Energy efficient engine high-pressure turbine component rig performance test report

'I TABLEOF CONTENTS Section Title Page 1.0 SL_IMARY l 2.0 INTRODUCTION 2 3.0 HIGH-PRESSURE TURBINERIG DESIGN 4 3.1 DesignGoalsand Requirements 4 3.2 High-Pressure TurbineComponentDesign 4 3.2.1 Overview 4 3.2.2 Aerodynamic Design 6 3.2.3 Mechanical Design 9 3 . 3 High-Pressure T urbineC o m po nentRig Design 14 3.4 Rig Fabrication and Assembly 15 3 .5 Te s t Instrumentation 18 3. 5.1 Perf o rmance In s trumentation 18 " 3.5 2 StructuralIntegrityInstrumentation 24 3.5.3 Instrumentation Calibration and Accurac y 24 3.5.4 Data Acquisition System 28 4.0 TEST PR_RAM AND TEST FACILITIES 29 4.1 Test Program 29 4.1.1 Full StageTurbineTest 29 4.1.2 Vane CascadeTest 32 4.2 Test Facility 33 _ 4.3 Data Reduction and Analysis 33 4.3.1 Ana_lysis of Full StageData 35 4.3.2 Analysisof CascadeData 36 5.0 TEST RESULTSAND ANALYSIS 37 5.1 Introduction 37 5.2 Mechanical Performance 37 5.3 Aer o dynamic Performance 38 5. 3 .1 Tur b ineStage Pe rf o rmance Assessment 38 5.3.2 Vane CascadePerformance Assessment 55 i 5.3. 3 TurbineBlade P erf o rmance Analysis 69 5. 3 .4 Se co ndary F l ow SystemPerformance Characterization Studies 77 5.4 P o s t-Te s tIn s pection Resu l ts 8 6 5.5 Summaryof Result s 86 6.0 CONCLUDING REMARKS 90 APPENDIXA 91 APPENDIXB 94 LIS T OF SYMBOLS 235 REFERENCES 236 DISTRIBUTION LIST 237 : iii ,!

w_ f LIST OF ILLUSTRATIONS Number Title Page 2-I OverallProgramSchedule 3 3.2.I-I High-Pressure Tu bine Component for the Flight Propulsion System 5 3.2.2-I High-Pressure TurbineFlowpath 7 3.2.2-2 Vane Mea r , SectionAerodynamic Contourand Pressure Distributi,_n 7 3.2.2-3 BladeMean SectionAerodynamic Contourand Pressure • Pistribution 8 3.2.3-I High-Pressure TurbineRotorAssembly lO 3.2.3-2 TurbineBladeCoolingSyztem II 3.2.3-3 Vane ar.dInnerCase Assembly II 3.2.3-4 TurbineVane CoolingDesign 12 3.2. 3 -5 TurbineBlade Tip Seal Assembly 13 t 3.3-I High-Pressure TurbineComponent Test Rig 14 3.3-2 TurbineSecondary Flow SystemAir SupplyLines 15 3.4-I Completed TurbineRig Static Structure with PWA 1422 Directionally-Solidified Turbine Vanes 16 3.4-2 Completed TurbineRotorAssemblySingle Crystal (PWA 1480)Blades 16 3.4-3 Assembled TurbineTest Rig 17 3.5-i Instrumentation Map of High-Pressure TurbineComponent Rig 19 3.5.1- 1 Circumferential TraverseRakewith Instrumentation 21 Locations and TraversePath 4.1-I Test Envelopefor Full Stage TurbineTest Program 31 4.2-I Pratt& WhitneyAircraftX-203 Te__t Facility 34 !

iv LIST OF ILLUSTRATIONS (Continued) Number Title Page 5.3.1-I TurbineStageEfficiency Trendsas a Functionof Pressure Ratioand SpeedParameter 40 5.3.1- ? TurbineStageEfficiency 40 5.3.1-3 TurbineReactionCharacteristics 41 5.3.1-4 TurbineSecondary Flow SystemMap ShowingPredicted and MeasuredFlow Ratesand Pressures 43 5.3.1-5 TurbineInletSpanwiseTotalPressure 45 5.3.1-6 TurbineInletSpanwise TotalTemperature 46 5.3.1-7 Instrumentation and Circumferential Measurement Locations 47 5.3.1-8 Exit SpanwiseEfficiency by Quadrant 48 5.3.1-9 AverageSpanwiseEfficiency 48 5.3.1-I0Tfficiency ContourPlot of One Vane Gap in FirstQuadrant, ShowingMaximumEfficiency Near the MidspanRegion 49 5.3.1-IISpanwiseProfileof TurbineExit Total Pressure 50 5.3.1-12TurbineExit TotalPressureContourPl o t 50 5.3.1-13SpanwiseProfileof TurbineExit Total Temperature 51 5.3.1-14TurbineExit TotalTemperature ContourPlot 51 5.3.1-15BladeAir Exit AngleCharacteristics 52 5.3.1-16Air AngleContourPlot 53 5.3.1-17AverageSpanwiseAir AngleTrendsComparedto the Design Prediction and PrecedingUncooledRig Test Results 53 i 5.3.l-IBBladeExit Mach NumberCharacteristics 54 5.3.P-I Vane Loss Trends 56 5.3.2-2 Flow CapacityCharacteristics 57 i J v f, k LIST OF ILLUSTRATIONS (Continued) Number Title 5.3.2-3 Vane Deviation VersusMach Number 57 ; 5.3.2-4 Vane CascadeInletTotalPressure 58 5.3.2-5 Vane CascadeInletTotalTemperature 58 5.3.2-_ Circumferential Vane Exit PressureCharacteristics 59 5.3.2-7 SpanwiseVane Exit PressureProfile 60 5.3.2-8 SpanwiseVane Loss Characteristics 60 5.3.2-9 Vane Loss ProfileShowingthe Influence of Coolingon Performance 6 1 5.3.2-10ContourPlot of Vane Loss Characteristics 62 5.3.2-11Circumferential Trendsof Vane Exit Temperature 63 5.3.2-13ContourPlot of Vane Exit Temperature 64 5.3.2-12SpanwiseProfileof Vane Exit Temperature 64 ! 5.3.2-14SpanwiseVane Exit Air AngleTrend_ 66 5.3.2-15ContourPlot of Vane Exit Air Angle 66 5.3.2-16Comparison of Air Angle Trends 67 5. 3 .2-17SpanwiseProfileof Vane Exit Mach Number 67 5.3.2-18Vane Root Section(ll PercentSpan) PressureDistribution 68 5.3.2-19Vane Midspan(50 PercentSpan)PressureDistribution 68 5. 3 .2-20Vane Tip Section(89 PercentSpan) PressureDistribution 69 5.3.2-21Vane SuctionSurface F ilm CoolingEffectiveness 70 5.3.3-I Calculated Blade PressureLoss as a Functionof Exit Mach Number 71 5. 3 . 3 -2 B l ade D ev iation VersusMach Number 7l i • vi L.TST OF ILLUSTRATIONS (Continued) Number Title Page 5.3.3-3 SpanwiseBlade Exit Air Angle 73 5.3.3-4 SpanwiseBladeExit Mach Number 73 5.3.3-5 SpanwiseProfileof Vane InletAir Angle 75 : 5.3.3-6 BladeSpanwiseEfficiency 76 5.3.3-7 BladeTurningCharacteristics 76 5.3.4-I Tangential On-BoardInjection SystemRig Geometry 78 = 5.3.4-2 Tangential On-BoardInjection Rig Instrumentation Locations 78 5.3.4-3 Tangential On-BoardInjection Rig RadialPressures 79 5.3.4-4 FrontRim CavitySensitivity to TangentialOn-Board _- Injection Flow Rate Variations 81 5.3.4-5 Rim Cavity Pressures with BladeTangential On-Board Injection Flow Rate Variations 81 5.3.4-6 Rim CavityTemperatures with BladeTangentialOn-Board Injection Flow Rate Variations R2 5.3.4-7 Rim CavityPressures with Mini Tangential On-Board Injection Flow Rate Variations 83 5.3.4-8 Rim Cavity ' Temperatures with Mini Tangential On-Board Injection Flow Rate Variations 83 5.3.4-9 FrontRim CavitySensitivity to SwirlLevelDifferences 84 5.3.4-10FrontRim CavitySensitivity to PressureVariations 84 5.3. 4 -11 H ig h -Pressure C o mpressor DischargeSeal Leakage StudyResults 85 5.4 - I High-Pressure Tur b ineR o tatingDisk Assembly 87 5.4-2 Post-Test Condition of TurbineVanes 87 vii u LIST OF ILLIISTR#TIONS (Continued) : Number Ti t___l e 5.4-3 High-Pressure Turbine Outer Air Seal S e gments 88 5.4-4 High-PressureTurbine Full Stage Rig Exit Probe _' InstrumentationRing Showing Galling and Metal Pickup 89 - , _ viii . ,i LIST OF T_BLES Number Title Page 3.1-I High-PressureTurbine Efficiency 4 3.2.1-I High-PressureTurbine Technology Features 5 3.2.2-I General Aerodynamic Parameters 6 3.2.2oii Design Gas Triangles 8 3.2.2-III High-Pressure Turbine Aerodynamics After Restaggerin g 9 3.5.1-I F]owpath Instrumentation 20 i 3.5.1-I! Secondary Flow System Instrumentation 22 3.5.1-III Flow Measurement Instrumentation 25 L ; 3.5.1-IV Speed Humidity and Vibration Instrumentation 24 3.5.2-I Rig Safety System X-203 Stand 26 3.5.?-II Turbine Component Structural Integrity Instrumentation 27 3.5.3-I InstrumentationAccuracy 27 3.5.3-II Measurement Uncertaintyof Turbine Parameters 28 4.l-I Test Matrix for Full Stage Test 30 4.l-If Turbine Vane Annular Cascade Test Conditions 32 5.2-I Turbine Blade Tip Clearances 37 5.3.1-I Comparison of Performance Parameters 38 5.3.1-II Full Stage Turbine Warm Rig Test Results 39 5.3.1-III Clearance Adjustment 42 5.?.l-IV Secondary Flow System Flow Sensitivity Test Results 42 5.3.1-V Full Stage Turbine Warm Rig Secondary System Coolant Flow Test Results 44 5.3.I-VI Design Point Efficiency 47 _X a LIST Or TABLES (Continued) Number Tit]_ Page 5.3.2-I High-Pressure Turbine Annular Cascade Test Conditiors and Results 55 5.3.2-II Cascade Loss by Quadrant 62 5.3.3-I Wall Static Pressure 72 5.3.3-II Internal Aerodynamics 74 5.3.4-I Energy Efficient Engine Front Rim Cavity Data 80 5.3.4-II A_tachment Leakage for Vane 85 A-I Test Rig Prog r am 93 : B-I Full Stage Turbine Warm Rig Test Results 95 _._ B-II High-Pressure Turbine Cascade Test Results I Tl }

