EVALUA'T~ON OF AIRFLOW RING COMBUSTORS FROM COMPRESSORS · PDF fileevalua't~on of-circum ferental airflow uniformity ente ring combustors from compressors volume i dscussmon of data

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N73-14276

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EVALUA'T~ON OF-CIRCUM FERENTAL AIRFLOW

UNIFORMITY ENTE RING COMBUSTORSFROM COMPRESSORS

VOLUME I DSCUSSMON OF DATA AND RESULTS

Novmbier 1972

by J. H. Shadowen and W. J. Egan, Jr.

PRATT & WHITNEY AIRCRAFTFLORIDA RESEARCH AND DEVELOPMEN

Prepar d For NATIONAL AERONAUlHDCS AND SPACE ADM

NASA Lewis Research Center

Contracd NA$3-15693

NASA CR-121009

2NASA-CR-121009) EVALUATION OF

CIRCUM.FERENTIAL AIRFLOW UNIFORMITYENTERING COMBUSTORS FROM COMPRESSORS°

o.Ho Shadowen, et al (Pratt and WhitneyAircraft) Nov. 1972 122 p CSCL 20D

https://ntrs.nasa.gov/search.jsp?R=19730005549 2018-05-15T09:09:26+00:00Z

IR1eot.2. Governmenport Accession No. 3. Recipient's Catalog No.

NASA CR-121009

4. Title and Subtitle 5. Report Date

EVALUATION OF CIRCUMFERENTIAL AIRFLOW UNIFORMITY November 1972

ENTERING COMBUSTORS FROM COMPRESSORS, VOLUME I - 6. Performing Organization Code6. Performing Organization CodeDISCUSSION OF DATA AND RESULTS

7. Author(s) 8. Performing Organization Report No.

J. H. Shadowen and W. J. Egan, Jr. PWA FR-5183

10. Work Unit No.

9. Performing Organization Name and AddressPratt & Whitney AircraftDivision of United Aircraft Corporation 11. Contract or Grant No.

Florida Research and Development Center NAS3-15693West Palm Beach, Florida 33402

13. Type of Report and Period Covered

12. Sponsoring Agency Name and Address Contractor Report

National Aeronautics and Space AdministrationWashington, D. C. 20546 14. Sponsoring Agency Code

15. Supplementary Notes

Project Manager, Everett E. Bailey, Fluid Systems Components Division, NASA Lewis ResearchCenter, Cleveland, Ohio 44135

16. Abstract

A study of the compressor discharge airflow uniformity of two compressors from advanced engines, the

J58 and F100/F401, was made. Compressor discharge pressures and temperatures at up to 33 circum-

ferential rake locations allowed the airflow distribution to be ascertained and computer plotted. Several

flight conditions and compressor variables, i.e., inlet distortion, modified seals, etc., were analyzed.An unexpectedly high nonuniform airflow was found for both compressors. Circumferential airflowdeviation differences of up to 52% from maximum to minimum were found for the J58, and up to 40% for

the F100/F401. The effects of aerodynamic and thermal distortion were found to be additive. The datawere analyzed for influence of exit guide vane wakes and found free of any effect. Data system errors

were small in relation to the measured pressure and temperature variations.

17. Key Words (Suggested by Author(s)) 18. Distribution StatementCompressor Temperature Pattern Unclassified- unlimitedCombustor Pressure PatternDiffuserAirflow UniformityAirflow Deviation

19. Security Classif. (of this report) 20. Security Classif. (of this page) 21. No. of Pages 22. Price

Unclassified Unclassified 122 $3.00

* For sale by the National Technical Information Service, Springfield, Virginia 22151

INASA-C-168 (Rev. 6-71)

Preceding page blankC ONTE NTS-

PAGE

LIST OF ILLUSTRATIONS ....... .................... iv

LIST OF TABLES ................................ viii

SUMMARY ..................................... 1

INTRODUCTION ................................. 1

SCOPE OF INVESTIGATION ......................... 2

DESCRIPTION OF RIGS AND INSTRUMENTATION .......... -3

J58 Compressor Rig and Instrumentation ............. 3F100/F401 High Compressor Rig and Instrumentation .... 4

DATA ANALYSIS ................................ 6

RESULTS ...................................... 9

J58 Compressor Rig ........................... 9F100/F401 High Compressor Rig .................. 12

DISCUSSION OF RESULTS .......................... 14

SUMMARY OF RESULTS ............................ 19

LIST OF SYMBOLS ..... ......................... 20

REFERENCES .................................. 115

iii

ILLUSTRATIONS

FIGURE PAGE

1 J58 Compressor Rig Schematic .................. 29

2 F100/F401 High Compressor Rig Schematic .......... 31

3 Typical J58 Compressor Discharge Instru- mentation .... .... . . . ........... 33

4 Typical Schematic of J58 Compressor DischargeInstrumentation, Looking Upstream ............... 34

5 Schematic of J58 Compressor Discharge Instru-mentation, Looking Upstream; Rig 30239,Build 206, Runs 136 and 182 .................... 35

6 J58 Cruise Inlet Distortion Screen, LookingUpstream ................................ 36

7 J58 Compressor Rig Inlet Total Pressure Distribution,Looking Upstream; Cruise-Distorted Inlet, Rig 30204,Build 10, Run 569 ........................... 37

8 Typical F100/F401 High Compressor DischargeInstrumentation ............................ 38

9 Schematic of F100/F401 Compressor DischargeInstrumentation, Looking Upstream ............... 39

10 J58 Circumferential Airflow Deviation;30204-10-0531 ............................. 40

11 J58 Compressor Discharge Flow Map; 30204-10-0531 . .. 41

12 J58 Circumferential Airflow Deviation; 30204-10-0552 . .. 42

13 J58 Compressor Discharge Flow Map; 30204-10-0552 . . . 43

14 J58 Circumferential Airflow Deviation; 30204-10-0569 . . . 44



17 J58 Compressor Discharge Flow Map; 30204-11-0185 .. . 47

18 J58 Circumferential Airflow Deviation; 30204-11-0171 ... 48

19 J58 Compressor Discharge Flow Map; 30204-11-0171 . .. 49




23 J58 Compressor Discharge Flow Map; 30204-12-0453 .. . 53

24 J58 Circumferential Airflow Deviation; 30239-202-0008 . . 54

25 J58 Compressor Discharge Flow Map; 30239-202-0008 . . 55

iv

ILLUSTRATIONS (Continued)

FIGURE PAGE

26 J58 Circumferential Airflow Deviation; 30239-203-0014 . . 56

27 J58 Compressor Discharge Flow Map; 30239-203-0014... 57


29 J58 Compressor Discharge Flow Map; 30239-204-0037 .... 59

30 J58 Circumferential Airflow Deviation; 30239-206-136 .... 60


32 J58 Circumferential Airflow Deviation; 30239-206-182 .... 62


34 F100/F401 Circumferential Airflow Deviation;34017-204-8031 .............................. 64

35 F100/F401 Compressor Discharge Flow Map;34017-204-8031 ...... . ....................... 65


37 F100/F401 Compressor Discharge Flow Map;34017-204-8023 .............................. 67


39 F100/F401 Compressor Discharge Flow Map;34017-204-8014 ............................... 69

40 F100/F401 High Compressor Inlet Pressure Map,No Distortion; Build 205, Run 4313 ................. 70

41 F100/F401 High Compressor Inlet Temperature Map,No Distortion; Build 205, Run 4313 .................. 71







v


FIGURE PAGE

48 F100/F401 High Compressor Inlet Pressure MapWith Aerodynamic Distortion; Build 205, Run 4233 ....... 78


50 F100/F401 Compressor Discharge Flow Map;34017-205-4233 .. ............................ 80





55 F100/F401 High Compressor Inlet Pressure Map WithAerodynamic and Thermal Distortion, AT = 22°K (40°F);Build 205, Run 4453 ........................... 85

56 F100/F401 High Compressor Inlet Temperature MapWith Aerodynamic and Thermal Distortion, AT = 22 ° K(40 0F); Build 205, Run 4453 ...................... 86







63 F100/F401 High Compressor Inlet Pressure Map WithAerodynamic and Thermal Distortion, AT = 55 ° K(100°F); Build 205, Run 5313 ..................... 93

64 F100/F401 High Compressor Inlet Temperature Map WithAerodynamic and Thermal Distortion, AT = 55°K (100°F);Build 205, Run 5313 ........................... 94


vi


FIGURE PAGE





70 F100/F401 Compressor Discharge Flow Map;34017-205-4513 ............................... 100

71 F100/F401 High Compressor Inlet Pressure Map WithThermal Distortion, AT = 22°K (40°F); Build 205Run 4653 ................................... 101

72 F100/F401 High Compressor Inlet Temperature Map WithThermal Distortion, AT = 22°K (40°F); Build 205,Run 4653 .......... . ............ .......... 102

73 F100/F401 Circumferential Airflow Deviation;34017-205-4653 ............................. 103

74 F100/F401 Compressor Discharge Flow Map;Y 34017-205-4653 .............................. 104





79 F100/F401 High Compressor Inlet Pressure Map WithThermal Distortion, AT = 55°K (100°F); Build 205,Run 4753 .................................. 109

80 F100/F401 High Compressor Inlet Temperature MapWith Thermal Distortion, AT = 55 °K (100 °F);Build 205, Run 4753 ........................... 110

81 F100/F401 Circumferential Airflow Deviation;34017-205-4753 .......................... .. 111


83 J58 Compressor Discharge Total Pressure DistributionRelative to Exit Guide Vane (EGV) Trailing Edge ........ 113

vii


FIGURE PAGE

84 F100/F401 High Compressor Discharge Total PressureDistribution Relative to Exit Guide Vane (EGV) TrailingEdge ..................................... 114

TABLES

TABLE PAGE

I Test Conditions and Variables for CompressorsSurveyed ............................. 22

II J58 Compressor Rig Discharge InstrumentationCombinations .............................. 23

III J58 Compressor Rig Data Comparison ............. 25

IV F100/F401 High Compressor Rig Data Comparison ..... 27

viii

SUMMARY

A study of the airflow uniformity leaving compressors and entering com-bustors was made using compressors from two advanced engines, the J58 andthe F100/F401. Compressor discharge total pressures, total temperatures, andstatic pressures were obtained from rig tests at up to 33 circumferential rakelocations allowing the calculation of local airflow ratio and circumferentialairflow deviation. The data analyzed encompassed test conditions of sea leveltakeoff, cruise, and combat, and compressor variables such as aerodynamic andthermal distortion, modified shrouds and seals, and indexed stators. The resultsare presented as computer-generated plots.

For all cases analyzed, the J58 compressor showed highly nonuniform air-flow leaving the compressor. Local airflow values ranged from 52% above to88% below average. In the circumferential direction airflow values ranged from20% above to 32% below average. At least four areas of high airflow on the mid-span are noticeable in each case. The data were analyzed for influence of exitguide vane (EGV) wakes and found free of any effect.

The F100/F401 high compressor showed more uniform airflow than did theJ58 compressor but it still has local airflow values ranging from 35% above to 52%below average. In the circumferential direction, airflow values ranged from 20%above to 21% below average. The areas of high airflow are concentrated on thecompressor hub. Again the data were analyzed for influence of exit guide vane(EGV) wakes and found free of any effect. The effects of aerodynamic and thermaldistortion were found to be additive.

The information presented shows that an unexpected compressor dischargeairflow nonuniformity problem exists in state-of-the-art compressors. Methodsof distributing the airflow more evenly both in the compressor and in the diffuser-combustor system should be investigated.

