0 16 .8136 .9011 .8613 34 .8136 .8691 .9248 43 221.4 17 1.0317 .9368 1.0824 35 1.0317 .9084 1.1465 44 272.9 18 1.3590 .9836 1.4130 36 1.3590 .9741 1.4803 45 324.3

-- - FAN FAN FACE

RFAN =.6985 m

HI-LITE t

Figure 5. CoiS Static Ptowure Instrumentation

FAN FACE RAKE PROBE COORDINATES

AND NU1BERING

RAKE RAKE RAKE RAKE RAKE RAKE RAKE RING % AREA1 2 3 4 5 6 7 RADIUS FOR R/RFAN RING

RING 1 PTF 1 PTF 11 PTF 21 PTF 31 PTF 41 PTF 51 PTF 61 .9901 5%RING 2 2 12 22 32 42 52 62 .9700 5%RING 3 3 13 23 33 43 53 63 .9286 15%RING 4 4 14 24 34 1 44 6454 .8627 15%RING 5 5 15 25 35 45 55 65 .7914 15%RING 6 6 16 26 36 46 56 66 .7129 15%RING 7 7 17 27 37 47 57 67 .6247 15%RING 8 8 2818 38 48 58 68 .5582 5%RING 9 PTF 9 PTF 19 4F 29 PTF 39 2 PTF 49 PTF 59 PTF 69 .5218 5% PPI PP2 PP3 PP4 PP5 PP6 PP7 .5017 --RING 10 PTF10 PTF 20 PTF 30 PTF 40 3 PTF 50 PTF 60 PTF 70 .4813 5%

RAKE ANGLE 25.7 77.1 128.6 180.0 282.3213.4 334.3 (deg)

1 POF 112 PDF 2 Dynamic total pressure probe mounted side by side with steady state nrobe 3 PDF 3 TopNo

PC 40 / 10OTYPICAL

VIEW LOOKING AFT

RAKE151.430 TYPICAL

PC 42

LEEWARD 0 1800 WINDWARDRAKE 4

Figure 6 Fan Face Instrumentation

F

4!

..~~RAKE*

Figure 7 Fan Face (Front View)

OIGINAL PAtt25 Of POOR UALt

00 180 RADIUS L/AR PROBES PROBES R/R FAN

1.0676NOZZLE WALL NOZZLE WALL

PmI PM3 1.0640

PThl1 1.0447 .142PTMI PTM2 PTM1 2 .9964 .134

TTMI TTM4 .9738 PTM3 PTM13 .9513 .121 PTM4 PTM1I4 .9084 .110 PTM5 PTM1I5 .8676 .100 PTM6 PTM16 .8287 .091 T112 TTM5 .8102 PTM7 PTM17 .7916 .083 PTh8 PTM18 .7556 .077 PTM9 PTM1I9 .7211 .071

TTM6 .7044TTM3 .6876 .073PTMI0 PTM20

CORE CASE CORE CASE .6676

AA: Area assigned to total pressure probe 2 = 1.064 mFlow area at rake faceAR:

FAN DUCT RAKES, 0 & 180

F LOW// FO 0.456 m TO NOZZLE EXIT PLANE "

PM3 (1750)

Figure 8. Fan Duct Instrumentation

26

Figure 9. Aft View Fan Nozzle Exit

IGINAL PAGE 15 OF POOR QUALM'

27

VIEW LOOKING FROM FRONT TO REAR

- ALL PT EXCEPT AS NOTED

P-7

Ps-Psi((I a

SS

KU LIT E

Figure 10 Compressor Face Instrumentation

28

PTF

Figure 11. Core Engine Inlet Instrumentation

,N PN 2 1

180 + 00

-1 .35"' 0 PON 3 PN4

2700

Figure 12. Core Engine Nozzle Instrumentation

C

INLET DUCT VIBRATION HORIZ. & VERT.

GEARBOX HORIZ. & VERT.

X

PROXIMITY PROBES SUN GEAR RING GEAR

RETAINING NUT

;;

'GEARBOX VIBRATION ., HORIZ. & VERT. HMOUNTRING

VIBRATION HORIZ.

O F.PO

~TURBINE VIBR

RIGHT

S HORIZ.& VERT.