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b SECTION l.n , _ SUMMARY I As part of the NASA-sponsored Energy Efficient Engine program, Pratt & Whltney Aircraft successfully completed a rig test of the cooled high-pressure t,mbine . component. The principal objective of this test was to substantiate the tur- ' bine design point performance as well as determine off-design performance with the interaction of the secondary flow system. The program was organized into both full stage and vane cascade tests. A total of 216 hours of testing was accomplished, and there were no major difficulties that prevented the acquisi- tion of performance data. Performance data were acquired at 13 principal con- , ditions during full stage testing, covering the intermediate and high power operating range of the ;ntegratedcore / low spool.

The measured efficiency of the turbine component was 88.5 percent. This sur- passed the rig design goal of 88.5 percent, the estimated component efficiency _ for the integrated core / low spool of 87.1 percent and the goal for the flight ' , propulsion system of 88.2 percent. De measured efficiency was repeatable, In- 1 dic_ting that no performance deteriorationoccurred during the test.

l Cascade testing with the rotor removed showed that vane performance was gen- t erally in good agreement wilh the design prediction. In addition, these series of tests verified the effectivenessof the film cooling design.

Blade performance, on the basis of analysis, exceeded expectations. Perfor- mance trends showed that low loss design concepts, in conjunction with the in- I troduction of trailing edge cooling, are effective in redu:ing losses at high exit Mach numbers. Results also showed that there is a slight compromise in rig performance when using engine hardware in a rig environment.

The secondary flow system in the turbine performed according to the design in- tent. C h aracterization studies showed that system performance, in particular the pressure-balancedtangential on-board injection system, is insensitive to flow an d pressure variations.

Overall, this test demoqstrated that a highly-loaded, transonic, single-stage turbine can achieve a high level of operatihg efficiency. In addition, it p_ovides the confidence that the compor, cqt is suitable for testing in t|,_in- tegrated core / low spool.

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,I SECTION 2.0 I NTROD UCT ION The Energy Efficient Engine Component Development and Integration Program, sponsored by the National Aeronautics and Space Administration (NASA), is directed toward demonstrating the technology to improve fuel efficiency and to reduce operating economics of future commercial gas-turbine engines. The pro- gram 9 oals include a reduction in fuel consumption by at least 12 percent ard a reduction in direct operatina cost by at least 5 percent relative to a base Pratt & Whitney JT9D-7 turbofan engine. To demonstrate the technology to ac- complish these goals, the program is o, anized into two main technical tasks: Task l Flight Propulsion Syste,_Analysis,Design and Integration Task 2 Component Analysis, Design and Development Under Task 2, an advanced high-pressure turbine component was designed for the Energy Efficient Engine. This turbine is a single-stage configuration and has various t_chnology features in the areas of aerodynamics, structures ard materials / coolinL_.

_I Substantiationof this turbine design involved two component rig test programs.

" _ The first, as shown in Figure 2-I, wa_ a test of the uncooled turbine rig and _ was conducted as part of the Uncooled Rig Supporting Technology Program (Ref.

l). The results Crom this effort corroborated the aerod3mamic design assump- • tions and established the uncooled efficiency b_se for the turbine component detailed de s ign. The se c ond test, w h ich i s the s ubje c t of this report, veri- fied the performance of tt'ecooled turbine co,lponent.This program involved both full stage and annular cascade testing, ap! it was directeJ toward de- monstrating design and off-d_;signperformance as well as assessing perforrr_nce sensitivity to the secondary flow system. With some exceptions, the turbine component used in this te.;t is the sap_ a.: designed .'orboth the flight propulsion system -- the analytical study engine in the Energy Efficient En g ine program -- and the integrated core / low spool test vehicle.

The results of the Cooled H;gh-Pressure turbine Component Rig Test program are summarized in this report. The followin c section, Section 3, presents a de- scription of the turbine component as designed for the integrated core / low sponl and the design similarities and differences in the turbine rig. Section 4 outlines the test program and describes the test facility and data reduction method. Section 5 presents the results of both the Full stage and annular cas- cade tests, in addition to an analysis of these results. This section also contains results from a related test on the blade tangential on-be'_d injec- tion system. Concluding remarks are presented in Section 6.

Two appendixes are included in this report. Appendix A contains additional in- formation on the supporting blade tangential on-board injection rig test.

AC endix I_ contains a presentation of performance data from both the full stage test and the annular cascade test.

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SECTION 3.0 HIGH-PRESSURE TURBINERiG DESIGN 3,1 DESIGN GOALS ANDREQUIRE_NTS # The high-pressure turbine component efficiency estimates and goals for the rig are summarized in Table 3.]-I. The estimatefor the integ;-ated core / lowspool was based largelyupon the resultsobtainedfrom the earlierUncooledRig Test Programand then updatedto reflectrefinements in the designas the component : definition evolved.

TABLE3.1-I HIGH-PRESSURE TURBINEEFFICIENCY STATUS Aerodynamic DesignPoint-- 10,668m (35,000 ft), 0.8 Mn Standard Day Condition i_ Integrated Core / Low ,_ FlightPropulsion Spool Component Test Rig System Goal (%) Status (%) Goal (%) 88.2 87.1 86.5 The goal for the test rig was definedat the beginning of the Energy Efficient i EngineProgram.Other key design parametersincludea specificwork output of 448,000J / kgm (192.96Btu / lbm), an expansionratio of 4.0, a combinedturbine cooling / leakage flow rate of 13.2 percentof core engin_ flow,and a rim speed of 527 m / sec (1730ft / sec).

3.2 HIGH-PRESSURE TURBINECOMPONENT DESIGN 3.2.1 Overview The high-pressure turbinecomponentis shown in Figure 3.2.1-Ias definedfor the Energy EfficientEngine flight propulsionsystem. It is a single-stage configuration designedto operate at a high velocityratio and low ratio of throu_hflow to wheel speed (Cx / U).The design is based on advances in aero- dynamics,structuresand materials / cooling management.Some of the advanced design featuresare listed in Table 3.2.1-I. Detailsof the design are con- t_ined in Reference 2.

I f ,- i LAB SEAL 9 - LIP DISK PRE S WtRLER SIDEPL A T E S LOW L E A KAGEA T TACHM| T HERMA L BARR E_ _ R C O ATED S C 2 0 00 '- PLATFORMS CR Y STAL PWA S C 20Q0 BLADES (_ ) ; S INGLECRY S TAL ;fVE B LA DE T IP V ANES ( 2 4 ) ( S I C GRIT _I C ERAMIC '_ O UT E RAIR S EA L INTERNA L AC TIVE CL E A R A NCE C O N TRO L Figur e 3.2.1-I Hig h -Pr e ssure Tu rbin e C o mp o nent f o r th e Fl ig h t Pr o p ul si o n Systern TABLE3.2. 1 -I HIG H -PRESSURE T URBINE T ECHNOLOGY FEA T URES REDUCEDCOST CONCEPTS: REDUCEDCOOLANTFLOW CONCEPTS: Sing l e-Stage Turbine Sing l e-Stage Turbine ReducedNum b erof Airfoils Improved Airfoi l Cooling Effect ivenes s Sing l eCrysta l Airf o ilMat e ria l s INCREASED AERODYNA H IC EFFICIENCY ThermalBa r rier CONCEPTS: _ Coatings High AN Z / HighRim Speed Efficient CoolantSupp l ySystem Contoured Vane End Walls Low Windage Low Loss Airfoi l s ReducedTip Loss Configuration ActiveC l earance Contro l High Airfoi l Loadings REDUCEDLEAKAGECONCEPTS: ReducedLeakageLength J I mp r o v e d Gap S e a l ing I I mp r o v:d R i m S e a l ing ' t W-S e al s 5 w_ , 3.2.2 Aerodynamic Design The general parameters governing the aerodynamic design of the turbine compo- nent are listed in Table 3.2.2-I.

TABLE 3.2.2-I GENERAL AERODYNAMIC PARAMETERS (Aerodynamic Design Point - Mn 0.8; 10,668 m (35,000 ft)) PTIN, MPa (psia) 1.324 (192.1) CET, K(°R) 1633 (2940) RIT, K (OR) 1561 (2811) • -._ N (RPM) 13232 AH, (Btu / se_ ) 13384 -= FPin, (W _ / TT / P T) i6.98 il _ / A, (%Wae) 4.0]4"10 Reaction 43.0 percent Velocity Ratio, _ / U2 / 2gJ_h 0.556 NASAWork Factor, (Z_h / UZ) 1,62 !