INTRODUCTION

One of the significant problems in the development of turbojet combustionsystems has been that of obtaining the uniform turbine inlet temperature patternsnecessary for long turbine life. During engine development programs it hasoften been observed that large-scale circumferential turbine inlet temperaturevariations are insensitive to changes in combustor geometry, giving rise to thetheory that these circumferential temperature nonuniformities are caused bycompressor discharge airflow nonuniformities. Compressor discharge instru-mentation for past engine development programs, as contrasted with researchrequirements, has been inadequate to prove or disprove the presence of non-uniform airflow. Airflow has therefore usually been assumed to be uniformbecause of the apparent mechanical uniformity of the compressor.

The objective of this program was to determine if, and to what extent, air-flow nonuniformities arepresent in the discharge of advanced turbine enginecompressors. For this purpose, compressors from two engines were chosenfor investigation: the J58 and the F100/F401. The J58 is a high-altitude supersoniccruise, turbojet engine, whereas the F100/F401 is an advanced turbofan engine forthe F-15/F-14B air superiority fighter aircraft.

I

Data from several J58 compressor rig tests with extensive instrumentationwere available. For the F100/F401 high compressor, the necessary com-pressor discharge instrumentation was designed and fabricated under this pro-gram, and data were gathered during the normal course of compressor rigtesting. Data from both compressors were analyzed under this program andthe results are presented herein. Test variables included sea level takeoff(SLTO), cruise and combat operating conditions, indexing of stators, modifi-cations to seals and shrouds, and inlet total pressure and temperature distor-tion. The inlet total pressure and temperature distortions will hereafter betermed aerodynamic and thermal distortion, respectively.

The data were obtained from up to 33 total pressure rakes and up to 12 totaltemperature rakes installed in the compressor discharge. Each rake had five orseven heads spaced across the annulus. Static pressures in the same plane at upto ten locations on each wall were also obtained.

This report discusses the acquisition, analysis technique, and results ofthe data obtained from the J58 compressor rig and the F100/F401 high compressorrig. It presents the results in nondimensionalized plot form for ease of com-parison between test conditions and variables.

A companion report, NASA CR-121010 "Evaluation of CircumferentialAirflow Uniformity Entering Combustors From Compressors, Volume II -Data Supplement" (Reference 1), contains the input data for each test case inboth tabulated and computer-plotted form. It is designed for the use of thosedesirous of implementing the results of this report, or in further exploration ofthe subject of compressor discharge uniformity.

SCOPE OF INVESTIGATION

To investigate the possibility of nonuniform airflow leaving the compressorand entering the diffuser-combustor system of advanced turbine engines, twocompressors were chosen for evaluation. The J58 turbojet is a large, supersoniccruise engine. The compressor has nine stages, a compressor pressure ratio of8.75:1, and is shown schematically in figure 1. The other compressor chosenwas the F100/F401 high compressor. This engine is an advanced turbofan and iscurrently in the development stage. It will power the F-15/F-14B air superiorityfighter series. The high compressor of this engine has ten stages, a compressorpressure ratio of 8:1, and is shown schematically in figure 2.

These compressors were analyzed for several variables at sea level takeoff(SLTO), cruise, and combat conditions. The variables included aerodynamic and/or thermal distortion, indexing of stators, and modified seals and shrouds. Notall conditions and variables were studied for both compressors, however. Thedata were obtained from seven builds of two J58 compressor rigs and two buildsof an F100/F401 high compressor rig. Run conditions for each compressor aregiven in table I, which shows the wide variation for the 31 test points consideredin this study.

The data obtained for these test points included total pressures and tem-peratures and static pressures at the compressor discharge plane. These dataallowed the radial and circumferential total pressure and temperature and the

2

circumferential static pressure distributions to be obtained. From these profiles,the average circumferential airflow deviation and the local airflow deviation werefound for each test point. Inlet distortion patterns were obtained from total pres-sure and total temperature rakes installed in the compressor inlet.

DESCRIPTION OF RIGS AND INSTRUMENTATION

J58 Compressor Rig and Instrumentation

The J58 compressor is used in a high altitude, high flight Mach numberengine that is considered state-of-the-art in turbojet engines. This compressoris a single-spool, variable inlet guide vane, fixed stator design with an interstagebleed and has a design pressure ratio of 8.75:1. The compressor is shownschematically in figure 1 along with a table describing each stage.

The data used in this program to determine the J58 compressor dischargeairflow uniformity were all generated in the course of the engine developmentprogram and were made available to this program. During compressor develop-ment, two compressor rigs were instrumented at the compressor dischargeplane with up to 33 seven-point radial total pressure rakes, up to 8 static pres-sure taps on the diffuser annulus outer wall, and up to 11 total temperatureprobes with one or two half-shielded temperature elements. Typical instru-mentation is shown in figure 3. A schematic view of the circumferential loca-tions of the instrumentation is shown in figure 4, while the axial location isshown in figure 1. For one build, 206, the one- and two-point total temperatureelements were replaced with five-point radial temperature rakes and eightadditional static pressures were installed on the diffuser annulus ID. This con-figuration is shown schematically in figure 5. The different builds used variouscombinations of static pressures, total pressures, and total temperatures.Table II describes the various combinations of instrumentation used.

The compressor rig was tested in the Florida Research and DevelopmentCenter C-3 altitude test facility. This facility has an atmospheric intake upstreamof a flow-controlling butterfly valve. The air is fed to a cylindrical plenum ap-proximately 6. 096 m (20 ft) dia x 6. 096,m (20 ft) long. Flow straightening in theplenum entrance is accomplished by a means of baffles and flow straighteningscreens. A standard bellmouth inlet is installed in the plenum just forward ofthe compressor rig with enough room for inlet distortion screens between thecompressor and the bellmouth, if necessary. Interstage compressor bleed isextracted between the fourth and fifth stages through an annular duct around thecompressor approximately 61 cm (24 in.) long, which feeds six collector mani-folds. These six manifolds each feed 12.7 cm (5 in.) diameter pipes, which arelocated at 60 deg, 90 deg, 120 deg, 240 deg, 270 deg, and 300 deg from TDC,respectively. The compressor rig is tested at a reduced inlet pressure but atproper corrected flow and rotor speed. Since the bleeds are usually subatmos-pheric, they are ducted back to the stand inlet flow duct between the controllingorifice and the plenum, upstream of the flow straightening baffles. The inter-stage bleed flows were run per normal operating schedule except for one test,during which the bleeds were not opened. Provisions for customer bleed arelocated on the inner combustor case downstream of the diffuser struts anddiffuser dump section (figure 1). However, the customer bleed was not con-sidered to have any significance in this program because of its distance fromthe compressor discharge instrumentation plane. The airflow undergoes dif-fusion and a sudden expansion between the compressor discharge and

3

customer bleed ports. The exhaust from the rig is collected in a manifold anddumped to atmosphere, or put through an exhauster. The compressor is coupledto a steam turbine that drives the compressor clockwise, looking upstream.

All pressures were recorded using a scanivalve and pressure transducersystem. The scanivalves were connected to AP transducers with a range of10.34 N/cm2 (15 psid). All transducers were referenced to the same pressure,which was accurately determined. An error analysis on the C-3 stand pressuresindicated a measurement uncertainty within +0. 0758 N/cm2 (0.11 psi). Temper-ature data were taken using half-shielded chromel/alumel (C/A) thermocouples,except in Build 206, which used 5-point C/A thermocouple rakes with Kiel-typeheads. Both measured the temperature with an uncertainty less than ±1. 111°K(2°R). The output of the transducers and thermocouples was fed into an auto-matic data recording system, which supplied reduced data in engineering units.

Several variables were studied in the J58 compressor rig at two importantconditions: SLTO and cruise. Among those variables were aerodynamic inletdistortion, indexing of stators, modified shrouds, and modified seals. Table Ishows the comparisons to be presented in this study by run number. The inletdistortion was produced by an inlet screen (figure 6), which was used at thecruise condition to simulate inlet total pressure distortion seen at that flightcondition. The distortion typically generated by the screen is shown in figure 7as an inlet pressure distribution map. The indexing of stators consisted ofrotating the first four stages 36 deg counterclockwise (ccw), looking upstream.The modification of interstage seals consisted of replacing, depending on thebuild, the sheet metal or honeycomb seals with Feltmetal or filled honeycomb,abradable-type seals and reducing the seal clearances. The modification to theshrouds consisted of replacing the sheet metal with Feltmetal in stages 5through 9.

F100/F401 High Compressor Rig and Instrumentation

The F100/F401 high compressor is used in the advanced state-of-the-artturbofan engine that will power the F-15/F-14B air superiority fighter aircraft.The high compressor is a single-spool, variable stator design with a nominalpressure ratio of 8:1. The compressor is shown schematically in figure 2 alongwith a table describing each stage.

To analyze the compressor discharge airflow of the F100/F401, it was firstnecessary to obtain data with a sufficient number of total pressure and total tem-perature probes at the compressor discharge plane. As the F100/F401 is in thedevelopment phase, a high compressor rig was being tested while this programwas being worked on. Through cooperation of the Air Force and Navy, per-mission was obtained to install the necessary instrumentation and gather dataon a noninterference basis. The instrumentation was designed and fabricatedunder this program and consisted of total pressure rakes with five Kiel-typeheads located on centers of equal area, total temperature rakes with five Kiel-type heads shielding chromel/alumel thermocouples, again on centers of equalarea, and finally, static pressure taps for the annulus inner and outer walls.Typical probes are shown in figure 8. Twenty-two total pressure rakes (fourof which were regular rig instrumentation), ten total temperature rakes (againfour of which were regular rig instrumentation), ten OD static pressure taps,and ten ID static pressure taps were used as shown in figure 9 to obtain the

4

circumferential and radial pressure and temperature profiles necessary tocalculate airflow deviation.

The compressor tested during this program is the version used in develop-ment flight engines and is typical of the production engine high compressor. Thecompressor rig was tested at a reduced pressure, but at proper corrected flowand rotor speed. The compressor rotates clockwise, looking upstream. Thestators are variable in stages 4 through 6, and are adjustable in the remainingstages. The variable stators were operated per normal schedule with the ad-justable stators in an optimized configuration for the rig build. The rig wastested in the East Hartford test stand X-27. This facility is essentially thesame type of stand as that used to test the J58 compressor rig. The facility hasan atmospheric inlet upstream of a flow-controlling butterfly valve. The air isfed to a cylindrical plenum, which is approximately 2. 134 m (7 ft) dia x 3.048 m(10 ft) long. Although this plenum is considerably smaller than the FRDC C-3stand plenum, the ratio of rig inlet flowpath diameter to plenum diameter foreach rig is similar. Flow straightening is accomplished by baffles and screensin the forward section of the plenum. A standard bellmouth inlet is installed inthe plenum just forward of the compressor rig with space for inlet distortionscreens downstream of the bellmouth. Provisions for extracting customer bleedthrough the trailing edge of the blunt-ended diffuser struts are included in the rig.However, the customer bleed was not considered significant because of its distantlocation from the compressor discharge instrurmentation plane. The trailing edgeof the diffuser struts, where the bleed ports are located is in the plane of the dif-fuser dump section so the airflow is diffused and subjected to a sudden expansionbetween the compressor discharge and the customer bleed location. The exhaustfrom the rig is collected in a doughnut shaped manifold downstream of the dif-fuser case and exhausted to ambient or through an exhauster. The compressoris coupled to a turboshaft engine through a gearbox which rotates the rig clock-wise looking upstream.

All pressures were recorded using a scanivalve and pressure transducersystem. The pressures were read as absolute values and every other port wasreferenced to vacuum. The vacuum readings were checked to ensure agreementbetween transducers. The uncertainty is less than ±0. 069 N/cm2 (0. 1 psi) onabsolute pressures. The total temperatures were taken by five-point temper-ature rakes with chromel/alumel thermocouples with an uncertainty of less than:1. 111°K (2°F). The output of the transducers and thermocouples was fed intoan automatic data recording system, which supplied reduced data in engineeringunits.