-W-THERMOCOUPLE

Figure 13. Nacelle Operational Health Monitoring Locations

ORIGINAL PAGE IS OF POOR QUA, ,Y, 31

VARIABLE

V 'A" THRU "F"

TT; PST

SENSOR (mSIG.

[6 SCANV-VALVE XDCRS

2 INDIV. XDCRS.

IICOND.

6 POWER + BAL. UNITS

.___2.

2 POWER + BAL.UNITS

CHANNEL NO.

&05H

6&7

DATA SYSTEM

HEWLETT PACKARD 3052A

I.ISCANNER 4. LINE DIGITAL PRINTER

VOLTMETER 5. TAPE 3. CALCULATOR PUNCH

t

TT"

TM 1-6; TC

E N ET EGT)

PLATINUMRTD

7 CU/CN THERMOCOUPLE

HROUPLE CHRMOIL E

LINEAR ID

150F

REFERENCE 9-15I

IOA

28V DC, 10A

POWER SUPPLY

~I

SCANIVALVE INTERFACE

BOEING 64-33083

SCANIVALVE, 6X48 PORTSENGINE MOUNTED

I

I

N1 &N2 RPM

2TACH GENERATORS

2 F. TOV.

CONVERTERS

j; PLA; TORQUE

SIGNALS FROM HAM. STD.

3 BUFFER AMPLIFIERS

19-21

8 DYNAMIC PRESSURES

18 KULITE XDCRS

8 SIGNAL CONDITIONERS

]22-2a

ca Figure 14. Block Diagram, Steady State Data Acquisiton System

J7 WIND TUNNEL PTO PST noDAFX CFZ NFY DISPLAY PANELSINLET FAN ENGINE W/T RMX PMZ YMY

A MODEL

ALPHA ENGINE OIL PRESS.7OPT I2OPT 47 PT PirT 52PS 4PS 9PS PSTDATA 3PDF 6iT 4PDC iTT_ FUEL PRESS.

SYSTEM I PDS N 1 ALPHA GIB OIL PRESS.

R GIB OIL FLOW

I'LA ENGINE OIL LEVEL ENGINE OIL TEMP.

I I I IDLER GEAR TEMP(31 TORQUE

EGT - M"7

N1 RADI US XYFNGBPOIEES()N

N2FAN BLADE STRESS 15) PT RAKE (1800) 4 PDC VIBRATION (6) ENG. CHIP 13)OO~~~ 3 PDFl N2Zfi TORO, Ip ENG CHI\0

1 PSD" FAN CHIP DETECTOR2 BLADE STRAINPTF-RMS Y, X-YY 1 RAKE STRAIN BLADE ANGLEP YBLADE ANGLE

;PDS-RMS 222X- POWER LEVER ANG a or PAV

N2 N2 PURGE

123 856 ~cRMY11L 34 DORIC DISPLAY RMS METERS . C S Y2 FAN FACE RAKE ENGINE CASE (4)

STRAIN GAGES ENGINE CASE (4) C1 or PAV TT2 ENGINE_______- 1 213145 61

G/B OILTEMP. - IN "ENG. CASE & COWL ff] TEMPFigure 15. Lift/Cruise Fan Inlet Test SUN GEAR GIB OIL TEMP. - OUT

Instrumentation System PROXIMETER 1112131415161 xY scoPE GB BRG TErMP VIBRATIONS

L/CFA 0-FAN' INLET TEST NASA-ARC aXSO WIND TUNNEL

RUN CONb TEST NO, VO ALPNA QPSF PTO -- PO TTO CONF DATE PSAR1, 1, 2593, 0,0 0.00 0.0 14,676 14,676 531.0 1. 91577, 14,660

WK1 ($ PTFA, ('Q),gS 527.F8A 1WKF, GPLAB @1WKE' @ PYCA DISC 107VN PTT & PSTA-TTT17'39 996 0246 1w7 307 15,8 15,2 1,013 0,M212 -0. 14,682 14,690 528.7SFWtR~iwxa C.61PTFAV D WF @ : @ _FTT, WKA19 2@1FPR 0WE ~ 3 KN1 EP 16N (3 1fT7wi 103,8 14,621 0,998 3p,2 - 15,0 9858, -1. -0,1 1.26518, 8,0

W3 18 We c~a.PTMAV @j YM 10TTC 24KN2 14aET PKN' 45,2 171.5 11,598 0. 529,9 4489,. 1, -0.1