Cx_U 0.351 AN _, (IN 2 RPM 2) 4.U6 x I0 I0 URIM, m / sec (ft / sec) 481 (1580) CTearance, cm (in) 0.0469 (0.0185) The turbine flowpath is shown in Figure 3.2.2-I. In the single-stage configu- ration, there is a total of 24 vanes and 54 blades. The vane is characterized by aerodynamic sections having a blunt leading edge and a long chord with the maximum airfoil thickness near the leading edge. Th,e inner vane endwall is cylindrical,while the outer wall is contoured In an _" shape. The blades are highly tapered with a conical inner wall.

Figure 3.2.2-2 shows the aerodynamic definition of the vane mean section and corre s ponding predicted pressure distribution. Similar information for the mean section of the blade is shown in Figure 3.2.2-3. Velocity triangle data for both the vane and blade are contained in Table 3.2.2-II.

To achieve the desired low-pressure turbine inlet aerodynamic conditions in the integrated core / low sponl test hardware, the turbine blade was restaggered opened 0.25 degree from its aerodynamic definition. The effect of restaggering is shown in Table 3.2.2-III I \ 43 (I 7 ) -- ROOT MEA. . . _ _ N TI. _ P 4.356 CM V I _ NE BX _ (1.715 IN) i '_ _ B L ADE BX 3.4 2 4 C M 2 .9 46 C M 2 .54 0 C M ( 1 . 348 I N ) (1 , 16 IN) (1. 00 IN) - 42 . 11 8 (1 6 .._ 8 2) _.

" _ , _ 40 . 949 41.018 _J 40 (16) -- (16.122) (16.1 4 9) z OC LU I - Z 'J ANE L U (,_ 38 ( 15) -- u , J z ' ' Z " - • , j j' '" OF PO OR QUAL I T Y O 35 (14) _=_ _ 3 5 .255 (13.880 / I I , =, u _ 35 . 135 (13.833) 34.742 m C3 (13.678) < n - 33(13) 1 J I } :, 0 2 . 5 5. 0 7. 5 1 0. 0 . (0) ( 1 ) ( 2 ) ( 3) (4 ) : ' A X IA L L EN GT H, C M ( IN) ' Figure3,2.2-] High-Pressure TurbineF]owpath o 0 5( 02) 1 0 ( 0 4) V ANE MEAN SE C T IO N 1 5 ( o 6) PRESSURE O ISTR I BUTION 20( 0 8 l 2 5 ( 101 3 5 (14 1 j 4011 L E 45(1 riP" 07 _, s o(2 De 301 1 2 1 09 A _ OIN G E_ ' _ ' __ _; 55( 2 05 -- i u 6 01 2 o4 I I I | I C 6 S ( 2 O0 02 04 08 09 10 _ o,2 X / B 75{ 3 B 0 ( 3 RAD I US 3 11. 1 0 3 CM (r E 0 01 IN I 5 5 ( 3 I F OI L S 24 90 13 AXIAL CHORD 4 _ CM (1,71§ IN) * ( D IA M ET E R I 3 34 CM (0, S : t l 5 I N I 9 _ >t3 T E D IAMETER 0 111 _ CM (0 . 0 6 4 IN) UN C OVER E DTUR N ING 9J) ° I0 0 ( 4 0 1 EXIT W E DG E ANGLE 4.0e 1 0 5 14 2; M MA x O 927 INLET F O IL ANGL E 9 0 O" O0 10 20 3 0 40 EX fT F OIL _ NGLE 1 0 312 = I O0 _ ( 0 41 (081 (1 2I I1 el , CHORD 10 Il lg CM (42 E I1 IN) CM (IN) Fig u r e 3.2.2-2 Van e Mean S e ctionAer od yna m ic C o nt ou rand P r essu r e Di s tri bu ti o n ,_ , ', F_,.. , I " o R | G h' . _,L 45118= O F pOO R Q_A I' i ' I_ B LADE MEAN SE C TION 4 0 ( 1 6 _ PRESSURE DISTRIBUTION 1 0 35(1 4 ) _ 09 FAD I NGE D GE P / PT o. E _ 25(10) 05 3011 2} _ _ 0 7 z v :s 0 .4 ( - _ 2 0 I 0 8 ) -- 0 . 3' O2 000 0.20 0.40 0 60 0 8 0 1 . 00 1 5 (0 6) -- X / B = RADIUS 37,879 C M 114 ,0 13 I N ) I 0 ( 0 4) -- # F O I L S 54 AX I AL CHORD 2.946 CM (1 . 160 IN) L E DIAMETER 0 . 394 CM 10.155 IN) 0 5 (0 2 ) -- T.E . D IAMETER 0.1 5 3 C M (0.OGO I N) UNC O VERED TURNING 6 . 0 ° EXIT _= :DGE AN ,_L E 2.00 • INLET WED G E ANGLE 30,0 ° ' _ ; 0 0 (0 O ) " ] I M N INLET 0 .32 0 O 0 1 0 2. 0 3 . 0 M N E XIT 1 . 217 - _ (0 0) ( 0 4) ( 0 8) 11 2) I NLET F OIL ANGLE 34,0 ° , | C M ( IN ) EXIT FOILANGL E 16 80 = - MMA x 1 702 • i CHORD 4 956CM11 771N) Fi gu r e 3. 2 .2 - 3 B l ade Mean S ec ti on A e r od y na mi c Con t ou r and Pr e s s ur e Distributi o n _ TABLE 3 .2 . 2-I I DESIGNGAS TRIANGLES R o o t M e an Ti p VANE InletAir Ang l e (deg) 90 90 90 Exit Air Ang l e (deg) 1 1 .6 10. 3 9.1 InletMach No. 0.09 0.08 0.07 Exit Mach No. 1.0 0.92 0.85 Gas Turning(deg) 78.4 79.7 80.9 BLADE In l etAir Angle (deg)* 38.5 34.0 74.0 InletAir Angle (deg)** 33.5 42.7 63.6 ExitAir Ang l e (deg) 15.9 16.9 1 7.7 In l etMach No. 0.36 0.25 0. 1 4 i ExitMach N o . 1.22 1.2 4 1 .28 r Gas Turning ( deg) 13 0.6 1 20.4 98.7 i ExitAbs Air An g l e (deg) 38.0 43.8 48.4 i E x it Abs Mach N o . 0 . 54 0 . 52 0.5 2

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* With inl e tt em p e rature and vane ]oss profile ** F l at in l ett e mperature and fiat vane lossprofile 8 , ,' _ J ] • L TABLE 3.2.2_ - III HiGH-PRESSURE TURBINE AERODYNAMICS AFTER RESTAGGERING HPT Restaggere _ HPT Designed HPT Run Run at LPT FP (Initial IC /LS) At LPT FP (Final IC/LS) FPHp T IN 76.983 16,983 17.023 t FPHp T OUT 66.562 68. 165 68.165 PR HPT 3.98 4.093 4.084 Reaction (%) 43.0 43.8 42.4 A_HPT (%) BASE 0 to -0.3 0 to -0.15 Mn HPT OUT 0.523 0.554 0.539 _HPT OUT (deg) 43.8 43.0 44.0 LPT Convergence Vl Root 1.4 1.35 1 . 4 B1 Root 1.3 1.25 1.3 : The secondary flow system in the hig h -pressure turbine is d e sign e d to maximize the use of secondary air for cooling and thrust balance as well as minimize . parasitic leakage and the attendant performance penalty. The primary design features that enhance leakage control include: o A tangential on-board injection (TOBI) system for positive blade coolant flow supply o A front rim cavity mini tangential on-board injection (TOBI) system o Boltless and full ring rotor sideplates o A multi knife-edge, stepped high-pressure compressor discharge seal.

3.2.3 Mechanical Design The major subsystems in the turbine component are the rotor system, vane and inner case assembly, and outer air seal. i Turbine Rotor Assembly The Energy Efficitnt Engine high-pressure rotor construction is different from most previous Pratt & Whitney Aircraft designs in that the rotor is straddle mounted. This arrangement e l irnlnates the bearing compartment forward of the high-pressure turbine disk and places it after the disk. The turbine rotor assembly is illustrated in Figure 3.P_.3-1. Because of high rim speeds, the de- sign is characterized by a thick bore region. The rim has a f;rtree attachment to hold the blades, shelves to support the front and rear sideplates, and a flange to support the vortex plate. The vortex plate is used to contain biade cooling air ana provide a passage for free-vortex pressure rise to augment the pressure of the flow exiting the tangential on-board injection nozzle. The curved elliptical cooling air holes supply coolant from the vortex plate to the blade root. Pumping action through the curved elliptical hole also in- creases the pressure before the flow enters the blade root cavity.

The air-cooled turbine blades are retained in the disk by boltless sideplates.

The ful_ ring sideplates perform _ dual function of blade retention and seal- ing in the rim areas.