During all high compressor rig tests, a total pressure distortion screenwas present over the outer 30% of the inlet annulus. Distortion generated bythis screen simulated the fan pressure profile entering the high compressor.The screen mesh used was 3 x 3 x 0.0248 cm (0. 063 in.). An additional totalpressure distortion screen, 4 x 4 x 0. 0248 cm (0. 063 in.) mesh, was locatedover the compressor inlet from 165 deg to 345 deg clockwise, looking upstream,when desired, to provide circumferential aerodynamic distortion. Thermaldistortion is accomplished by splitting the inlet duct in half and heating half theair by means of a heater burner system. The same 180 deg sector covered byscreens for aerodynamic distortion was heated for thermal distortion. Theheater burner system allowed for AT's of 22°K (40°F) and 55°K (100°F) byblending of ambient air downstream of the heater burners to reach the desired AT.

5

DATA ANALYSIS

To analyze the information obtained at the compressor discharge planeof both the J58 and F100/F401 compressor rigs, a data reduction program wasformulated that allowed for computer-generated plots of the input total tem-perature and pressure profiles as well as the local and circumferential airflowdeviation for each test case. The latter sets of computer plots show conclu-sively the areas of high and low airflow at the compressor discharge that arebeing fed into the diffuser-combustor system. These plots allow for comparisonsbetween variables in different builds to indicate their effect on airflow deviation.

The input necessary for this program consists of total pressure profiles,Pti,i, total temperature profiles or points, Tti , and static pressures, Psi k,at various circumferential locations. The profi'e descriptions can have up toseven values per circumferential location. The diffuser annulus radii to theinner and outer walls are required at the plane of the static pressures and theplane of the total pressures. The average static pressure at the plane of thetotal pressures is found by adjusting the static pressure for the area change to theplane of the total pressure if the static is not in the exact plane of the total pres-sure. The span location, i.e., probe head location between inner and outer wallsof the annulus, is also necessary for the total pressures and temperatures. Fromthese input items the data reduction program can calculate and generate plotsat the compressor discharge of average circumferential airflow deviation,average square root of dynamic pressure ratio, and local airflow deviation,usually referred to as the local airflow map.

The engineering formulation of the data reduction program is presentedbelow. All symbols not defined in the text are in the list of symbols. Severalintegers and subscripts are used in this formulation and are defined here forclarity:

i is circumferential location

j is radial location or associated with a radial location

k is indicator for ID or OD

L is indicator, = 1 for J58, = 2 for F100/F401

m is number of total pressure rakes

ml is number of total temperature rakes

n is number of static pressure circumferential positions.

The compressor discharge area at the plane of the total pressure rakesis:

At

= ?r(R 2 ) 2t~OD t,L ID t,L

and the area at the plane of the static pressures is:

As= r (ROD 2. RID 2

s,L s,L

6

The average static pressure at an input circumferential location is:2

P =

i, avgk1p

si,k

while the average total temperature, at an input circumferential location is:nn

ti, avg = nnj=1

Tti, j + 459.7

where when L = 1, nn can be = 1, 2, or 5when L 2, nn = 5.

Note that, if any values to be used in the summations are zero, they are elimi-nated from the average. These quantities are now used with an interpolationscheme so that appropriate values at the total pressure circumferential locationscan be obtained.

The overall average total temperature and static pressure are calculated by:

mlagoTt m l iaavg, ov ml i=l i, avg

and PSavg ovavg, ov

The input values of total pressure are scanned and any values below localaverage static pressure are discarded. The average total pressure for each rakeis then computed as:

nmtiavgPti, avg J= A j Pti, j)/At

where

nm= 7 when L = 1

nm =.5 when L= 2.

The average static pressure at the plane and location of the total pressurerakes is found by interpolation at the appropriate circumferential location, andthen adjusting that value for the area change to the plane of the total pressuresto obtain:

(A )2

Psi,avg =] iavg A I Pti,avg - P'i,avg (8 i' Pti,'avg)}

It is recognized that this expression is for incompressible flow but the Machnumber is low and the area ratio is small so that the discrepancy is negligible.

To obtain the airflow parameters it is first necessary to evaluate the Machnumber at each probe head and for each rake. This is done in an iteration loop

7

iPs i, avg

1n

n i=l

that initially guesses the static temperature to evaluate )' i or Y i, j (Reference 2),then solves for

2Mij = (- 1

1_,J

[(Pti, j/PSi, avg

(Pti, avg/P Si, avg)

,) 1]

])1/2'Y-1 ])1/2

) - 1 ]

'Y = Yiiq 3

Y =Y.I 'Y = 'Y 1

The static temperature is then calculated from the isentropic relation

T,si, j ,= T avg-1 M2

iT S=Tt. avg2 i I i, avg

and used as the guess in the next iteration. This iteration loop is closed, andthe correct values obtained, when the static temperature is within one degreeof the guessed value. This is less than 0. 2% in most cases.o

The airflow parameters and final averages are obtained by:

Local airflow per unit area is:

/i1/2(W/A) i j = ik)

i, avg

Average airflow per unit area is:

(W/A) i = T 1/2

1

\ Y+1+- 1M 2 2('Y-1) M P Y Y 21 _+ i j) Mi, j Pt..' SY = e

'Y +12(Y -1)

Mi Pt 'i1, avg

The uncertainties involved in the local airflow per unit area and the averageairflow per unit area were calculated from the measurement inaccuracies of thepressures and temperatures. For the J58 compressor, the uncertainty in localairflow per unit area was 1, 6% and the uncertainty in average airflow per unitarea was 2. 5%. For the F100/F401 compressor, these uncertainties were 1. 6%and 1.4%, respectively.

Overall average airflow per unit area is:

(W/A)avg, ovm

= - E (W/A)m i=1

Local airflow ratio is:

Wi, j/Wavg,ov = (W/A)i, j Aj/(W/A)avg, ov A

8

Y= 'Y .ii

1

Percent airflow deviation is:

Deviation = [(W/A)i/(W/A)avg ov - . x100

Average total pressure is:

m -IPt Pt (W/A)i / (W/A)

avg, ov Lt i, avg /=1

Dynamic pressure at each rake is:

Qi, avg = Pt. - PS.i, avg i, avg

Average dynamic pressure is:

Qavg, ov [ P E[t Ps ] (W/A)i (W/A) iavg, ov i=l , avg i, avg i=1

The output from the data reduction program consists of computer-generatedplots of (1) radial total pressure and total temperature profiles; (2) circum-ferential total pressure, total temperature, and static pressure profiles; (3)circumferential airflow deviation profiles; (4) circumferential square root ofdynamic pressure ratio variation; and (5) the local airflow deviation map. Allof these plots will be discussed in detail.

RESULTS

J58 Compressor Rig)

The data analyzed during this study were acquired from two J58 compressorrigs during seven builds. Twelve points were chosen at conditions of sea leveltakeoff (SLTO) and supersonic cruise with variables, including inlet aerodynamicdistortion, modified seals, modified shrouds, and indexing of stators. Table Iindicates the matrix of variables examined. Because of the large quantity of datagenerated at each test point, only the results are presented here. The input datafor each case, along with the radial and circumferential computer-generatedprofile plots of the compressor discharge total pressures, temperatures, andstatic pressure are presented in Reference 1. Also given for each case inReference 1 is a computer-generated plot of the average circumferential varia-tion of the square root of dynamic pressure. It should be pointed out that theterm IDENT, used in all figures, refers to the rig-build-run numbers and thate/emax, used to indicate circumferential location, starts at top dead center andproceeds clockwise, looking upstream.

The compressor for the first build, rig 30204, build 10, was an engineBill-of-Material (B/M) configuration. Three cases were analyzed: SLTO, andcruise with and without inlet aerodynamic distortion. The compressor dischargeaverage circumferential airflow deviation for the SLTO condition is shown infigure 10. A four-lobe, nearly cyclic pattern of high and 16 w flow areas isapparent, with peaks ranging from 20% above average to 33% below average, fora difference of 52%. This term, the difference between the highest and lowestvalues of circumferential airflow deviation, is a convenient indicator of overallairflow nonuniformity and is used throughout this report for comparing differentvariables and builds. The airflow map for this test is shown in figure 11. The

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four areas of high and low airflow are again apparent, with the areas of highflow located near midspan of the compressor on horizontal and vertical center-lines. For this case the maximum local deviation is 52% above average whereasthe minimum is 88% below average. The absolute values of the maximum andminimum local airflows are not considered to be particularly significant inthemselves but do indicate the extremes of airflow variation when they lie in alarge area of high or low airflow. Results associated with each test case aregiven in table III and provide the best method of comparing results of the dif-ferent variables. The circumferential airflow deviations and local airflowratios are presented, along with the average test conditions at the compressordischarge.

Results of the test at cruise conditions are shown in figures 12 and 13.The four-lobe pattern of high and low airflow is still present but the magnitudeof flow variation is greatly reduced. The circumferential airflow difference isonly 27% as compared to 52% at the SLTO condition. The airflow map shows theareas of high flow have moved towards the hub of the compressor and the flowpattern has shifted about 45 deg in angular location relative to the SLTO test(figure 11). Results of the test at cruise conditions with inlet aerodynamicdistortion are shown in figures 14 and 15. The distortion imposed at the inletwas presented as a pressure distribution map in figure 7. The presence of inletdistortion produced little change in the circumferential airflow deviation, otherthan to increase the difference between highest and lowest deviations by 4% to 31%.The airflow maps with and without inlet distortion are nearly identical.

For the next two tests, the compressor was modified by installing reducedclearance, abradable type seals, in the first four stages to reduce interstageleakage. This configuration, Build 11, was tested at SLTO and cruise-distortedconditions. Results are shown in figures 16 through 19. The airflow uniformityis greatly improved, with the circumferential airflow deviation difference drop-ping to 19% at SLTO and 18% at cruise-distorted conditions. These values com-pare with 52% and 31%, respectively, for the compressor with B/M seals.Detailed comparison of the airflow maps for this compressor and the B/M com-pressor (figures 11 and 17 for SLTO and figures 13 and 19 for cruise-distorted)shows the reduced clearance interstage seals did, in fact, reduce interstageleakage because the airflow rate is increased at the hub of the compressor. Thefour-lobe cyclic airflow distribution observed with the B/M compressor is stillclearly present at cruise-distorted conditions. Its presence at SLTO is question-able but the airflow map does show high flow areas at midspan on the horizontaland vertical centerlines, as were present in the B/M compressor.

The compressor was next modified by installing reduced clearance inter-stage seals throughout the compressor to further reduce interstage leakage(Build 12). These seals were constructed of Feltmetal rather than the abradablematerial used in the first four stages of the previous build. The compressor wastested at SLTO and cruise-distorted conditions. The results are shown infigures 20 through 23. The circumferential airflow deviation difference is 19and 27% for SLTO and cruise-distorted conditions, respectively. This is muchimproved over the B/M compressor with values of 52 and 31%, but is inferiorto the 19 and 18% obtained from the previous build with reduced clearance inter-stage seals in the first four stages. Detailed comparison of the airflow mapsexplains this discrepancy. The compressor with tight seals throughout actually

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had more interstage leakage effects than the previous build with tight seals inonly the first four stages. This is evidenced by the larger areas of low airflowrate at the hub of the compressor with tight seals throughout (compare figure 17with 21 and figure 19 with 23) and indicates that the Feltmetal seals were lesseffective than the abradable seals. They were, however, more effective thanthe B/M seals during the SLTO test (compare figure 21 with 11) and henceproduced more uniform airflow than the B/M configuration. When the airflowmaps for tight interstage seals throughout and B/M configurations at cruise-distorted conditions are compared, it appears that the tight seal configurationhas more interstage leakage than the B/M (compare figure 23 with 15). Thiscan be explained in part by the fact that about 65 hours of testing occurred be-tween the two data points on the tight seal compressor. Thus, the seals mayhave worn considerably by the time the cruise-distorted point was taken. Inspite of the apparent relatively high seal leakage of the tight seal compressorat the cruise-distorted test condition, it provided more uniform airflow than theB/M configuration. The circumferential airflow deviation differences for thetwo are 27 and 31%. This deviation is explained by the fact that the tight sealcompressor, even after wear had occurred, had more circumferentially uniformairflow on the hub than the B/M compressor (figures 23 and 15). The tight sealcompressor had a line of average airflow of 70% around most of the hub circum-ference and an 80% line, which is continuous at a near-constant distance off thehub. This is in contrast to the B/M airflow map, figure 15, which shows air-flow rates around the hub ranging from 70 to 110% of average. Apparently theleakage of the worn tight seals is more circumferentially uniform than the B/Mseals and thus circumferential airflow distribution is more uniform. Examina-tion of the tight seal compressor airflow maps also shows the same characteristicfour-lobe pattern of high flow at midspan seen in previous builds.