STEADY STATE RAKE PRESSURES FAN FACE & Z ZLE COMPRESSOR FACE RK25.7 RK77,1 RK12B, RK180 RK231 RK282 RX334 RKt RK2 RKI Ri2 RK3 RK4 RX5 Ri6 RK7 RK8 Pc39 PC40 PC41 PC42 PC43 PC44 PC45 PM1 PM3 @ PSCI PSC 2 PSC 3 PSC4 PSC5 PSC6 PSC7 PSC8 PTFI PTF1I PTF21 PTF31 PTF41 PTF51 PTF6I PTM 1 PT1i1 (9 PTC1 PTC 7 PTC13 PTC19 PTC25 PTC31 PTC37 PTC43

2 12 22 32 42 52 62 2 12 2 8 14 20 26 32 38 44 3 13 23 33 43 53 63 3 13 3 0 16 21* 27 34 40 454 14 24 34 44 54 64 4 14 4 10 16 22 28 34 40 46 5 15 25 35 45 55 65 5 15 5 11 17 23*** 29 35 41 476 16 26 36 46 56 66 6 16 PTC6 PTCI2 PTCI8 PTC24**** PTC30 PTC36 PTC42 PTC48 7 17 27 37 47 57 67 7 17 8 18 28 38 48 38 68 8 18 9 19 29 39 49 59 69 9 19 @ TEMPERATURES@ 2DYNAMIC RMSPTF1O PTF20 PTF30 PTF40 PTF50 PTFGO PTF70 PTi1O PTM20 FAN DUCT FAN COMP

PPl PP2 PP3 PP4 PP5 PP6 RKI RK2 PDFI PDC1PP7 PM2 PM4 TTMI TTM4 PDF2 PDC2TTM2* TTM4 PDF3 PDC3TTI13 TTM6 PSP0

INLET COWL STATIC PRESSURES @COWL SURFACE MACH NUMERS PS PDC 1) PCI PC2 PC3 PC4 PC5 PC, Pc2 P 3 PC4 PC5 CORE INLET LIP PRESSURES6) 6 7 8 9 10 PC6 .......... same as static ...... PIC PpTIC

11) 11 12 13 14 15 pressure array 0.975 TIO1 16) 16 17 18 19 20 PTIC2 21) 21 22 23 24 25 PTI03 26) 26 27 28 29 30 31) 31 32 33 34 35 36) 36 37 38 39 40 4M) PC41 PC42 PC43 PC44 PC45

* Repl'aced with TT3 - Runs 19 & on CORE NOZZLE STATIC PRESSURES (PSIA) ** Replaced with PTC16 - Runs 23 & on

''PNIl PN2 PN3 PN4 *** Replaced with PTC17 - Runs 23 & on FORCE BALANCE **** Replaced with PTC18 - Runs 23 & on

LBS FT-LBS FX FY FZ FXY aY OX OZ

-549,0 50 28,0 549,0 1 -248.0 113,0 56,,0

FIGURE 16 - DATA SHEET FORMAT

Vo I,!1 44*+ PITCHING

1+ YAW + POET(OZ) MOMENTM(y)

349 -O-NTEhI*L

LIFT WFY)

+

DRAG (FX) I DRAG (FX)

ROL 1 MOMENT (Ox) .+.m ...T

LIFT (FY?)

ORIGINAL PAGE IS OF POOR QUALITY

O +... SIDE FORCE (FZ)

Figure 17. Nacelle Force Measurements -Nomenclature and Sign Convention

35

APPENDIX A

Q-FAN DEMONSTRATOR

LIFT FAN TECHNOLOGY BLADE STRESS REPORT

-36

Q-FAN DEMONSTRATOR

LIFT FAN TECHNOLOGY PROGRAM

BLADE STRESS REPORT

BLADE VIBRATORY STRESSES

FOR

Q-FAN DEMONSTRATOR

AT

NASA AMES 40' x 80' WIND TUNNEL

Purchase Contract No. N-947295-9578 "NASA Prime Contract 2-9640

Prepared by Project Manager

Reviewed by:

Prepared for:

____ ___o Project Engineer D. P. Currie

L 4 Head of Applied Mechanics Dr. Robert W. Cornell

Program Manager DFJ. Nelson

Boeing Aerospace Company, Seattle, Washington 37

SUMMARY

The Hamilton Standard.QFT-55 full'scale Q-Fan Demonstrator propulsor was utilized by Boeing Aerospace Company to fabricate a large scale variable pitch Lift/Cruise fan nacelle. The unit was tested in 1976 at Boeing's Tulalip, Washington static test facility and the NASA Ames 40'.x 80' wind tunnel under a NASA Prime Contract NAS2-9215. A Blade Stress Report covering those tests was submitted to Boeing on 16 November 1976.