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TO HIGItPRESSURE C O MPRE S SOR _} R ADII A RE ME A S U RE D FRO M BOLTED JOINT ENGINE C ENTER LINE 22.3 CM i8.8 IN) R (8 . 0 IN) R • r,, VORTEX PLATE DISK . ..,i OF pOOP, QL,,_\ L _ ' T'i 1 SIDEPLATES 34 . 5 C 1_4113.6 IN) R i : 40 _ CM t16 . 0 I N) F ADE q Figure 3. 2 .3-] High-Pressure Turbine Rotor Assembly A schematic of the cooling design is s h o wn in Figure 3.2.3-2. The blades re- quire only 2 . 75 percent of the core e ngine inl e t flow for cooling. They are cooled by a combination of internal convection and local film cooling. The internal geometry is designed to enhance cooling convection while external surfaces are ]oc d lly film cooled from the leading edge showerhead holes and tip pressure side holes . There are no film cooling holes on either the pres- sure or suction surfaces of the airfoil, Vane and Inner Case Assembly The vane and inner case assembly is illustrated in Figure 3.2.3-3. The primary elements are the vanes, the tangential on-board injection (IuBI) system and the high-pressure compressor discharge seal.

The turbine vanes are also an advanced air-cooled design. Excluding the inner and outer platform surfaces, the vanes require only 6.4] percent of the core engine inlet flow for cooling. A schematic of the vane cooling system is pre- , sented in Figure 3 . 2.3-4. Effective internal cooling of the vane is achieved by cross flow impingement, augmented by strategically placed external film cooling holes. Cooling air enters the vane at both the tip and root and is distributed within the three internal cavities. The showerhead holes are an g led to provide maximum heat transfer in the thick leading edge region. The f _ pressure surface is fil,_cooled by two sets of double row holes. The suction { surfar_ has three rows of holes. Th e platform cooling scheme consists of a - combination of impingement cooling under the platform and convection cooling on the platform gas path surface.

: 1 0 w & < BLADE COOLING FLOWS {TOTAL 2.75%) 0 2 6% _ o_, O F POOR QUALI T Y 0 4 0 % _ 2 0 5 % 0 7 0 % 1 72% 033 % Figure 3.2.3-2 Turbine Blade Cooling System i lIOGsI _P H _ A _ R _ii_ S _ E AL _LIb_ " ___ :i: : S Y STEM "'__ ' _ " _"" ' _ V ANE INNER SUPPORT Figure 3 .2. 3 - 3 Van e a n d Inn e r C ase A sse mb l y l I ' ' t ; ' 0 .325% | 1 . 52% TOTAL COOLING FL O W 6.41 % WAE ,| - t , Figure3.2.3-4 TurbineVane CoolingDesign To minimizeleakagecaused by vane twisting, both the innerand outer surfaces are clampedalonga chordalcut. By having a chordalcut, axial tiltingof the vane, introduced by d ifferential axial growthbetweenthe innercase and outer combustor case, is allowedto occurwithoutbindingup or openinga leak path.

T he vane platforms are sealedby featherseals to preventleakageof compressor discharge air into the turbineflowpath.

The main or b l ade tangentialon-b o ard injectionnozzle i_ a cascade design that providescoo l ingf l ow at a positivesupp l ypressureto the disk rim and b l ade. A secondaryor mini tangentia l on-boara injectionnozz l e is used to swirl th e coo l antf l ow to the front side of the turbine disk to reduce disk hea+-upcausedby windagesheareffectson the frontsidep l ate.

The high-pressure compre s sordischargesea l is a knife-edge l abyrinthcon- figuration designedto maintainminimumc l earance at a 11 operatingconditions.

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\ Or,,_ • OF PC , C, : '_ : Turbine Outer Case and Airsea] The turbine Llade tip seal is part of the internal a c tive clearan c e control system . The turbine outer case and seal design are shown in Figure 3.2.3-5.

The major c omponents in c lude the fr o nt and rear outer airseal support rails, outer air sealshoe and impingement ring. To minimize c ooling air leakage, "W" seals are used on the front and rear hooks of the shoe, and feather seals are used at the cir c umferential ship lap joint between shoes. The sea] assembly is supported by the high-pressure turbine outer c ase whi c h incorporates the mani- fold for the active clearance control system.

The active clearance control system maintains c lose blade tip clearances at all operating c onditions by impinging c ontrolled temperature air from the high-pressure compressor on the outer air seal support rails. The cooler tem- perature air reduces the rate of thermal expansion to c ontrol the radial move- ment of the seal shoes towards the blade tip. The blade tip clearance at the cruise condition (the aerodynami c deRign point) is 0.047 cm (0.0]86 in).

2 IMPINGEMENT RING

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)UTER AIRSEAL SHOE "W" SEAL , i _-COO L I NG AI R MAN IFO L D HPT INN ER G UI D E REAR O UTER A I RSEAL V A N E SUP POR T S U P PO RT RA IL i ACTIV E CL EARAN C E CO NTR OL AIR MANIF O LD Figure 3.2.3-5 Turblne Blade Tip Seal Assembly

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\ 1 3 _._ OF l_J ' -_FI (;_ ; _ _'' 3.3 HIGH-PRESSURE TURBINE COMPONENT RIG DESIGN The high-pressurc turbine component rig was designed to confirm the aero- dynamic performance of actual engine-type hardware prior to testing in the integrated c ore / low spool. A cross-sectional view of the rig is shown in Figure 3 .3-I. Basically, it consists of an inlet section, test section .qd exhaust section. The inlet, exhaust and outer case sections are rig hardware.

In contrast, the components in the test section are suitable for use in the integrated core / lo w spool. These compor.ent s have been designed to meet the structural requirements of the integrated core / low spool and contain nearly all the technology features of the Energy Efficient Engine high-pressure tur- bine component described in the preceding section. The main exceptions are the • use of a directionally solidified material for the vanes instead of a single crystal material and a metal,ic blade tip seal instead of ceramic material. In addition, since the rig was run at relatively low temperatures the airfoils were not coated. They were, however, restaggered closed O.3-degree to account for the coating thickness.

DRIVE FRONT BEARING REAR B E ARING COUPLING CO _ ; PARTMENT COMPARTMENT

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Figure 3.3-I High- P ressure T urbine C o mp o nent T est Rig The rig s econdary flow system was designed t o simulate the Energy Efficient Engine requirements. Air supply lines, as shown in Figure 3.3-2, supply meter- ed cooling air to the vane, t_ngential on-board injection (TOBI) system, bore cavity, mini tangential on-board injection system, and active clearance con- tr o l system. For e ac h metered flow, spe c ific coolant and leakage flow splits were calculated using pretest flow calibrations of rig hardware. In addition, a separate cooling supply system was provided for the active clearance control system. This permits the air temperature to be varied over an approximate 167°C (lf_(l°F) temperature range to facilitate clearance change.

Special consideration was given to the type of material used in certain areas • _ of the rig. For example, rig hardware exposed to main and secondary airflow was fabricated of stainless steel or comparable rust resistant alloy to pre- vent contamination of coolant passages HiGh strength materials were used _q high temperature regions of the rig, while less expensive low carbon steel was used for external rig hardware.

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MAIN TOBI VAh / E COOLING : _. Figure 3.3-2 Turbine Secondary F low System Air Supply Lines The rig has two main safety systems. The first is a dump valve, which bypasses the rig airflow upon detection of an overspeed condition or loss of bearing , oil. The second safety system is an alarm that is actuated when pre - established rig operating parameters are exceeded.

3.4 RIG FABRICATION AND ASSEMBLY Components for the test rig were manufactur ' . , d according to Pratt & Whitney Aircraft's standards for experimental test hardware. Figures 3.4-I through -3 show some of the major comporonts in the various phases of fabrication. Figure 3.4-] shews the finished tur o ine vane and case subassembly. The completed rotor subassembly, includi g the installation of the full ring, bolt]ess side- 4-2.

plates, is shown in Figuren3.

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J 15 w_- W Figure 3.4-] Completed Turbine Rig Static Structure with PWA ]422 Directlonally-SolidifiedTurbine Vanes A J L .... - " 4. _ - _ - ' , , ., ' ,r_ II' _ . ._

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Figure 3.4-3 Assembled T u rbine Test Rig Following fabricatio n , 11 parts were inspected to ens u re conformity to blue- print dimensions. Also, gas path and secondary flow system components were individually flow checked. Test hardware was generally in good agreement with the design specifications.A dimensior1 check of the blades showed that the throat flow a:ea was 4 percent less than the design. Measurements of the vane throat area showed the area to be 2 percent higher tha n desired.

An airflow check of che individual va n es showed that the cooling passage aver- age flow rate was 14 percent less than the design requirement, resulting from slightly smaller than intended holes in both the impingement insert and air- foil walls. The vanes were considered acceptable for rig testing since the in- let temperaturewould not exceed 426 ° C (800 ° F)and the cooling air exit veloc- ity was close to the design level. A flow check of the blade showed that the average cooling passage flow was within 1 percent of the design intent.

A cold flow calibration was conducted during assembly of the major rig com- ponents to ascertain specific secondary flow system coolant and leakage rates as well as provide baseline data for subsequent comparison of test results.

The vane inner platform area leakage was approximately two times the predicted flow, while the outer platform area overflowed by approxi n ately 1.5 times the predicted level. The vane subassembly cooling airflow was 15 percent under prediction,which confirmed results of airflow tests of individua_ vanes.

17 _ , _.

The active clearance control system showed a 20 percent higher flow rate at a 2.0 pressure ratio. The blade outer air seal cooling air supply holes over- flowed by 15 percent at a pressure ratio of 1.5. Four of these holes were plugged to reduce flow to the design intent.

The blade tangential on-board injection nozzle had a 4 percent higher than predicteH flow, but was judged to be acceptable for testing. The mini tangen- tial on-board injection nozzle, whirh swirls air in front of the turbine disk, had a flow rate 18 percent higher than the design intent. As a result, eight holes were plugged to reduce the flow rate.

Cooling airflow to the blades was only slightly underflowing the prediction and was in good agreement with the bench cooling flow results. Leakage past the attachment area of the blade was two times higher than the p,ediction.