A second J58 compressor rig, 30239, was used for the rest of the J58compressor testing described in this report. This rig is similar to rig 30204,but it brought a different set of compressor hardware into the program. Thefirst test with rig 30239, Build 202, was with a B/M compressor except thatthe B/M honeycomb tip shrouds had been replaced with Feltmetal tip shroudsin stages 5 through 9. This configuration was tested only at SLTO. Resultsare shown in figures 24 and 25. The circumferential airflow deviation differenceis 29%. This value, compared to the 52% obtained with a B/M compressor inrig 30204, indicates a significant improvement in airflow uniformity attributableeither to the Feltmetal tip shrouds or to unknown differences between the twocompressors. This deviation difference is not as good, however, as the tightinterstage seal compressor values of 19% obtained in the previous rig. Theairflow map again shows four areas of high airflow at midspan on the verticaland horizontal centerlines.

For the next test, the Feltmetal tip shrouds were replaced with B/Mhoneycomb shrouds, and the stators in stages 1 through 4 were indexed 36 degcounterclockwise, looking upstream. The configuration (Build 203) was there-fore B/M except for the indexing of the stators. This configuration was tested atSLTO giving the results shown in figures 26 and 27. The circumferential air-flow deviation difference dropped to 19%, which is similar to that obtained fromthe other compressor when reduced clearance interstage seals were installed.The airflow map again displays four areas of high airflow at midspan on thevertical and horizontal centerlines.

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The forward stators (stages 1-4) were indexed back to the B/M positionin Build 204 and testing was again performed at SLTO conditions. This compres-sor was entirely B/M. The results are shown in figures 28 and 29 and arepractically identical to the results of the previous test with the forward statorsindexed. The circumferential airflow deviation difference for both compressorsis similar, about 20%. Comparison of the airflow map of this compressor withthat of the compressor with Feltmetal tip shrouds (figures 24 and 25) shows thatthe two machines have similar airflow patterns except on the tip in the areabetween 1 and 5 o'clock. This difference apparently produced the 10% highercircumferential airflow difference for the Feltmetal shrouded compressor. Thelocalized low tip airflow of the Feltmetal shrouded compressor leads to thespeculation that one of the shrouds was eccentric with large clearance in thearea of low tip airflow. Comparison of results from this compressor with theother B/M compressor tested during the program reveals striking differencesfor supposedly identical compressors (compare figures 32 and 33 with 10 and 11).The airflow map for the first B/M compressor shows much larger areas of lowairflow on both the hub and tip of the compressor than the second B/M compressor.The circumferential airflow deviations for the two compressors are 52 and 20%,respectively. As the two compressors are thought to be identical, the cause ofthis large difference cannot be isolated, other than to speculate that the secondmachine was "tighter" in hub and tip seal clearances than the first. The four-lobe pattern of high airflow at midspan is present in both compressors.

For the next Build, 206, the second B/M compressor was rebuilt withforeign object damaged rotor blades in stages 2 through 9. The machine wasagain B/M except for the damaged blading. Tests were conducted at both SLTOand cruise-distorted conditions with this compressor. Results are shown infigures 30 through 33. Circumferential airflow deviation differences for SLTOand cruise-distorted conditions are 19 and 23%, respectively. The four-lobehigh airflow pattern is again clearly present in the cruise-distorted test but isindexed 45 deg from its usual orientation half-way between horizontal and verticalcenterlines. At SLTO, the airflow map is more indicative of an eight-lobe pat-tern. Except for the damaged blading used in this build, the results are directlycomparable to the other two B/M builds. Overall airflow characteristics aresimilar to the second B/M compressor but quite different from the first B/Mcompressor. No specific cause for these variations can be offered. Airflowdistribution is obviously sensitive to subtle build-to-build variations.

F100/F401 High Compressor Rig

The compressor tested during this program is the version of the F100/F401high compressor used in development flight engines. The rig, 34017, underwenttwo builds during the program, with the detailed compressor discharge instru-mentation shown in figure 9. During the first build, 204, the only variablesavailable were the test conditions: SLTO, subsonic cruise, and combat. Thesecond build, 205, was essentially the same compressor, but it was tested withinlet aerodynamic distortion and two levels of thermal distortion at the varioustest conditions. The combination of variables analyzed for these builds is shownin table I. As in the J58 compressor results, the compressor discharge totalpressure, total temperature, and static pressure are given as computer-generated radial and circumferential plots and in tabular form in Reference 1.Also given in Reference 1 is a plot of average circumferential variation of thesquare root of the dynamic pressure for each case.

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Results from tests of Build 204 are presented in figures 34 through 39,and in table IV. (Results for all F100/F401 tests are presented in this table.)Circumferential airflow deviation differences for the SLTO, subsonic cruise,and combat test conditions are 13, 11, and 16%, respectively. The airflowmaps for the three tests are quite similar and show none of the cyclic circum-ferential airflow variation observed in the J58. A second difference relativeto the J58 is the presence of high airflow on the compressor hub. The J58compressor consistently delivered low airflow on the hub and nonuniformityappeared to be related to the radial depth of the low airflow. The F100/F401compressor has its highest airflow near the hub and in this test series providedmore uniform airflow than the J58 compressor.

One of the purposes of Build 205 was to permit study of distortion effectson compressor performance. Several sets of data were obtained with and with-out aerodynamic distortion and different amounts of thermal distortion. Thefirst set of tests generated nondistorted inlet data for comparison with Build 204and subsequent tests with inlet distortion. The inlet pressure and temperaturemaps are given in figures 40 and 41 at the SLTO condition. They were obtainedwith eight equally spaced five-point total pressure rakes and an equal number oftotal temperature rakes at the locations shown. Maps for the cruise and combatpoints were very similar and are not shown. Results of the baselinecase arepresented in figures 42 through 47. During the taking of data for these cases,a scanivalve switch was not activated. Thus, no pressures were recorded overthe segment 83.5 deg to 150 deg. For the SLTO test case, the average circum-ferential airflow deviation (figure 42) ranges from 8% above to 11% below average.The local airflow deviation map (figure 43) still has the areas of high airflow onthe hub. Both the cruise and combat conditions show similar characteristicswith about 15% in circumferential airflow deviation difference and high airflowareas located on the hub.

Aerodynamic distortion was generated by a screen installed at the com-pressor inlet from 165 deg to 345 deg clockwise, looking upstream. The pres-sure map obtained at the inlet is presented in figure 48; one side is mostly belowaverage, the other above. The temperature map is similar to figure 41. Re-sults of tests at all flight conditions are presented in figures 49 through 54. TheSLTO condition results (figures 49 and 50) indicate an increase from 19 to 23%in airflow deviation due to the screen. The airflow map is similar to the base-line (figure 43). As in the baseline case, cruise and combat conditions givesimilar airflow deviations (figures 51 and 53) and flow maps (figures 52 and 54).

Combined aerodynamic and thermal inlet distortion was tested in the nextseries. The thermal distortion was imposed by heating the inlet airflow overthe same 180-deg sector on which screens were installed for aerodynamic dis-tortion. Two thermal distortions were tested, AT = 22 °K (40° F) and 55°K(100°F), over the range of test conditions. For AT = 22°K (40° F), the com-pressor inlet total pressure and temperature maps are given in figures 55 and56. Results with a thermal distortion of AT = 22°K (40°F) are presented infigures 57 through 62. The SLTO case shows another increase in averagecircumferential airflow deviation to 27% (figure 57). The local airflow map(figure 58) indicates an area of high airflow from approximately 30 deg to230 deg and low airflow over the remainder. This indicates a rotation or swirlof the distortion of about 45 deg through the compressor. The cruise and com-bat test data show the same results (figures 59 to 62). An increase in thermal

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distortion to AT = 55°K (100°F) drastically increased the average circum-ferential airflow deviation difference to 34% for SLTO and higher for the othertest conditions. The inlet pressure and temperature maps for the SLTO con-dition are given in figures 63 and 64. The results of all three test conditions(figures 65 through 70) show the increase in airflow deviation and the rotationof the distortion by about 45 deg. The areas of high airflow are still on the hubbut seem to have expanded.

When the aerodynamic distortion is removed and the thermal distortion setto AT = 22°K (40°F), the compressor inlet characteristics, given as total pres-sure and temperature maps, are shown in figures 71 and 72. Results fromthese test conditions are given in figures 73 through 78. The average circum-ferential airflow deviation is about 22% for all cases. At SLTO the local air-flow map shows high airflow areas in four locations off the hub. The cruiseand combat points do not show the same characteristic, however. When thethermal distortion was increased to AT = 55°K (100 °F), the inlet total pressuremap (figure 79) remained the same but the inlet total temperature variations(figure 80) were exaggerated. The SLTO compressor discharge characteristicsare similar to those found previously with combined aerodynamic distortion andAT = 55 °K (100° F). The average circumferential airflow deviation (figure 81)is 17% from average. The distortion remains rotated about 45 deg clockwiselooking upstream. The areas of high airflow as shown on the flow map(figure 82) are mainly on the hub.

DISCUSSION OF RESULTS

The J58 and F100/F401 compressors show a degree of compressor dis-charge airflow nonuniformity not previously documented. As these two com-pressors encompass state-of-the-art compressor design, it is highly probablethat comparable airflow nonuniformity exists in other compressors. This non-uniformity had not been previously observed becausc normal compressordevelopment instrumentation does not provide discharge measurements at alarge number of circumferential locations, and the variations that are observedare usually dismissed as instrumentation errors or exit guide vane (EGV) wakeeffects. In this program, however, the compressors were instrumented atmany circumferential locations, care was taken to ensure accuracy of themeasurements, and exit guide vane wake effects can be judged insignificant byinspection of the data.

Rigs were instrumented with five- or seven-point total pressure rakes ata minimum of twenty circumferential locations, single or five-point total tem-perature rakes at a minimum of ten circumferential locations, and wall staticpressure taps were spaced around the annulus on the outer and inner diametersat up to ten locations. Typical instrumentation locations were shown in fig-ures 4, 5 and 9. The measurement uncertainties were within ±0. 0758 N/cm2

(0.11 psi) on pressures and ±1. 11°K (2°R) on temperatures. The difference be-tween average total and static pressures on all tests was 1.38 N/cm2 (2 psi) orgreater and the variation within total pressures on each test was 2.76 N/cm2

(4 psi) or greater. Thus the ±0. 0758 N/cm2 (0. 11 psi) instrumentation uncer-tainty is small relative to the pressure variations encountered. Temperaturevariations were on the order of 11. 1°K (20 °R) or higher, so the temperatureuncertainties are a small effect, particularly since the absolute temperatureswere about 555. 6°K (1000 °R). Thus, a 1. 11°K (2 °R) error is only 0.2% errorin absolute temperature.