This report covers the more recent tests which were conducted in the NASA Ames 40' x 80' Wind Tunnel during September-October 1977 under the NASA Prime Contract NAS2-9640. These tests included operation at higher nacelle tilt angles and free stream velocities than those tested in 1976.

38

I

TEST DESCRIPTION

In connection with the Boeing-NASA V/STOL aircraft program, blade vibratory stresses were measured during the testing of the Hamilton Standard Q-Fan Demonstrator in the NASA Ames 40' x 80' wind tunnel, This document reports on theresults of these measurements,

The rotating components of the test unit were identical to those described inReport NAS CR-121265 (HSER 6163) which covered the stresses asmeasured at Hamilton Standard. Referring to Figure 6 (Pg, 239Q of that report or Figure I included in the appendix of this report, stresses were measured using the bending strain gages located at 362 mm (14.25 in.) from the blade tip on Blades No. 1, 6 and 7 and the vee strain gages located at 89 mm (3.5 in.) from the blade tip on blades 2 and 9 during this test program. These stress measurements were recorded on magnetic tape and played back onto Sanborn records. Also recorded and played back were torque, fan speed and fan blade angle. Wind tunnel air speed, fan angle of attack (actually yaw inflow angle) and the power lever angle (PLA) were noted on the test log. An X-Y plotter was also used to record Blade No. I bending stress vs. N2 (engine rpm). A number of the recordings were played back to determine the dominant frequencies of the vibratory stressing at this gage location.

The tests were conduct6d on the 15th thru the 21st Sept. 1977 and the 13th thru 15th October 1977. The interruption in testing was caused by an unscheduled shutdown of the wind tunnel.

Tunnel wind speeds set during the tests were 0, 40, 60, 75, 90, 105, 125, 140 and 160 khots. Inflow angles (yaw angles) were varied from 00 to 1200. The maximum fan speed and fan blade angle were 3250 RPM and 560 respectively.

OjfqNAL PAGE IS

OB-DQOR QUALITY

39

TEST RESULTS

Bending stresses were generally acceptable using a continuous limit of + 38 MN/m 2 (+ 5500 psi) observed for these tests. This level was slightly exceeded when operating at combinations of angle of attack, tunnel: speed and torque that produced flow separation at the duct inlet. This was evident inTest Points 27 and 31.

Bending stresses also became significant when operating near a first flatwise (If)mode critical speed; that is,where the aerodynamic excitation at a multiple (nF) of the fan speed is near the first flatwise natural bending frequency of the blade. A critical speed diagram is shown in Figure 2. This diagram is also shown on Figure 5-39 of Report NASA CR-121265 (HSER 6163). This type of stressing was particularly evident in Test Points 30 with a maximum duct angle of attack of 1200. For instance, Test Point 30.2 showed moderate 2F/lf response at 3250 RPM and Test Point 30.7 showed significant 3F/If response at 2300 RPM. This stress behavior was also evident in previous running of this q-Fan.

Maximum shear stressing measured was 10.0 MN/m 2 (+1450 psi) occurring during operation at maximum power with a maximum duct angle of attack of 1200. This amplitude of shear stressing isacceptable.

Appended to this report are copies of the on-line X-Y plots relating Blade No. 1 total vibratory stress (+ psi) at 362 mm (14.25 in.) from the tip and the power.turbine speed, N2. -To obtain.the fanspeed, the N2 speed is divided by the 4.75 gear ratio. --MarkHed-ori these plots-are the freestream tunnel velocity,"'V; duct.angle df attack .(actuallyoayaw angle in the tunnel); fan blade angle,1., at the 75%blde -radiusstation,-date of the test, and the Run Number'(same as Test Point Number) designated by Boeing. The sequence of the testing, which at times influenced the stressing, was in the order of increasing Run Number.