However, this result was expected since the rotor was stationary and, there- fore, not encountering the normal centrifugal loads that would effectively seal the sideplates and damper seals.

The assembled test rig is shown in Figure 3.4-3 prior to shipment to the test stand. The riq was delivered to the Pratt & Whitney Aircraft Andrew Willgoos Laboratory in East Hartford, Connecticut in mid-February 1982.

At the test facility, all flows _nd clearances were rechecked. Flow ratios were in excellent agreement with the preceding calibration r_zults. This in- formation provided baseline data for subsequent comparison of test results.

3.5 TEST INSTRU E NTATION Test instrumentation, including type of sensor, quantity and location, was = determined on the basis of analyses and previous turbine test experience. As shown by the instrumentation map in Figure 3.5-I, the rig was equipped with sufficient instrumentationto monitor and record turbine aerodynamic behavior.

In certain areas, instrumentation was provided for redundant measurements to enable supervisory control of test conditions, establish data validity, en- hance measurement precision, and minimize down time resulting from sensor failures.

3.5.1 Performance Instrumentation A complete list of the flowpath instrumentation used in both the full stage and cascarJetests is contained in Table 3.5.1-!. The inlet was instrumented with stationary pressure and temperature instrumentation rakes to properly establish ri g inlet conditions. Three vanes were instrumented at the locations identified in Table 3.5.1-I for the purpose of determining pressure distribu- tions and cooling film effectiveness.

Laser proximity probes were used at four circumferential locations to measure blade tip clearance at different operating conditions during testing of the full stage. This system consists of e helium-neon laser, fiber optics and readout reticle, video camera, and monitoring / recordingequipment.

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.umb e r o f s e ns o r s ( n u _r_ c ai ) F O L DOUT F RA ME I_1 T y pe o f Pr o be A - t it ingle # - rp m B - a c oustics P - t ot al preSSure O - ue t s ur e t en t (bla d e ttp , e tc .) $ - static pressure E - missions T - t a pe ra ture F - a t H low V - vi b r at ion G - st rai n g a ge Z - atscell a neous M - l et ll teme rat ure k Fi gure 3.5 -1 I ns tr umen t a ti on Ma p of H i g h- Pressu r e T urbi ne C om p one nt R i g 1 9 I N t..- _ U 4) I-- 4.) _ " _" I.-- _ '- I-- _ _ 4 _ _ *. n _ L n I.. i= I=L • _m_ , NN( _ UUU U I I I I I I I I I I I I I I I I I _ I v tt n i - i ..-. _ i . n _ XXXX X)<XXX: _ '<XXX: _ 'EXXXXX XXX XXXXXXX 4-) W U g --- 20 [: J

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t i Circumferential traversing instrumentation was used to record turbine exit temperature, pressure and air angle. The traversing instrumentation ring contained four kiel-head total pressure pole rakes, four kiel-head total temperature rakes and four radially traversing air angle wedge probes. The locations of these probes are shown in Figure 3.5.1-I. The total range of the circumferential travel for the ring was 30 degrees, corresponding to two vane pitches. The circumferential positions of the rakes and wedge probes were selected to provide performance data in four quadrants of the rig to account for circumferentialvariations in vane hardware and to permit blade tip clear- ances to be properly documented.

AA - PROBE @ 33 0 ° PT RAKE @ 340 ° T ' I - - RAKE @ 3 50 ° , _- " _ TRAV TR A V E RS E R A NGE AA P ROB E @ 6 0 ° T T RA KE @ 7 0 ° P T - R A KE @ 8 0 ° F, AA - PRO B E @ 240 ° P T - RAKE @ 250 ° TT - R A KE @ 260 ° A A PROBE @ 1 50 = TT - RAKE @ 160 ° P T - RAKE @ 1 70 ° Figure 3.5.1-I Circumferential Traverse Rake with Instrumentation Locations and Traverse Path Static pressure taps were installed at three radial locations on selected vanes to determine the pressure distribution. Also, pressure taps were in- stalled on the endwalls to provide information pertaining to blade and vane exit flow conditions.

A summary of instrumentation in the secondary flow system is contained in Table 3.5.l-II. A part of this instrumentationwas used to record the thermo- dynamic states of the secondary flows in order to bookkeep turbine efficiency.

The remaining instrumentation was used to record secondary system flows and thermal performance. For cascade testing, only pressure and temperature sen- " _ sors in the vane inner and outer supply cavities were operative.

_ X Table 3.5.1-111 presents a summary of the flow measurement instrumentation, and Table 3.5.l-IV identifies the location and type of instrumentation for : recording rotor speed and humidity.

3.5.2 Structu , "alIntegrity Instrumentation To ensure the structural integrity of the test rig, sensors were installed to record vibration as well as bearing temperature and oil pressure. This instru- , mentation is listed in Table 3.5.2-I.

In addition, several stationary components in the turbine test section were _i instrumented with strain gages as a precaution against a potential fatigue failure. The location and quantity of the strain gage instrumentation is contained in Table 3.5.2-II.

3.5.3 InstrumentationCalibration and Accuracy ' ii Test instrumentationwas calibrated, as appropriate, to ensure a high level of , _l measurement accuracy. Five types of calibrations were conducted. These : i included: _ (1) Calibrator curves generated at Pratt & Whitney Aircraft and trace- i;! able to the National Bureau of Standards for voltage, pressure and ice point reference units.

(2) Thermocouple wire electromotive force correction curves generated at " Pratt & Whitney Aircraft and traceable to the National Bureau of Standards.

(3) Aerodynamic recovery curves for flowpath mounted total temperature and total pressure rake kiel head sensors and wedge probes (estab- lished in the Pratt & Whitney Aircraft Free Air Jet Test Facility).

(4) Choked venturi discharge flow coefficient curves established at the " Colorado Experimental Engineering Station and traceable to the National Bureau of Standards.

(5) Scanivalve pressure transducer curves which were updated during each performance acquisition by a secondary calibration employing dead weight tests traceable to the National Bureau of Standards.

, TABLE 3.5.l-lV SPEED, HUMIDITY ANDVIBRATION INSTRUMENTATION Te s t Mea s urem e nt T o tal L oc ati o n CascadeR o tating Type . _ MainShaf t Sp eed Es tabli sh Rl g S p eed X N - E le ctr o magnetic P i ckup 3 R igI nle t D e wC ell X TA- D e w C e ll Air Te m pe ra tu r e 1 Es ta b lish I n let Hu m ldlty X T _ . Dew Cell L ICI T e mp e ra t ur e 1 : X P ; - De w CellStati c P r essu re 1 TOBISu ppl y DewCe ll J v T A -De w C e ll A lr T em pe r a t u r e l Esta bli sh TO BI H u mi d l * C / ,-_ X T V -De w Cell LtCI Te m per a ture l X PS , De w C e l l Stati c Pr essu r e I ;_ 24 \

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I,.- XX XX _X X X X X XX XX XX X X X X_ X X *P _ _ " _ _ 0 _ , _ _ 0 _ _ _ U TABLE 3.5.2-I RIG SAFETY SYSTEM X-203 STAND ANNUNCIATOR AND WARNING LIGHT IN CONTROL ROOM WHEN LIMIT EXCEEDED Parameter O uantity ; External Vibration Front Case Horizontal I 0.005 cm (0.002 in) Front Case Vertical l 0.005 cm (0.002 in) Rear Case Horizontal l 0.005 cm (0.002 in) Rear Case Vertical 1 0.005 cm (0.002 in) Internal Vibration Front Bearing Horizontal l 0.005 cm (0.002 in) Front Bearing Vertical 1 0.005 cm (0.002 in) Rear Bearing Horizontal l 0.005 cm (0.002 in) i Rear Bearing Vertical I 0.005 cm (0.002 in) Rig O verspeed l 9500 RPM* Thrust Balance Cavity Pressure 1 206,844 Pa (30 psi) Rig Low Oil Pressure Front* l 241,318 Pa (35 psig) _ Rig Low Oil Pressure Rear* 1 241,318 Pa (35 psig) Rig Oil Out High Temperature 1 776 ° C (350°F) Rig Oil In Low Temperature 1 71 ° C (160°F) Rig Oil Filter Differential Pressure 1 34,474 Pa (5 psid) Front Compa,_tment 1 34,474 Pa (5 psid) Rear Compartment l 34,474 Pa (5 psid) Rig 0ii Tank Low Level 2 Low Level Rig Oil Tank High Level 2 r_ighLevel Rig Bearing Temperature Front Bearing 1 176 ° C (350 ° F) Rear Bea_'ing l 176 ° C (350 ° F) Loss of Rig Oil Scavenge 2 68,948 Pa (lO psia) Main Stream Filter Differential Pressure 2 34,474 Pa (5 psid) Low Bearing Differential Pressure Front 2 13,78 9 Pa (2 psid) Low Bearing Differential , P r ess ure Rear 2 1 3 ,789 Pa (2 psid) Active Clearance Control Cal Rod Heat Exchanger 2 i i * Over bo ard dump val v e ; nitiatedat 9700 RPM • 2 6 w_t TABLE 3.5.2-11 T I_ I f T TURBINE CO M PO_IENT STI_,UCTURAL INTEGRITY ,,,STRUMENIATION Location Quantity (I) Midpoint aft bearing center- Two sets of 0 / 45 / 90 degree rosettes line spring 180-degreeapart (2) Carbon seal support ring Two gages with axial grids close to aft fillet • (3) Inner aft turbine seal Two radial grids (4) Thrust balance seal Three equally spaced circumferential grids on the seal land (5) High-pressurecompressor Two radial grids on forward portion discharge seal of seal land (6) High-pressurecompressor Two radial grids d ischarge seal support cone (7) Inlet probes Two radial grids at center of flat : portion of inlet temperature probes close to the outer attachment and two grids on the pressure probes i (8) Inner vane outer spring seal Two axial grids i Table 3.5.3-I presents the error / accuracy of the instrumentation in test stand 1 X-203 that was used to conduct the test program. The analysis included the precision and bias associated with both the hardware and related calibrations used in establishing the engineering unit value for each parameter.