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The criticism that compressor discharge airflow nonuniformity data areunreliable because the exit guide vanes influence only some of the total pres-sure rakes and not others, is refuted by a study of a specific case for eachcompressor. For the J58 compressor rig, the data from the-cruise-distortedtest in rig 30204, Build 10 were used. The circumferential position of eachrake was determined relative to its nearest stator trailing edge and the localtotal pressure ratio plotted (figure 83). No pattern of EGV wake influence isdiscernable, although enough data points were taken to give adequate definitionto any trends. Since there are twice as many 9th-stage stator vanes as EGV'sthis plot also indicates no pattern of 9th-stage stator vane influence. For theF100/F401 high compressor, the subsonic cruise condition with aerodynamicand thermal distortion was analyzed in a manner similar to the J58 case men-tioned previously. Again, although sufficient data were available (figure 84),no pattern of EGV wake effects could be found. Stator vanes and EGV's areknown to produce local wakes but the instrumentation planes in both rigs arefar enough downstream that the local wakes have dissipated.

Results of this study have been presented in two forms: computer-generated plots (figures 10 through 82) and tables (tables III and IV). Becauseof the many test conditions and combinations examined while reducing the datafrom the J58 and F100/F401 compressor rigs, tables I, III, and IV are theeasiest to iuse. Table I presents the matrix of variables and test conditionsanalyzed. Table III gives the reduced data for the J58 compressor, whereastable IV has the F100/F401 high compressor reduced data. Several parametersare presented for comparison in tables III and IV. Among these are the circum-ferential airflow maximum deviation, minimum deviation, and difference; thelocal airflow ratio maximum and minimum; the number of high and low devia-tion areas; and the number or location of the locally high airflow areas. Alsopresented are the overall averages of the compressor discharge variables.

The parameter that displays the nonuniformity of the compressor dischargeairflow most dramatically for each case is the circumferential airflow deviationdifference. - For the J58 compressor, this parameter ranges from 18 to 52%.The greater the deviation difference, the greater the nonuniformity of flowleaving the compressor. All the J58 compressor tests investigated show rela-tively high nonuniformity of compressor discharge airflow. The worst caseinvestigated was the SLTO test condition of rig 30204, Build 10, which had acircumferential airflow deviation of +20% to -33% and a maximum local airflowratio of 52%. Four areas of locally high airflow were distributed about 90 degapart starting at TDC. The other two test conditions for this build were slightlybetter, with a low of 27% circumferential airflow deviation difference but thesame local characteristics.

In Builds 11 and 12 of the same rig, modifications to the interstage sealsbrought the circumferential airflow deviation differences down below 20%, butthe maximum local airflow ratios remained high, 33 to 39%. The reduced clear-ance interstage seals in stages 1 through 4 of Build 11 reduced the interstageleakage as evidenced by the increase in airflow rate at the compressor hub. Afurther modification to the interstage seals in Build 12 extended the tightenedseals through stage 9 and made use of Feltmetal to replace the abradable-typeseal. This change, however, showed no improvement over Build 11. In factthe cruise-distorted test point was more nonuniform than the same case inBuild 11, probably due to seal wear since about 65 hours of testing occurredbetween the two test points of Build 12. During both builds the four-lobe pat-tern remained in evidence, although somewhat hidden in the SLTO case ofBuild 12. 15

Rig 30239, Build 202, was a B/M configuration with modified tip shrouds.This build was tested at SLTO only and showed a circumferential airflow deviationdifference of 29%, as compared to the 19% deviation obtained in the previous rigwith tight interstage seals. The differences, however, can not be attributedstrictly to the shrouds, but could be caused by unknown differences in the twocompressor rigs. The next build, 203, returned to the B/M honeycomb shroudsbut indexed the first four stator stages 36 deg clockwise, looking upstream.This configuration was also tested at SLTO only and resulted in a circumfer-ential airflow deviation difference of 19%, which is the same as the SLTO casesof the previous rig with tight interstage seals. The stators were returned toB/M position for Build 204, giving a complete B/M configuration. The SLTOtests showed essentially no difference between Builds 203 and 204. Indexing ofstators, therefore, does not seem to have an effect on discharge airflow uni-formity. Comparison of the SLTO points from Builds 11 and 12 of rig 30204 andBuilds 203 and 204 of rig 30239 shows about the same uniformity for all builds;this indicates that the second rig, 30239, probably had tighter B/M seals tostart with. The final build, 206, was a B/M compressor with foreign-object-damaged blades. The circumferential airflow deviation difference for the twoconditions analyzed, SLTO and cruise-distorted, was similar to those of theprevious builds and no conclusion about foreign-object-damaged blading can bedrawn. The four-lobe pattern of airflow ratio and deviation is still evident inall cases.

For the F100/F401 high compressor, an average circumferential airflowdeviation difference for the baseline case is about 13% for Build 204 and 15%for Build 205 of rig 34017. The high and low deviations (table IV) are somewhatdifferent for each build. Whcn aerodynamic distortion is imposed, the averagecircumferential airflow deviation difference increases to about 19%. For aero-dynamic and thermal distortion, the difference becomes 26% for AT = 22 °K(40°F) and 37% for AT = 55°K (100°F). When only thermal distortion is imposed,the average circumferential airflow deviation difference is 22% for the low AT and35% for the high. These comparisons show that for the levels of distortion testedthe thermal distortion increases nonuniformity more than does aerodynamic dis-tortion by a substantial amount. Typically, aerodynamic distortion increasedthe circumferential airflow deviation difference about 3-4%, thermal distortionof 22°K (40°F) increased the difference by 6-7%, and thermal distortion of 55°K(100°F) increased the difference by 18-19%. Effects of the aerodynamic andthermal distortion seem to be additive in that the increases in average circum-ferential airflow deviation difference attributable to the aerodynamic or thermaldistortions separately may be numerically added to the baseline to obtain thesame deviation difference resulting from testing with combined distortions. Inmost cases the local high flow is concentrated on the compressor hub. Themaximum local airflow ratio for all cases was 40% above average. In general,the F100/F401 high compressor airflow is more uniform than that of the J58compressor but it is still relatively nonuniform.

Possible causes of compressor airflow nonuniformity are many, and whileisolation of specific causes and effects is generally beyond the scope of thisprogram, a discussion of some of the more obvious possible causes and perti-nent results follows.

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Stator Orientation - Because the performance of a stator passage is affectedby upstream flow properties, a repeating pattern of varying stator wake alignmentaround the compressor could produce a repeating pattern of varying airflow ratearound the compressor. The number of times the pattern repeats around thecircumference should be equal to the highest common integral factor of the numberof vanes per stage. In the case of the J58 the highest common factor for all stagesis two, but no two-cycle pattern was observed. The first four rows of statorvanes were indexed to break up the pattern and no change in the outlet flow fieldwas noted. However, the last five stages, which were not indexed, have a commonfactor of four and a four-cycle pattern of high and low airflow around the circum-ference was observed. This suggests that the four-cycle pattern was a result ofthe orientation of the stator vanes in the last five stages of the J58 compressor.This cannot be construed as positive proof of the source of the four-cycle patternas it was not possible to index the individual stages and change the orientation ofthe stator vanes. In the F100/F401 compressor, the largest common factor forall stages is two, and the last three stages have a common factor of six, but nocyclic pattern corresponding to these values was observed. Thus, the dataavailable on this subject are inconclusive but do lend some support to the theoryof stator vane orientation being a source for generating circumferential variationsin compressor discharge airflow.

Bleed Ports and Bleed Extraction - The presence of bleed ports or localextraction of airflow could locally alter the compressor aerodynamics and producenonuniform airflow at the compressor discharge. Both the J58 and F100/F401compressor rigs had provision for customer bleed. The J58 compressor rigalso had provision for interstage bleed. Customer bleed was not considered inthis program because its location was far removed from the compressor dis-charge instrumentation plane. In the J58, the customer bleed ports were locateddownstream of the diffuser, on the inner burner case wall. Thus, the airflowunderwent diffusion and a sudden expansion between the compressor dischargeand the location of the customer bleed ports. In the F100/F401 compressor, thecustomer bleed ports were located in the blunt trailing edge of the diffuser struts.The diffuser struts ended at the diffuser dump section, so again the airflowencounters diffusion and sudden expansion between the compressor discharge andcustomer bleed port locations.

The J58 interstage bleed was operated per normal schedule for all but onetest, but no data were generated with varied bleed flow at the same conditionwith the same compressor. Thus, no conclusions can be drawn concerning bleedeffects.

Circumferentially Nonuniform Tip and Seal Clearances - Any circumferentiallynonuniform feature of the compressor, such as nonuniform tip or seal clearances,is likely to cause nonuniform airflow due to circumferentially varying leakage andperformance levels. No known hardware circumferential nonuniformities weretested during this program, so no conclusive data are available on this subject.

Seal Leakage - Throughout this program, airflow nonuniformity appearedto be strongly related to interstage seal leakage. Seal leakage could not becorrelated with known variations in seal clearances but whenever seal leakagewas small, as evidenced by the presence of high airflow near the compressorhub, the compressor discharge airflow was relatively uniform. Increased sealleakage increased airflow nonuniformity. This trend persisted throughout boththe J58 and F100/F401 compressor testing.

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Exit Guide Vanes or Last Stage Stators - A circumferentially nonuniformairflow pattern is likely to exist immediately downstream of EGV's or statorvanes because of wake effects and the velocity profiles generated by turning inthe vanes. These effects are localized due to the close spacing and thinnessof the vanes and may be expected to dissipate a short distance downstream ofthe vanes. In this program, the compressor discharge total pressure probeswere randomly spaced relative to the stator vanes or EGV's, so that the probesmight be influenced randomly by the local airflow profiles coming off the statorvanes or EGV's. The data from both compressors studied in this program weresorted out in terms of probe orientation relative to the stator vanes and EGV'sand analyzed for local effects but none could be found. The local stator vaneand EGV effects appear to have fully dissipated at the instrumentation planesso the measured nonuniformities are truly present.

Downstream Struts - The diffuser case struts located downstream of thecompressor could conceivably cause circumferential airflow variation by blockageeffects. The rigs used in this program were all built with engine-type diffusercases with actual-size struts. The number of struts in all the rigs was eightbut no eight-cycle airflow variations were consistently observed. Only one testof the J58 compressor displayed an eight-cycle pattern, so downstream struteffects are not considered significant in the compressors studied in this program.

Inlet Distortion - Inlet total pressure distortion of the degree imposed inthe J58 compressor testing was found to have little effect on discharge airflownonuniformity. The total pressure distortion imposed simulated that distortionencountered in actual flight operation. In the F100/F401 compressor testing,classical 180-deg patterns of both total pressure and temperature distortionswere imposed. This distortion carried through the compressor, producing apattern of 180-deg nonuniformity with one high and one low airflow area at thecompressor discharge. The basic airflow nonuniformity pattern generated bythe compressor without inlet distortion was still present with the distortion;the 180-deg pattern was simply imposed over it. The effects of inlet distortionappear related to the compressor design and to the type of distortion imposed.

Build-to-Build Variations in Hardware - Unknown build-to-build variationsin compressors built to the same design produced large variations in dischargeairflow nonuniformity; these variations were as large as the changes producedby the modifications tested during the program. This phenomenon is mentionedto point out the subtleness of the causes of compressor discharge airflow non-uniformity.

Data Accuracy - Very accurate measurement systems are required inthe study of compressor discharge airflow distribution because small variationsin total or static pressure represent large variations in airflow. Most com-pressors operate with a discharge Mach number of about 0. 30. The velocityhead at this Mach number is 6% of total pressure; therefore, a 1% low readingfrom a total pressure probe would indicate an area about 9% low in airflow(assuming a uniform static pressure). A 1% high reading from a total pressureprobe indicates an area about 7% high in airflow. Variations of ±1% in pressurereadings from compressors are usually regarded as negligible, but they representsignificant variations in local airflow if the data system is accurate. In this

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program, pressure measurement uncertainties were within -0.2%. Erroranalysis predicts an uncertainty of ±2.9% on local airflow and ±2.5% on averageairflow at each rake location. These errors are small compared to the magni-tude of airflow nonuniformity encountered.