Twenty-four-such plots.are givens-- (Not all of the runs are shown on these plots).- The430 blade anglepoints for -Boeing.RuniNos--.13 and 25 show moderate stressing; Ltkewise, stressing-is moderate 'for-Rin No. 30. All of -theseconditions-arenat dcmbinations~of-high duct angle of attacki tunnel speeds-of 39 mh/s (75 knots-).oless -and comparatively 19w torque. In addition5 the.stress peaks generally-dccur near the N2' speed of 10,000 RPM, corresponding to a fan'speed bf about-'2100 RPM..- Inlet flow separation'is believ6d to have occurred duringthe deceleration'for Boeing Run No. 13.

Also givenin ihe-App6ddix'Aresix-pages summarizing the stresses measured during-the-testing---These-: tables "were prepared from the Sanborn playbacks.

-Idetifca-tion'of-the'-test:cond-iti6n-is diven'by the HS Run'No. correlated' with:the-Boeing Test Point.No.- Maximum bending stresses were noted during Test:Points-Nos.27, 30, and 31.2As above, the points for Nos. 27 and 30 were at combinations of high 'angle of attack,-moderate tunnel speeds, and low-torquecTest Points for No.'31 were unique in that this is the only run-in which, the variable.xas- the:duct-angle of attack. Tunnel speed and' fan' power-were kept-constant.= Upon' increasing angle-of attack, significantstressing:occurredonly-as the maximum angle of 800 was,approached. This s,tressngwas:mainta-ined until: af angle of less than 78P was reached?. This hysteres's,may be a characteristic of duct inlet.'separation.

http:Test:Points-Nos.27http:Point.Nohttp:Boeing.RuniNos--.13

Test Results (continued)

The Appendix also includes 15 frequency spectrum plots for conditions selected from the stress tabulations. These results confirm that the major bending stress response is at frequencies which are integer multiples of the fan speed - IF,2F, 3F, etc. Stressing can become significant wben one of these frequencies is near the blade natural frequency just above 100 Hz. Test Point 30.7, for instance, shows predominant response near 112 Hz. This is the 3F/lf condition, Likewise, Test Points 30.l and 30.2 show moderate stressing near 105 Hz. This isthe 2F/lf condition.

These frequency spectrum plots also show stress response in the first flatwise natural befding-mode when thi-frequency is-not an integer multiple of the fan speed. This response was evident, for instance, for Boeing Test Points 27.5, 31.4, and 31.5. Itwas also evident in Points 27.6 and 27.8 where the response was very close to but not exactly at the 3F frequency. This type of response could very well be caused by turbulence associated with duct inlet separation.

Insummary, the vibratory stress levels measured in this investigation were acceptable. No significant structural fatigue damage accrued from these tests. Major stress response was due to the proximity of a frequency of aerodynamic excitation to the first natural blade bending frequency. This type of response was evident in previous tests of this Q-Fan. Some stress response was noted in the4f.irst-natural bending -nodeat a frequency-not equal to an integer multiple-of the fan -speed.-e This tesponse ould- be expected from the turbulent nature of the air flow associated mith separation withinthe'duct.

ORIGINAL PAGE IS

OF POOR QUALITY

41

APPENDIX

TO

Q-FAN DEMONSTRATOR

LIFT FAN TECHNOLOGY PROGRAM

BLADE STRESS REPORT

42

CORE.

'V

I.

t -89MM F.T.

I- - 27MM F.T.

362MM F.T.

387MM F.T.

FT. - FROM TIP

F oIGRE1

*43FIGURE 1 43

A-RP ROXIMATE .FREQUENCY OFFIRST fLATWISEBLADE 'MODE INrFORWARD THRUST

- ,CALCULATED

S..--- MEASJURED

- :200

A 'VANE..OgDER-.

7F

_ _ _ ._ 6F/

N2 + IF ,(ENGINE UNBALANCE ORDER)

-5F 4F

__-_ _ 4'...

u

Z35"

2'2F

50

0 1500 -- 2000 2500 "FAN SPEED-RPM . "

3000 3500

1.4 M DIAM. Q-FAN DEMONSTRATOR CRITICAL SPEED DIAGRAM

FIGURE Z

4t'

-- --

____

0

20 X 20 PR INCH ACW

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