TABLE 3.5.3-I INSTRUMENTATION ACCURACY (95 Percent Confidence Level) Parameter Bias* Precision* Pressure, % + 0.05 + 0.10 (% of transducer range!

Temperature, °C (°F) T0.23 (+ 0.4l) _-0.21 (+ 0.38) Air Angle, deg T0.22 - T 0.65 - Tip Clearance, cm (in) Negligible _0.002 (+ O.OOl) Air Flow, % + 0.20 T 0.30 (%-of indicated flow) Radial Position, cm (in) T0.0063 (+ 0.0025) ¥0.0088 (-0.0035) Circumferential T O.Ol - +-0.04 Position, deg - Speed, rpm Negligible + 5 m * 2 standard deviations J 27 _ t A standard error propagation analysis for key performance parameters p_ovided the absolutp uncertainties (precision p!us bias) shown in Table 3.5.3-II at the 95 percent confidence level (2 standard deviatlo;,s).

TABLE 3.5.3-11 MEASUREMENT UNCERTAINTY OF TURBINE PARAMETERS (°5 Percent Confidence Level) Efficiency,% + 0.38 Pressure Ratio T 0.04 Speed Parameter, ¥ 0.16 ( N / V _r ) L 3.5.4 Data Acquisition System _ _j Test data were acquired with the Computerized High Accuracy Data (CHAD) system, ; _ located at the X-203 test stand. This system is designed to obtain pressure, temperature, speed, and spatial position data from both stationary and traver- t sing instrumentation. Traversing instrumentation is regulated by a Computerized • Automatic Traversing System (CATS) that controls probe movement through a digi- tal fe e dba c k loop. On c e th e computer lo c at e s the prob e s, position data are ac- quired from radial and circumferential digital shaft encoders and yaw balance data from analog differential pressure transducers. The computer for the CHAD system acts as a master control station, issuing commands not only for its own internal acquisition hardware, but also to the automatic traversing system. Scan lists _re incorporated into the data acquisition s y stem to define the order in which rig instrumentationis sampled.

SECTION4.0 • TTT TEST PROGRAM A,_ID TEST FACI-_ I r_ , L.)

4.1 TEST PROGRAM The principal objective of the High-Pressure Turbine Component Riq Test Pro- gram was to substaqtiate turbine design point performance as well as determine off-design perfoFm:'.nce. Oth_ test objectives included evaluating performance _.,Ining secondary system sensitivity sensitivity to bl_.de tip clearances and -_ to tangential on-board injection (TOBI), mini TOBI and compressor discharge seal flow rate variations.

To accomplish these objectives, the test pro_am was organized into two parts.

The first part consisted of full stage testing for c,_aracterization of turbine and secondary flow system performance. The f, 11 stage t.=st was structured into the five phases indicated by the test ma_ x i_ Table 4.1-I. Figure 4.1-1 shows an envelope of the test points along with the corresponding operating conditions of the test rig.

Vane cascade testing, the second part of the program, was directed toward eva!,,ating turbine vane aerodynamic performance over a range of exit Mach numbers and Reynolds numbers for insight in the analysis of ful3 stage perfor- mance. The test matrix is presented in Table 4.1-11.

The test program, consisting of rig installation, testing, and data analysls, was a nine-month effort. A total 216 hours of testing was conducted. Perfor- mance data were acquired for 13 principal test conditions _ur,ng the full stage rig test, covering the high and intermediate power range of the :_te- grated cor e/ low spool.

4.1.I Full Stage Turbine Test Before actual testing was initiated, a series of pretest checks was conducted to (l) verify the proper operation of all test support systems and equipment and (2) demonstrate the mechanical integrity of the test rig. Checks of the instrumentation and data acquisition systems were performed both during rig assembly and after it was installed in the test facility to ensure proper ; operation. Also, a pretest shakedown of the facility was conducted to check out the r,g supervisory cor.trol and each secondary flow system conlant line.

Finally, a shakedown test of the turbine rig was accomplished to verify both the mechanical integrity of the rig and proper functioning of rig instrumenta- tion.

The full stage test, as indicated by the test matrix in Table 4.!-I . consisted of five phas_'s. The first was directed toward substantiating turbine design point performance as well as determining off-design performance. Secondary cooling flows were set at the values listed in Table 4.1-I, and an average blade tip clearance of 0.0472 cm (0.0186 in) was achieved by regulating th_ • active clearance control air temperature.

In the next phase, an assessment of performance sensitivity to blade tip clearance was made by operating the turbine rig at tlu design condition and varying the blade tip clearance. The desired clearance_ were obtained by re- gulating the temperature of the active clearance control air. Clearances were _L measured by four laser proximity probes and traversing data provided a ba_!s \ for identifyingperforr,_nce changes, 30 i , _

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Vane* A c ti v e * Tu r bin e and B u ff e r* R eJ r , C]ear J nc e _ Pl l n TTin PT ou I I nl et Flo w S _ 'ed Torque B O A S Mt nI - * Bo r e, B r 9. Dis k C ont r ol _'t . w , ( p s ,,) K _ - r ) NP , tpst , I _ _tj k _ ' s e c(l_ / s e (I ( R_ . ) ( n - L e d H .P . C o o l , n t t O _f . TO el C o o l,n t C,. lty T rl . (A( C ) D .P. 0 441 7 0 0 I I0 4 72 1 4. 2 R6 _ n _ T _O q2 1 0 _l _ 3 3 . _ 5 U. _V 0 .4 7 0,2 3 0,1 2 O.2B (64 O) (1 26 01 (16.n ) ( 8 5 01 (3 1 .4 J B 0 .55 2 7 0 0 1 1 0 450 1 7 . B 86 60 3 5 9 0 593 0 ( 8 0.0 1 ( 12 6 0 ) ( ]6 .01 (8 10 1 ( 3 9 .3J C 0 44 1 7 0 0 1 1 0 4 7Z | 4 . 2 7 8 _0 2 800 4210 (64 , 0 1 (I 2 DO ) (l& . O ) ( 850 ) ( 3 1 .41 D 353 7 00 l l O 49 7 1} .3 8660 1750 289 0 ( 5 1 . 2 1 ( 12 60) ( 16. 0 1 (895; ( 2 5. 1 1 A

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220_ C .. I 1 i .... __ L __ _I 3 0 _. 0 5 , 0 PRESSURE RATI O (PTIN / PTExIT) *_ te t e redS eco ndary F low s - % o f HP T I nle t F lo w Figure 4.1- I Test Env e l o pe for Ful l Stage Turbin e Test Program The influence of both the blade and mini tangentia l on-board injection (TOBI) systems was evaluated duri , g the third phase of the program as part of the overal l performance characterizationof the secondary f l ow system. An eva l ua- tion of the b l ade f l ow sensitivity provided pertinent infcrmation to assess the inf l uence of b l ade off-design cooling f l ows on turbine performance as we l l _s TOBI performance. Next, the sensitivity of the secondary coo l ing system to variations in the mini tangential on-board injection flow was investigated.

Testing included operation with no mini tangentia l on-board injection flow.

The fourth phase of the program was aimed at providing additional mapping of turbine off-desig r ,performance. Testing was conduc t ed at more exaggerated off-design c o n d itions in comparison t o t ho se in t h e initial ph ase o f th e program.

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• TABLE 4.1-11 TURBINE VANEANNULAR CASCADE TEST CONDITIONS PHASE i - Mach NumberEffects (Ambient Discharge) Exit Reynolds Test Ma ch TTi n Number Cooli ng Point Number °C (°F) (105) PR TR I 0.70 426 (800) 1.97 1,027 0.48 2 0.85 426 (800) 2.56 1,027 0.48 3 0.91 426 (800) 2.82 1.027 0.48 " 4 l.O0 426 (800) 3.24 1,027 0.48 5 1.08 426 (800) 3.62 1,027 0.48 6 1,16 426 (800) 3,92 1.027 0,48 7 0.56 426 (800) 1.52 ].027 0.48 PHASE 11 - ReynoldsNumber Effects- High-Pressure Discharge :i :_ 8 0,92 426 (800) 4,66 1,027 0.48 _ 9 0.92 426 (800) 6,44 1.027 0.48 , g The final seriesof tests evaluatedthe influence of compressordischarge seal flow rates on the pressuredistributionin the front disk cavity and blade _i TOBI supplyplenumas well as the resultant effectson bladecoolantflow, 4.1.2 Vane CascadeTest The annularcascaderig -_'s_entially consistsof all the compo n entsin the f_J]1 stagerig, except that the turbinerotorwa,_removed.With the removalof the . _ rotor, the exit _raversinginstrumentation system was positionedbehind the vane stage to obtair aerodynamicmeasurementsat different circumferential locations.

Variouspretestchecks were conductedto verify the proper operationof the C test rig and supportsystems.These checks followedthe same format as those performedin the full stagerig program.

The vane cascade te s t consis t edof two phases. The first phase of testing I evaluatedvane performance under varyingexit Mach number conditions with an

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ambientdischarge.The range of Mach numbers tested (0.56 - 1.18) bracketed q the design pointfor the integrated co_e / 1owspoolhigh-pressure turbine.Test _ ' point 3 (Table 4,l-II)correspondsto design point conditionsor test point I No. 1 in the fu I l stage rig test matrix.Data acquiredat this condition ;'ere i usedwith the full stagedata to ,_ssess stagegas path aerody n _ =mics.