SUMMARY OF RESULTS

Both the J58 compressor and the F100/F401 high compressor show com-pressor discharge airflow nonuniformities not previously realized. The J58compressor shows the greatest airflow deviation in terms of average circum-ferential deviation and local airflow ratio. The F100/F401 high compressoris considerably better. However, it is more nonuniform than desirable fromthe standpoint of providing uniform turbine inlet temperature distribution.

The J58 compressor test cases indicated that no less than a ±10% averagecircumferential airflow deviation with a local airflow ratio 30% above average istypical of the nonuniformity to be expected from this compressor. The worstcase showed that the circumferential airflow deviation was +20% to -33% at-SLTO.Four distinct areas of high flow were found. Although several variables wereexamined, no significant improvements could be discerned. In fact, the build-to-build variations in the rig were larger than those produced by the variationsin compressor geometry. Comparison of probe readings at various locationsfrom the exit guide vane trailing edges showed no discernable pattern of EGVinfluence. These measurements are therefore believed to be unaffected by localstator or EGV airflow patterns.

The F100/F401 high compressor baseline test case is typically a +7%average circumferential airflow deviation compressor. The local airflowratio maximum is typically 20% above average. The high airflow regions areon the compressor discharge inner wall with only a few high spots as far awayas the 50% span. The typical average circumferential airflow deviation dif-ference for the baseline case is 16%. When aerodynamic distortion is added,this increases by 3-4%. When thermal distortion only is added, this increasesby 6-7% for AT = 22°K (40°F) and 18-19% for AT = 55°K (100°F). The com-bined effect of aerodynamic and thermal distortion seems to be additive, sothat the difference becomes 10% for AT = 22°K (40°F) and 22% for AT = 55°K(100°F) when both distortions are considered. Comparison of probe readingsat various circumferential spacings from the EGV's reinforce the conclusionthat the EGV's do not affect the compressor discharge airflow uniformitymeasurements as taken in this program.

The results of this program indicate that large airflow nonuniformitiesexist at the discharge of advanced turbine engine compressors. The onlyvariable that could be defined during these tests as having a marked effect onairflow uniformity was interstage seal leakage. If not attenuated, these airflownonuniformities could cause significant local increases in turbine inlet tempera-ture. Any contribution to these locally high temperatures from combustor mixinginadequacies could be additive. Efforts should be initiated to improve flow uni-formity in the compressor and to design diffuser-combustor systems that attenu-ate inlet airflow nonuniformity.

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LIST OF SYMBOLS

Area

Bill-of-Material

Clockwise

Counterclockwise

Airflow deviation

Local gravitational acceleration,9.807 m/sec2 (32. 174 ft/sec2 )

Inner Diameter

Rig-Build-Run number

Mach number

Number of total pressure rake locations

Number of total temperature rake locations

Number of static pressure tap locations

Outer diameter

Pressure

Dynamic pressure

Radius

Gas constant for air, 0.287 J/mol °K(53. 3 ft lbf/lbm °R)

Temperature

Top dead center

Airflow

Airflow per unit area

Local airflow/overall average airflow

Specific heat ratio

Circumferential location

3.1416

2 2 2 2m ,cm , (ft2, in.)

%

cm (in.)

cm (in.)

N/cm2 (psia)

N/cm2 (psia)

cm (in.)

OK (OR)

kg/s (lbm/sec)

g/s/cm2

(lbm/sec/in 2 )

rad (deg)

A

B/M

cw

ccw

Deviation

g

ID

IDENT

M

mn

ml

n

OD

p

Q

R

T

TDC

W

W/A

WA/WAA

-y

7r

20

LIST OF SYMBOLS (Continued)

Subscripts

avg or ave

avg. ov

ID

i

j

k

L

Local

max

OD

S

S

T

t

Superscript

' (Prime)

Average

Overall average

Inner diameter

Circumferential location

Radial location, or associated with a radial location

Inner or outer diameter location

Indicator, = 1 for J58, = 2 for F100/F401 case

Value at probe head

Maximum value

Outer diameter

Static

Static or plane of statics

Total

Total or plane of totals

Input static average

21

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Table II. J58 Compressor Rig Discharge Instrumentation Combinations

Noo Pt No. Tt No. Ps (OD)

26262626263333

3332322727

11111110101111

1010101212

8886666

33388

No. Ps (ID)

0000000

00088

23/24

C

Rig Build

30204302043020430204302043020430204

3023930239302393023930239

10101011111212

202203204206206

Run

531552569185171114453

81437

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ical

J58

Co

mp

ress

or

Dis

char

ge

Inst

rum

enta

tion

FE

120

469

B

w

GO

1.105 (63 deg 30 min

5.07 (291 deg) 1.187(68deg)

09 (287 deg) 1.34 (7 deg)

2 (282 d 1.344 (77 deg)

5 (273 deg) _ .(.1.501 (86 deg)2 2 deg) 1.571 (90 deg)

5)8 26 de)_ - 1.693 (97 deg)

4512 )1.850 (106 deg)

1.955 (112 deg)

)8(2464 deg)(~

41 deg 30min)4.136 (237 deg ) 2 .129 (122 deg)

3~ ~~~~~~~~~ ~' .822 (297 deg)- .~.5dg

3.979 (228 dc/l @%eg)/ 2.286 (131 deg)3.927 (225 deg) / / ;A\ \ .356(135deg)

3.822 24 4deg) /

3.665 (210degg' 3 //|\ 2.601 149deg)

3.508 (201 deg )_\ 2.758 (158 deg)3.430 (196 deg 30 min 2.880(165 deg)

3.351 192deg) 3.037 (174deg)3.19 (183 deg) 3.142 (180 deg)

Note: All locations were not used for every build. Typical of all builds except 206. Angles are given in radians (degrees).Key:

o Total Pressure Thirty-Three 7-Point Rakes( Total Temperature - Eleven Probes* Static Pressure - Eight Wall Taps, OD

Figure 4. Typical Schematic of J58 Compressor FD 63335Discharge Instrumentation, LookingUpstream

34

Note: Angles are given in radians (degrees).

Key:o Total Pressure - Twenty-Seven 7-Point Rakes( Total Temperature - Twelve 5-Point Rakes* Static Pressure - Sixteen Wall Taps

Figure 5. Schematic of J58 Compressor Discharge FD 63334Instrumentation, Looking Upstream;Rig 30239, Build 206, Runs 136 and 182

35

TDC

Figure 6. J58 Cruise Inlet Distortion Screen, Looking Upstream

FE 104135B

36

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0-

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tatic

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ress

ure

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ure

8.

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ical

F10

0/F

401

Hig

h C

ompr

esso

r D

isch

arge

In

stru

men

tati

on

FE

112

884A

Note: Angles are given in radians (degrees).

Key:* Static Pressure0 Total Pressure$ Total Temperature

- Twenty Probes- Twenty-Two 5-Point Rakes. Ten 5-Point Rakes

Figure 9. Schematic of F100/F401 CompressorDischarge Instrumentation, LookingUpstream

FD 63333

39

_~~~~~~~~~~~~~7 -,7 - -- - - -

:._ _ .- _. ... _ ~ _.L " - '- - ' '_'

--- C TRCUMFERENT IRL RI'RFLOW DEVIRT I ON-58:-" " - - ~ ~~~ J'-5 8 - ----0

.... : ..-- IDENT = 30204- 10-0531:

-- -- : SLTO. UNDIST. B/M - K

0~~~~~~~~~~~~~~~~~~

-o

' °- tf PAX= :6. 283 RRD- 1360 DEG)

U~~~~~~~~~~~~~-'

~cc

0

-o

mLa:

LLr-0

ALICio

(c1:

L.5

Ah 3 .:.. . o

a-0

'0.00 20.00 40. 00 60.00 'B. 100.0

_ -:-:; '/e,,l0o -.: CIRCUMFERENT IRL LDRTI ON

Figure 10. DF 92056

40

I

COMPRESSOB D ISCHRRGE FLOW MAP

J-58 COMPRESSOR RIG - - WLOCAL/WRVE OVERALL

MRXIMUM FLOWRRTIO = 1.523

MIN FLOWRRTIO = 0.116

RIG NO.30204 BUILD NO.= 10RUN NO.=0531 CONDITION- -- SLTO UNDIST.

Figure 11.

41

DF 92063

,-'-r 1 .- - -, 1

-CICUM-FER:ENTIRL RIBFLOW- DEVIRTION;_ . . ..- :~' , . . -: - .. .. . ~- _£ ¢:' . ,'-i:T~'. ~ ' ! ,, ,.

- : '- ,J-58 8'IDENT = 302014- 10D-0552

::- - CRUISE. UNDIST. B/M - K _ --- -

-R= 6.283 RAD (360 DEG)

CD0cm

c:)_

0

(U

m

=CD

o--

crc

LU

CD

::o

4 :

I-, , ,t-c;

Z . .': ' - ''a:~~~~~~~~~~~~~~~~~~~~~~~~~~~~

19 . --0 . :-

~"J ;,' ' ,'T-- T ' · -- '' -.r

'0.00. 2000 qO.00 60.00 80.00 00-.0 · /ex'0 -CIRCU'FERENTIRI L'OC:RTION

Figure 12. 9

42

COMPRESSOR D I SCHRRGE FLOW MRP

J-58 COMPRESSOR RIG - - WWLOCRL/WRVE. OVERRLL



RIG NO.30204 BUILD NO.= 10RUN NO.=0552 CONDITION- - CRUISE. UNOISr.

Figure 13. DF 92070

43

C I RCU1F EBENT IRLA R AR FLON DEV IRTI ON, :i- - --8. :

-I -. ' " . -' .. . . . .L . .-7 ~.--

':~4 1 :; b N.T 4T-=', 302:4- 1 0-0569~-- - - IOE, 04-;~ -,

- - .-CRULSE._ I.ST. B/M - K

0_0 ---- - - /,' - :- ' ' r;'' - . ,-- - '

....... -- X-= 6. 283 RROB (360 DEG)

C2. : . - :

0

I,-- -

>_Lu

0

LILO

C;'

.0 8. a. o 6b. '' 80.00 '-1.1

e:" 0 A ,0 -, CIRCU.4FERENTIL LOCATION

-e ' 1..0 D : 9 205'- ,'''.\ , _-,ur 14 -F 920.;58:-:E

:'o-f o o ~b:-o'o f -0:oo"' ' - _ ~~~~~~~~~~~~~~~~~~~~~~~~~~~~~' 'tO0*s r\.00 .,~~~~~~~~~ ~0 ,0 80.O

~/,,xO -o X .RCFEETIR L;RT

Figure 14. DF 92058

44

COMPRESSOB D I SCHRRGE FLON MRP

J-58 COMPRESSOR RIG LOCAL / WRVE. OVERALL

+ MRXIMUM FLOWRATIO = 1.310

\\/ X MIN FLOW••~~~~~ ~RRTIO = 0.138

RIG NO.30204 BUILD NO.= 10RUN NO.=0569 CONDITION- - CRUISE, DIST.

Figure 15. DF 92064

45

CIRCUMFERENTIRL RIRFLOW DEVIRTIONJ-58

IOENT = 30204-11 -0185

TIGHT SEALS STRG. I-4

e& = 6.283 RRO (360 OEG1

'o oo 2b. ooe/ mAx xxI O0 - 7.

4b. oo 610.00 8b.00 0o.00ooCIRCUMFERENT IRL LOCRTION

Figure 16.