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In the second phase, the rig was back pressuredto evaluatevane performance at higher Reynolds numbers.

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4.2 TEST FACILITY • Test stand X-203, located at the Pratt & Whitney Aircraft Andrew Willgoos Tur- bine Labor-fury in East Hartford, Connecticut, was used for the Turbine Rig f Test Program. The test facility is shown in Figure 4.2-I, and is designed for development testing of full scale turbine component rigs.

The test stand consists of a large enclosed test cell and an adjacent control room. The component is mounted on an open bedplate, sealed to the air supply and exhaust ducts, and connected to the dynamometers by a suitable drive shaft through a gearbox. Air can be supplied at a rate up to 56 kg / sec (125 Ib / sec) i at 635 cm (250 in) Hg absolute. High pressure air can be heated to 537°C (IO00°F) by two natural gas-fired heaters. Through the co n trolled use of laboratory compressors, a refrigeration system, air heaters, and exhauster units, sea level and altitude operating conditions can be simulated over a wide range.

Two eddy-current dynamometers, each with a I0,000 horsepower absorption capa- :: c ity, are c onn ec ted to t he c o m ponent t h ro u gh a gearbox to reduce the output _ speed of the test unit to a level acceptable for the dynamometers. Only one ' dynamometer was used for this turbine rig test. Auxiliary equipment used to _ support testing included a remotely controlled lubrication system and a natural _I gas supply system. The stand is also equipped with an extensive safety and • _ fire protection system.

Stand services available to operate equipment in support o_ testing include 689,480 and 2,413,180 Pa (I00 and 350 psig) air supply, !20 and 480 volt, 60 cycle per second A.C. and 24 volt D.C. electrical power, 206,844 and 2,999,238 Pa (30 and 435 psig) steam supply, and city or river water supply.

All controls and instrumentation required to operate the test rig and monitor performanc_ are located in the adjacent control room. The test operator has direct control of the test vehicle, inlet and discharge air valving associated with the test stand and the power absorbing dynamometers.

During the test, a Rig Supervisory Control system was used to aid in setting performance points, maintaining test parameters within preprogrammed limits, and monitoring ri.g and facility safety parameters for early detection of prob- lems. The supervlsory control is a digital computer control with appropriate input and output signal circuiting. In addition to the Rig Supervisory Con- trol, certain additional monitoring systems were used to ensure both rig and facility safety.

4.3 DAT# REDUCTION AND ANALYSIS The analysis of data obtained from this program was accomplished in two steps.

The first step assessed the validity of the data and executed preliminary cal- culations to obtain turbine performance information. This was conducted in near real time.

The second step consisted of more comprehensive calculations to obtain final turbine performance values. In addition, the test results were compared to Design System calculations. This work was completed shortly after the comple- tion o f t es ting.

- _' 33 , " 3F PCt_ t ,:., _ , Figure 4.2-I Pratt & Whitney Aircraft X-203 Test Facility. (The top photograph shows the internal ducting as well as the location of the test component. The bottom photograph shows a typical test component without the ducting installed.) \ 34 _'_ ._ _. _,_'_ : " Data acquired at the test stand by a Honeywell 316 computer were sent to a Univac II00 computer, which converted raw data (electrical voltages) to engi- neering units. The engineering unit data sets were then sent to the iBM com- puter network for review and analysis in near real time. The culmination of the near real time effort was the execution of preliminary turbine efficiency and secondary system performance calculations employing existing computer programs, 4.3.1 Analysis of Full Stage Data Data obtained during testing were analyzed in detail to determine turbine and secondary system performance at design and off-design conditions. For each 1 test point, the thermodynamic efficiency of the component was calculated in ! accordance with the following definition.

turbine efficiency = _]i mi Ahi _E]i mi Z_hi ' . mi - mass flow of i th stream Ah i - actual enthalpy change of i th stream Z_h i ' ideal enthalpy change expanding flow to turbine exit pressure Where the numerator and denominator are summed over the primary and all cool- i ant and leakage flows, ,i C,_mparisons of test data with design system nredictions were made in the j_ following areas: f - turbine efficiency - turbine flow capacity (flow paran_ter) - stagP reaction - turbine exit air angle profile - turbine, Mach number profile - secondary system performance (pressures and temperatures) - active clearance control (ACC) response - vane surface pressure distributions (design point only) - vane suction surface cooling effectiveness (design point only) i 4.3.2 Analysis of Cascade Data For each test poi n t, comparisons of data with design system predictions were made in the following areas: - full passage total pressure loss - exit air angle profile exit Mach number profi]e vane flow parameter - vane surface pressure distributions (design point only)

i

SECTION5.0 TEST RESULTSAND ANALYSIS 5.1 INTRODUCTION i This section presents the results of the turbine rig test program. The mechan- ical performance of the turbine rig is discussed in the following section, and Section 5.3 presents the results and analysis of turbine aerodynamic perfor- mance. The presentation of aerodynamic performance results is divided into three categories. First, stage performance is discussed. Second, vane cascade performance and supporting data are presented. Third, the results from the vane cascade testing are used with results from stage testing for a character- ization of turbine blade performance. Finally, Section 5.4 presents the re- sults of a post-test inspection of the turbine rig hardware.

5.2 MECHANICAL PERFORMANCE The test rig was successfully operated over a range of conditions that simula- _ ted operation of the integrated core / low spool at both intermediate and hiuh " power levels. No major difficulties were encountered that prevented the acqui- sition of performance data. Thrvughout the test, the mechanical performance of . the turbine test section components -- the turbine gas path and secondary flow system components -- was flawless. This was confirmed by a post-test inspec- :_ tion. Component stress levels were low, on the order of 13.8 MPa (2000 psi), with no significant change during the course of the test.

During the initial shakedown of the rig, excessive bearing compartment and facility dynametric coupling shaft vibrations were encountered at mechanical speeds above 8300 rpm. Also, the exit circumferential instrumentation traverse :.

ring experienced bin,lingat temperatures above 315°C (600°F). To account for this, the maximum rig operating temperature was reduced from 426 to 315°C (800 to GOO°F), the sp_ed paraamter wa s held, and the air density ratio was in- creased from 0.48 to 0.52 at the design condition.

In addition, blade tip clearance measurements disclosed an eccentricity or displacement of the rotor / case assemb,y. As indicated in Table 5.2-I, clear- ances ran_ed from 0.020 cm (0.008 in) in quadrant l to as large as 0.078 cm (0.031 inj in quadrants 3 and 4. This condltion, which was attributed in part to tolerance differences in the experimental hardware, somewhat restricted the blade clearance investigatir planned in Phase II of the program. Selected clearance variations were ac h , red by regulating both the rig inlet tempera- ture to reduce rotor growth and the active clearance control air temperature.

TABLE 5.2-I TURBINE BLADE TIP CLEARANCES Location (De_) Clearance Readin_ cm (in) 42 0.020 (0.008) 132 0.033 (0.013) 222 0.078 (0.031) , 342 0.378 (0.03l) For the analysis of performance, an average tip clearance of 0.048 cm (0.019 , in) is used, and efficiency measurements have been adjusted accordingly. \ _C 37 ,_ 5.3 AERODYNAMI C PERFORM A NCE 5.3.1 Turbine Stage Performance As s essment The turbine rig overall performance parameters are listed in Table 5.3.1-I.

These nondimensional parameters show a relatively close modeling of the in- tegrated core / low spool design at the rig design point. At this point, rig i performance exceeded all efficiency goals and closely matched the design level _ of flow capacity and reaction. As indicated in Table 5.3.1-I, the measured flow capacity was 17.2 compared to a design value of 17.0 and the stage pres- : sure reaction was 43.1 percent compared to the design level of 42.4 percent.

The measured efficiency of 88.5 percent surpassed the rig efficiency goal of i 86.5 p e rcent. Moreover, it surpassed the expected test efficiency for the in- tegrated core / low spool (87.1 percent) and the goal for the flight propulsion system (88.2 percent). It is also noteworthy to compare this efficiency to the level measured in the preceding Uncooled Turbine Rig Program, which was 91.1 percent. Thus, a deficit of 2._ percent is imparted by the introduction of component cooling and the functional operation of the secondary flow system. , TABLE 5.3.1-I _| COMPARISON OF PERFORMANCE PARAMETERS INTEGRATED CORE / LOW TURBINE I SPOOL DESIGN RIG ! Mean Velocity Ratio 0.551 0.557 Work Factor ].65 1.61 Cx / U 0.360 0.332 : E x pansi on Ratio 4. 08 4 4. 0 14 Inlet Fl o w P arameter 17.02 3 17. 1 9 1 E_it Flow Parameter f q .165 65.215 Cooling Flow, % Wae 14.56 !2.36 Pressure Reaction,% 42.4 43.l Exit Mach Number 0.539 0.519 Exit Angle, deg 44.0 49.8 Ah / TT RIT 0.685 0.675 TT RIT, K ( O R) 1550 (2791) 566 (lOlg) 5.3. 1 . 1 C o mp o n e nt E fficiency and Reacti o n Test data were processed to determine the mass-averaged cooled turbine effi- ciency at design and off-design conditions evaluated in Phases I and IV of the program. The test conditions for these Phases are s l lown in Table 5.3. 1 -II along with the measured clearance, mass-averaged efficiency, mass-averaged efficiency adjusted to a clearance of 0.048 cm (0.019 in) _nd pressure r e - action. The design point was repeated in Phase IV of the progra m , as shown in Tab le 5.3.1-II. Resu l ts are in excellent agreement with the first design point F test and verify the leve l of e ffi c iency. In addi_ on, the results show that th e r e wa s n o p e rformance deterioration during the test.