- SLTO. UNOIST.

00

O-

0

C00u

m,

(U

-o

mR

30Cr5-I

D

JO

rr-

cc0

0

I0

C-,cr0

(U.

00D

0

46

DF 92055

>.~r't

COMPRESSOR D ISCHRBGE FLON MAPJ-58 COMPRESSOR RIG - - WLOCAL/WAVE. OVEALL

MAXIMUM FLOWRATIO = 1.393

X MIN FLOWRRTIO = 0.244

RIG NO.30204 BUILD NO.-=11l RUN NO.=0185 CONDITION- - SLTO, UNDIST.

Figure 17. DF 92062

47

UMFERENT I AL: J-58

RIRFLON DEVIRT ION

IDENT = 30204-11 -0171

- CRUISE, DIST. TIGHT SERLS STRG. I-4

= 6.283 RRD (360 DEG)

e/e Omxl O0 - . C I RL LOCRTI

Figure 18. DF 92054

48

CIRC,

CD -C

e"X

LIt::

,--

COMPRESSOR D ISCHRRGE FLOW MRP

J-58 COMPRESSOR RIG - WLOCAL/WAVE OVERALL

+ MRXIMUM FLOWRRTIO - 1.335


RIG NO.30204 BUILD NO.=11 RUN NO.=0171 CONDITION-- CRUISE.DIST.

.Figure 19. DF 92071

49

CIBCUMFEBRENTIRL' RIRFLOW DEV]J-58

IDENT = 30204- 12-0114

SLTO, UNDIST. TIGHT

e. RX= 6.283 RARO

SEALS

IRTION

i 1-9

t360 DEG)

'o.oo o5.00''o 0.00 .0 - 0 2. C O 0 R80.0 100.00 Oe/e,moxl0 - 7. CIRCUMFERENTIAL LOCATION

Figure 20. DF 92053

0oCD

CD

CD,

0

0

Cu

crc

woU-

!

d

0

00

LL~CD

30

CED°

LLiCE-

cc

Lio

0

Z __

0

0

aCDLc;

cuI

C3

50

1n ! : :; . :: e I : ' %:

I .

- I I I I . , - -- . - -. L I i~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~

COMPRESSOR DISCHRRGE FLOW MAPJ-58 COMPRESSOR RIG - - WLOCAL /NWAVE OVERALL


X MIN FLOWRATIO = 0.295

RIG NO.30204 BUILD NO.= 12RUN NO.=0114 CONDITION- SLTO, UNDIST.

Figure 21. DF 92061

51

C I RCUMFERENTIRL R I RFLONJ-58

IDENT = 30204- 12-0453

CRUISE. DIST.

DEVIRT ION

TIGHT SERLS. 1-9

e-:= 6.283 RAD (360 DEG)

1xlOO - % C80.00

RL LOCRT I

Figure 22.

52

DF 92050

0o0l

ZD

L.IO

COMPRESS09 DISCHRRGE FLON MRP

J-58 COMPRESSOR RIG - - WLOCAL/WvE. OVERALL

+ MRXIMUM FLOWRRTIO = 1.334

X MIN FLOWRRTIO - 0.306G

RIG NO.30204 BUILD NO.= 12 RUN NO.=0453 CONODITION- CRUISE, OIST.

Figure 23. DF 92069

53

CIBCUMFEBENTIRAL RIBFLONW DEVIARTI'ONJ-58

IDENT = 30239-202-0008

-- SLTO, UND

-I <3-110X = 6.

IST. MOD.SHROUDS, STRG 5-9

283 RRDA 360 DEG)

L ON

Figure 24. DF 92052

54

§ 00

LiM

a w % I11I . I S.w

COMPRESSOR DISCHRBGE FLOW MAPJ-58 COMPRESSOR RIG

LOCAL /AVE. OVERALL

MRXIMUM FLOWRATIO = 1.436

\N o\9 - - - - Do J X X MIN FLOWRATIO = 0.223

RIG NO.30239 BUILD NO.=202RUN NO.=0008 CONDITION- -- SLTO, UNDIST.

Figure 25.

55

DF 92065

C IBRCUMFERENT I-RL -R-IFLONW DEV IRT IONJ- 58

IDENT 30239-203-0014

--$T. UNIT-SLTO. UNDIST. INDEXED STRTORS 1-q

ex= 6.283 RRD 1360 DEG)

00D

I ,~~~~~~~~~o0.00oo 2.00oo9'/eRX xl O0O -

qb.oo 60.00o 8b.00 ibo.ooZ CIRCUMFERENTIAL LOCATION

DF 92051Figure 26.

56

00CD

aJ -

0

0-"" ..,

Q

Cu

z

coa -__

W

F- D

30

Wo

La...

b-q

cr0

0

tio

Cl.

!C

_ . 7 . fX

i

~~~~~~~~~~i

I j

COMPRESSOR DISCHRRGEJ-58 COMPRESSOR RIG - - WLOCRL/

FLON MRP

WAVE. OVERALL


\ _ ___ / X MIN FLOWRRTIO a 0.268

RIG NO.30239 BUILO NO.=203RUN NO.=0014 CONDITION- -- SLTO, UNOIST.

Figure 27. DF 92068DF 92068

57

_ ~~~~~~~~~~~~- - , . !' 0' :,- (z, _r ' - t ' _.. : , ,,,. .,,,

C--- C UMIE R FENNT IRL R I RFLO 0 DEV

IDENT = 30239-' 5 80037:~~~~ .:- :; DENT -- '3023920.4-0037-:

SLTO, UNDIST

e, , = 6.283 RR

B/M-K

0 (360 DEG)

;[ 00' ' ' 2'0.00. o.9oo 20. O0.'/eX x! 00 - '.

i ~~I I I40.00 60.00 80.00 100.00CIRCUMFERENTIRL LOCRTION

Figure 28. DF 92049

58

I RT ION

CO- CD

0

; l

0

zD

c-o-I-0

LU0

U-

,>,

cC0

3o

~- 0

,o°

LZ'"8

._ ,,.I

La:J

*0 t 'U-cI -io

0

0.

COMPRESSOR D I SCHRRGE FLOW MAPJ-58 COMPRESSOR RIG - - WLOCRL/WAVEOVERLL



RIG NO.30239 BUILD NO.=204RUN NO.= 0037 CONDITION- -- SLTO, UNOIST

Figure 29.

59

DF 92067

CII9CUMFERENTR- : -_ D .. :

C I RCUMFERENT I RL R I RFLOW DEV] I RTION-J-58

IDENT = 30239-206-136,

-- CRUISE DIST.

'Mx = 6.283

B/M - K

RRD (360 DEG)

xlOO - X. CI ERENTIAL LOCRTI

DF 92048Figure 30.

60

L*.Cl

oo

COMPRESSOR D I SCHPRGE FLOW MAPJ-58 COMPRESSOR RIG - - WLOCAL /WVE OVELLWLOCRL/WAVE. OVERALL

MRXIMUM FLOWRRrTIO = 1.277


RIG N0.30239 BUILD NO.=206RUN NO.= 136, CONDITION--- CRUISE DIST.

Figure 31.

61

'DF 92066

+,.' t '3I Lr -04

CI RC u M F R-E JN T:tN A - R FL -DE VIR IN

-I - . -- -- +.- IDENT -30239-2R0 3- :.0182- D. -- i

U-~~~~~~~~~~~~~~~~~~~~~~~~~~I

- -- ST- UN IT. -'- 'J- K'8 ........: .. ' w;: : 2; t ta 0 - -- ,: ;- : .' .:._ ̂ :.- -' - /- .:

...... --.: - u283 RRD 30 DEGI.

,~ '---: ,-:.'' $t; uNIS. -, --- ~t '1 Kz ' -. -R* i ...

* 0-' . : _ - -:: t-, , '-' . .. ' - .... '

'U-

.0 , 02 n .00 6

s: ' : 7 -] '

--o: 'i' '4' -- .. -. - '-z

1.1 .9

62

-i,

LL~~~~~~~.-

cr-

': ° .II!.-:~ - -' :.- ~o. :_: .-,.......,.' __0 O -'

0

LLJO - D -, :;X

' .00 -'2b. oo , b. oo 6b.0o0 8b.00 ' ibo. oo:/,mxl00 .. CIRCUMFERENTIRL LOCAlrION

Figure 32. DF 92059

62

COMPRESSOR D ISCHRRGE FLOW MAPJ-58 COMPRtSSOR 9IG

ILOCRAL /RVE. OVERALL

+ MRXIMUM FLOWRRTTO = L.379

\\/v 2/ X MTN I-LOW-X = ~~~~~RrGa = 0. O047

RTG NO.30239 BUILD NO.=206RdJN NO.- 182 CONOITTON- -- SLTO UNDIST.

Figure 33.

63

DF 92060

U~~E RE, -3S r- t. '. :'~ --O . -gE-;.v-i~:N/' -.! '~ ~ ~ ~ ~ -I - R1B O D' I' I

7- ,C-I RCU KF:ERES I RL-RF'-N;:X--D'V 'RT'O- . -. .-i./ i+0

,.~~~~~~~~~~~~~~~~~ .-..

I- : IDENT = 34q 17-20q4-8031

L.$LTO.OP LINE. ,UND.,-: I- I. -. : -I,

., i - ... - ;- t x -- '+ l t'S -E - t

: , I - .1

·. - 6. 2s- .AD oE)28 3 RFM0 (36 - DEG1:

Z4o. o b.0 o.0 8b.oo - bo.oo ----

CIRCUMFERENTIAL LOCRT I ON

Figure 34.

64

00

'-

DF 92091

- I L, --: I

7- - --":I

I I . I ..

.I -, i -. j; .. . .

I

I . ra-i - , -

COMPRESSOR D I SCHFRGE FLOW MAPFIOO/F401 HIGH COMPRESSOR RIG - - WLCRL/WAVE OVERALL



RIG NO.34017 BUILD N0.=204 RUN N0.=8031 CONDITION- SLTO,OP LINE, UND.

Figure 35. DF 92072

65

:,_ - :_:, _, _ ' _ ,:, _ /S, ¥ j , ~ _ -_.- _? _:: ... i:~H . . .J __~ _! -_

-C-1RTUNF ALENTIRRL-.RI RF:LOQ- DEV IRTJ I ON

: -iE T _ ~. I . r-- .;e --- 13N~ -§307 -- -80,23 ;- 0

-, SUBSONI C, CRUI:SEP UNO.-:4 : - ' ''- -; I' -. , . - II~ . I .

-0O

0l

, .,x8 "A..x= -". 283 RRD (360. OEG)

Figure 36. DF 92092

66

0

, ,tb.,,I:Ci .::

,me0 .

:sw

l , ; -,, - - .I ' .I - :,

COMPRESSOB D I SCHRRGE FLON MRP

F100/F401 HIGH COMPRESSOR RIG - - WLOCAL/WAVE. OVERALL

MRXIMUM FLOWRRTIO = 1.150O

X MIN PLOWRRTIO = 0.65G

RIG NO. 34017 BU I LD NO. -204 RUN NO.=8023 CONDITION-SUBSONIC CRUISE, LUND.

Figure 37. DF 92073

67

C I R CU MFE:R EN TI RE I-R:I R F -DEV i -T I ON- -F100O/F4 1:- :

-IDENT = 340.17-204-8014 ..... ., P- : . - , . , . . .~~~~~~~~~~~~~~~-

COMBRT OP LINE,UNO.

9RR0 (360 DEG)

.. ; - S f

, i . . -. - . .. .

. .

. . . .

: . - : -:: : E

CD-'O-- - ,- - !, - i , N i ~~~~~~~, . , .,

0.00 2b.0 o b.oo 6b. o o. o o 40 bo. ooe/eAx X1 00 -:i C I RCUMFERENTI-RL LOC-RT I ON-_--

Figure 38. DF 92093

- I {: ; 1 1 1

-r

- . .-

I: I ' . I

=.- 6. 283

0

C:)

d.

0CD

Cuj

z0

ED

.- o

O

0_

o:

co

0

z-r-Li

a: U

,_

LJoa - a

a0

I

68

I I , . -

d - : ..

,, , . ; -- I I

I . :.

.~~~~~~~~~~~

COMPRESSOR DISCHRGE FLOW MRPF100/F401 HIGH COMPRESSOR RIG

- -WLOCAL / WAVE. OVERALL


X MIN FLOWRrIO = 0.611

RIG NO.3U1017 BUILD NO.=204 RUN NO.=801LI CONDITION- COMBRT OP LINE,UND.

Figure 39. DF 92074

69

T CD- ,- -CD 0 -: q 0

- N M V M

ZU, z

00CV,

CNLin

a)

a)

a-co

0.ME

Cn

c

C

'a)a -

o

c

' (D)

'0a)

0

v

0)

L)

qcoCD,

CO

-S

10oCeC1i

0

PA

.F

o,-

._

0._

0

Cd

CI,

.4Ja1)

0o

o

0

U

C>

C)

bfl

'-4

0

00

'-4

70

CO

~~Z0)

0) z

l

0)0)

LqCN

CN6

0)0D

CO0Co

0)c7)0)

71

oa) CNo - -

0)

-

coI-

F-

n,.E

c',

wr_

cu.t

- N 0

.c_ CD

U, CDw a)

'ciLi,,CNA

CN0)

Ol

Ctr0

P-q

e~

co

--400

0z

Cab

F)

a5

©Q)

7_

.40ooCd

a)

0

0s

Cd

Is

ONL')

C IRCUMFERIENTItL Rt:IRFLON

':-E-N F 1 00/F40 51I DENT = 3.4O17-205-tL353

SLTO,UNDISTq

BASELINE CONFIG.

'.~ t- 6.283 RBAD (360,DEG)

; , I. I I;

.00 20.00, _ :i.0. O ..xl X~,

· i~~~~~~~~~~~~~~~~~~~~~~~~~~~.b.00o 6b. oo 0b.0 100o.ooC- IRCUMFEREN-T RL LOCAT I ON

Figure 42.

DEVIRTION

00

d.,

0.

I .

cu

z

Co

bo5o.

o

tir

LLJ

LU'

r:

o

. 0

C-)'LLJC

a 9

C

: C;a n-

72

DF 92094

COMPRESSOR DISCHRGE FLOW MAPFlOO/FLIO1 HIGH COMPRESSOR RIG - - WLOCRL/WAVE OVERALL


MIN FLOWRATIO = 0.580

RIG NO.34017 BUILD NO.=205RUN NO.=q4353 CONDITION- SLTO,UNDIST

Figure 43.

73

DF 92075

C I RCUMFEBRENTTIKRL RJI RHFLOW -~DE~YI'LRT I1,.-N ..:LF-1-00/F6401 -

IDENT = -34017-205:-4363

.SUBSONIC CRUISE.UND BRSEL I NE , C.ONF I.G .

6.283. RRD -1360 DEG)*. 1- f --

'o oo - 2b.ooe/ ex xl 00 -

- b. o - 6b.DO ^ b.'oo bo. oo -%. C I RCUMFERENT I-RL LOCRT ION

Figure 44.

-mx. -

00

0-Io

,00

01

: X

CU

-- o

"a~c. 0

5r-4

"o

Dotroc_j

U-

-I

:

I-o

CD-

InU o.

0Z-

0

0

74

DF 92095

. I

COMPRESSOR IDISCHRRGE FLOW MRP

F100/FL40O1 HIGH COMPRESSOR RIG - - WLOCRL/WAVE. OVERRLL


MIN FLOWRATIO = 0.575

RIG NO.31O17 BUILD NO.=205RUN NO.= 4363 CONDITION-SUBSONIC CRUISE,UNO

Figure 45.

75

DF 92076

~~~.R- v ': ti. ,i- F' Z' ¢t$., " -

" n" ;- __ C'I-R9CU'M-EHENT IRL-H "R9FLO DDV' IRT N- i - '-F -' DET.-; D1, OTYD 1

' ' : -- :" ~-IDENT =' 340~1:7-205'L313 - - -'

COMBRST, UNDIST+ -.W. i D . i.~~~~~~~'

or

, ~ -...

CD)

0

! i-7 0

LI-

0

'.

L'

0

Co

'0

Li~

L

s0

3-0

Lii

, a:

i jo

- 0.:

Figure 46.

76

e; -~~~MA

.BRSELINE CONFIG._I I

- 6.283 RRD -(360 OEGI

- -d fw

- l

* . : : I : .

L - | , 1 - :

! _ . _ _ . . . .

, . f? T-' - : .

. , . g i

s. ;:, i . . F . .

_ . #

: , F fl . , . F

:, f : . g , .

, ,

. n f ,7

- : . . R R . - -, - . .

, ' -. : .

. .

. , .

I I

I .

0oo 2 'b.oo : -.oo 6b.0oo eob.oo ibo.oo'e,:xl000: ' ...eC-I-RCULM:FERENTIRL-- LOCRTI ON-

DF 92096

I L,

COMPRESSOR D I SCHRRGE FLOW MAPF100/F401 HIGH COMPRESSOR RIG 7 - WLOCRL/ JVE. OVERLL


X MIN FLOW

RRTIO = 0.616

RIG NO.34017 BUILD NO.=205RUN NO.=4313 CONDITION- COMBRT, UNDIST

Figure 47. DF 92077

77

RT CD X0C0 qI C O

0 CNU)r*D q 0 .o _ _

- N M- T Lo (0 1,-

C..,XC o _4 C

o 00 o

- CN

0)-E~M _

o~~~~~~~~~~~~C J C

n

0 C

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d)

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iF

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d)

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4q)

. -4

CI,c*

C>

P4p

0s

-a0)

-o

co

CD)0

78

UJp2 19 im, 7- 'C tLW -

A~ S ~it0 FE ,6 t 8 SCtr+ ~mg :gEEN

p+ - * gW ij

ff~~~~~i. i

rrlt'f fl'

4~~~~~7

1i.t1

r~'~ --~-~I-y~--~~'-~_~~~~ ~~,~~~~~~~~~~,~~~: 4 7t,- 44,1 t'.

r&, T~~ ''--

I:

S S~~~~~~~~~~~~~~~~~~~~~~~4CW. :TT, tt ~ ~ OCI~O

Figure 49. DF 92097

79

COMPRESSOR D I SCHRRGE FLON MAPF100O/F401 HIGH COMPRESSOR RIG LOCAL /AVE. OVERALL


MIN FLOWRRTIO = O.GO1

RIG NO.34017 BUILD NO.=205RUN N0.=4233 CONDITION- SLTO, DISTORTED

Figure 50. DF 92078

80

: gi+~~~~.- 4- - " W.m- ~ ~ ~ f~~JT 3 fV - 2l Z 2 3

:s !I I .T- F I

J," rJ t E,

... ' ';

'1 '

rL-,;-l:;. ;' L4 L:;.:KS4J1';±t: :;'::c; ;;';:jbiTr:+.)Tl,

In S rW7t-1i5_W T , ' -

i ,1 1 7 -t .rj-,4-:J - -1

m mfrl 4 .+ r e -.1t .gr iS ~:~*-~f: ~~1 i t<95;5i+: t* .-~. r~t> : t _If.+ .i .-- ~ t- *'~. _ ." *zzjI. z'D¢.L w, ~- -~~ t-~. . I~_ =-~:-.:-;. ,:** 3f- -- ID .

'

' ** r z ;'-'*-I ,- . .z-3_ K ., -:- '::!

4. # tr;4 *Utr -- *' !

7:2I

S ;;i m g g . ,tt'e~~~~~~~~~~~~~~~a~ ;

,14w rt; -$i: En..t'. t,) --hij; .'!e'$tEi *r~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~W ' !*H,__ ,

I i+.ai .*.- *-o-ro - - -I I , r i

S eXX~~~~~~~~~~~~~~~~~t

Figure 51. DF 92098

81

~~~~P,4 ~ ~ '' . ,.=r1} 1, ._7:1 C I FrI .

'':;-*Ft, tffl-t#S# ,i ; t-4'~ -

"E i, fi ' t - +- F:-i*,- -4L ,, r^-t-Z

':4

--!±l

r[

I

IT

FT

r

&LZ

t-;. :, t-['

_ F, ; t5 + i=Ti-7-J-+nW . 7 47H+ ~ ~ ~ -I-+F LI4 3 ~s + Fttnv k fir! ~,=tl hh~ '84'RRntt8 n EJE

COMPRESSOR D ISCHRRGE FLON MAPFIOO/F40L1Ol HIGH COMPRESSOR RIG - - WLOCAL/WAVE. OVERALL



RIG NO.34017 BUILD NO.=205RUN NO.=4263 CONDOITION-SUBSONIC CRUISE.DIST

Figure 52. DF 92079

82

M ILSr I E RF T l aE F R I-.

t COMB~T NE X I 4 RER&1mO'~ t SCRl Wf>4yuo -tLI <I " 14 i >r7 B'; '. jrH -t ~, tH X+ X mt W-~'r~ 4, i1vrr. ~- QE -<.,

.

. r {!;i I ......

CIV~~~~~~~~~~~~~~~~~~~~~~

r-tdlL '1 + tt t;''_ 'r=FIt;1 S j ¢

t*+r -:--,t*- :ttir I:; t{4~-p: m r- - L~ .0-:t t - 7'''7: 'i +

(e *t t ~~~~~~~~~~~~~~~-'- '- =- ~_~ m .$t z-i-JXt--

t,-: i,.

t.E~.~2 00" j / L&0 ' :8

,tEL

k ~ ~~~~~~~~~~~~~~~~~~~~~~ --4

4,_~

~-~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~~-~! ~I',~ 4 4-~'

,;e-. :_. P-i,,~ '1- "- ~1P.'"' ..... ,,,t H

, <~~~~~~~-~; c ,-.:--~, ,; ' --

+o.;.oo/7.1'::t E+:[~;~o~.oo,~~.i/ ].~o, oo~-_~. ~o~:0o ~ "'; "~~~'4El- >'Pt-

ta r _Z; S _; r7 ̂ t*_ * r .ia 1-' T7% r .7

Figure 53. DF 92099

83

-- I: ... - , -~./ .... .. ..... -~~1E ,jQ-,U-~ rU .~ -..... ..-1

COMPBESSOBR DISCHRBGE FLOW MFRP

F100/FL401 HIGH COMPRESSOR RIG - - WLOCRL/WAVE OVERALL



RIG NO.34017 BUILD NO.=205RUN NO.=4213 CONDITION- COMBRT OP LINE,OIST

DF 92080Figure 54.

84

0N LD 0.O O 0 - - 96 a o-- -

_- N M d L1 (0 I-

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RIG NO.34017 BUILD NO.=205RUN NO.= 4453 CONDITION- SLTO, DISTORTED

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COMPRESSOR D ISCHRRGE FLOW MRPFlOO/FL401 HIGH COMPRESSOR RIG - - WLOCRL/RVE. OVERALL

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97

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Figure 84. F100/F401 High Compressor Discharge DF 92115Total Pressure Distribution Relative toExit Guide Vane (EGV) Trailing Edge

114

RE FERENCES

1. Shadowen, J. H., and W. J. Egan, Jr., "Evaluation of CircumferentialAirflow Uniformity Entering Combustors from Compressors, Volume II -Data Supplement," NASA CR 121010, 1972.

2. Keenan, J. H., and J. Kaye, Gas Tables, John Wiley & Sons, Inc.,New York, 1965.

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