38 ._ 0% ,_I" ¢_ O0 - "_I" 0 N' 4 P " " :c;, ORrc ' 'JAL _' ; ;_t , * ' : ES u E . • ° • * • • ° • • • • , "" -_ _ OF POOR QUALI T Y ' _ j ,_ - OF PO0_ , ; ...... " Figure 5.3.l-I shows turbin_efficiency trends as a function of pressure ratio and speed parameter (N / V-T-T),These results show efficiency at the design speed parameter of 244 increasing with pressure ratio to the design level of 4.0 and then decreasing. Efficiency also increases with speed pa r ameter at a given pressure ratio, This trend is more apparent in Figure 5.3.]-2, which is a cross plot of Figure 5.3. l -I with speed parameter as the primary variable.

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34 3 6 3B 40 42 44 46 4B P RE SS URE RA _O Figure 5.3.1-] Turbine Stage Efficiency Trends as a Function of Pressure Ratio and Speed Parameter P _ E S SURE RAT I O 0 , o 0 43 89 • ' p- w B E 220 2 24 2 2 B 219 23 B 240 244 2 q 25 2 S P EED PARAMETER , N / V _"T Figure 5.3. 1 - 2 Tur b ine Stage Efficiency \ " 40 ,i_ o | ,, Turbine reaction, which is defined as the ratio of the static pressure drop across the rotor to the static pressure drop across the stage, is presented in Figure 5.3.1-3. The static pressure reaction increases smoothly with pressure ratio through the design point level of 43.1 percent.

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, , 3 2 3 4 3 8 3. 1 4 0 4 ._ 44 4 6 41 PRESSURE RATIO + Fiqure 5 , 3.1-3 Turbine Reaction Characterist+cs 5.3.1.1.1 Blad e Tip Cl e aranc e Adjustment As discussed earlier, blade tip clearances showed a variation at the measuring locations. Since efficiency was not measured at the same locations as the blade tip clearances (refer to Figure 5.3.1-7), the average clearance was used to correct the average efficienc#. The sensitivity of turbine efficiency to tip clearance was determined on the basis of results of Phase If. This sensi- tivity was used to correct the efficiency _t the measured clearance to the design clearance o f 0.048 cm (0.019 i n ). During this test, two tip clearances : were evaluated at the design point conditions using the active clearance con- trol system, The results are presented in Table 5,3,1-III and sh o w the clear- !

ance change with the attendant change in efficiemcy for quadrants one through three. (Data frum the fourth quadrant were not used since that quadrant was partially affected by instrumented vanes and did not exhibit consistent results.) The clearan:e adjustm e nt factor for correcting efficiency was the average of quadrants one through three. The average c l earance adjustment factor is a 0,09 percent increase in efficiency for a 0.002 cm (0.0 0 1 in) de- crease in clearance.

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I i I "% TABLE 5.3.1-111 OF poor< Q,ji_Li ; _ f CLEARANrE ADJUSFMENT Quadrant Quadrant Quad r ant 1 2 3 Z_Clearance, cm (in) 0.015 (0.006) 0.020 (C.OOC) 0.015 (O.OG6) Z_% 0.61 0.61 0. 1 4 A n/ Z S clearance = 0.09% / mii tip clearance Reference: Z_TI / _clearance = O.08 , = '/ mil tip clearance uncooled rig 5.3.1.1.2 Secondary Flow System Performance All major components in the secondary flow system, which include the blade and mini tangential on-board injection systems, blade outer air seal, and com- pressor discharge seal, performed according to the design intent. The map of the secondary flow system shown in Figure S.3.1-4 compares measured flow rates and pressure levels with the prediction for the design condition. An excellent a g reement between the prediction and test d_t_ i_ apparent. With this good data correlation, corrections to performance data are not necessary to adjust for anomalies in the secondary flow _ystem.

In addition, as discussed in Sectio n 5. 3 .4, results of characterization studies showed thatjfor this turbine designasystem performance is relatively insensitive to variations in flow rates and pressures. Table 5.3.l-IV presents a summary of flow sensitivity tests, in which the effects of variations in flow rate to the r,ain and mini tangential on-board injection nozzles and vane case were evaluated. These results indicate a relatively small potential error in efficiency for the cooling flow variabl_ tolerance maintained during tt'e test. The cooling flow parameters are listed in able 5.3.1-V for each of tne i test points.

T ; BLE 5.3.l-lV SECONDARY FLOW SYSTEM FLOW SENSITIVITY TEST RESULTS Test Pressure Ratio Point No. luentification Main Tobi Mini Tobi Vane Case Efficienc), Deslgn Intent 0.946 1.027 1.027 l Design Point .0_942 1.030 1.027 Base 18 +23% Main TOBI l-._--_ l.C , 2l ].030 -0.22% Air Flow 21 +l_ Mini TOBI 0.949 _ 1.025 -0.10% Air Flow 28 -9% Vane Case 0.932 1.028 _ -0.14% Air Flow Design Point Error Z_ n % +0.008 -0.002 0.000 42 \ Figure 5.3.I- 4 Turbine Secondary Flow System Map Showing Predictea and Measured Flow Rates and Prp.ssures 9 43 ,_.

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. _ _ , o d oooo o o o o _ , ooo o o d _ c _ c _ c; d _ . c _ oo _ ( x _ lx J O _ C; _ O_ 00 C .) O _ 0 tO _ I _ '_ O00C_O0 0 0 O0 . _ 0 _ 0 00 0 "_ 00 0 C 3 0 00 o_ ' _ ooo dd o dd oo _ oooooo > _ _ dd o ooo h- • • • • • • • • • • • • • • • • • • " _ C_ C _ O C _ C_ C_ O _ (:30 0 00000 0 000 00 0 000000 0 C ) O 00, _ 1" t _ _ 0 0 000_ 0 O0 • • , • • * • * • * • * • • • , • • • • , , • • • • • , • • • . • • • , • I . -- OOC _ O 0 000 O0 0000 00 0 00 0 00_ mmm , , , • , • • • • • , • • • • • • , • • • • • 44 \ O RI ; OF F i3L 5.3.1.2 Design Point PerformanceCharacterization The following sections present the turbine rig inlet pressure and temperature profiles; exit pressure, temperature, and air angle profiles; and ca]culated efficiency and Mach number profiles. Also, the rig exit Mach number and air angle profiles are compared to the design intent.

5.3.1.2.1 inlet Profiles The inlet total pressure and total temperature profiles are shown in Figures 5.3.1-5 and -6. Figure 5.3.1-5 presents the spanwise total pressure profile at four circumferential locations, and shows a spanwise variation of only 1.5 cm (0.6 in) of mercury and a circumferentialvariation of 0.2 cm (O.l in) of mer- cury. The spanwise total temperature profile in Figure 5.3.1-6 shows the mid- span circumferential variation to be approximately 4 ° C (6 ° F) This figure also shows the lower temperature at the irner and outer walls resulting from heat transfer into the rig structure. These profiles show a consistent and well documented inlet for calculating t_rbine efficiency.

e = 3 41 ,C 0 QU A DRANT 1 [7 QUADRANT 2 _ QUADRANT 3 L_ ' /_ QUADRANT 4 "r 341 3 __ ( 1 346 i i :E U _- 340 , 8 1134.2 n (_ 3 49 , 3 I-- _I2 ' . , 01 P- _z _ 3 339 , 8 _ p_ (13381 u ) ( 1 3360 1 20 40 60 80 I(X) PERCENT SPAN Figure 5.3.1-5 Turbine Inlet Spanwise Total Pressure 5.3 . 1.2.2 Turbine Efficiency The design point turbine efficiency for the four quadrants defined in Figure 5.3.1-7 is prese n ted in Table 5.3.l-Vl and Figu , ' e5.3.1-8 as a function of span. The efficiency ievel shown is based on the spanwise measured inlet and exit temperature and pressure measurements. The resulting profile was adjusted to reflect the bulk cooling air effects, i.e., the efficiency profile was lowered such that the integrated average agreed with the cooled efficiency definition. These profiles show the maximum efficiency to be consistently in _ . t h e 30 to 50 percent span region.

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2 3 4 DEG _ . /'_ 2 7 3 DEG 304 0 3 1 6 DEG - 1 576 ) 0 20 4 0 6 8 8 0 1 00 PERCENT SPAN Figure 5.3.1-6 Turbine Inlet Spanwise Total Temperature • .

The profiles also sh o w the e fficiency change with circumf e rence. The blade tip clearances in Table 5.2-I show that the first quadrant has the tightest clear- ance. This is reflected in Figure 5 . 3.1-8 by the highest level of efficiency in the tip region. Similarly, quadrant three, which has the most open clear- ance, has the lowest outer wall efficiency. This tip clearance effect appears to extend toward the inner wall to approximately 60 percent span where effi- ciency in all four quadrants is within I percent. The data from these four quadrants were averaged to produce the profile in Figure 5.3.l-9.

An efficiency contour of the first quadrant is shown in Figure 5.3.l-10. This contour, which shows one vane gap, indicat ' . , s that the efficiency is relatively flat in the circumferential direction with the variation shown caused by the upstream vanes. Like the spanwise profile, the contour plot shows the maximum efficiency near m idspan and l o wer efficiency at b o th the inn e r and o uter walls.

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, A I R AN GLE LASER PRO B E 4 \ L ASE R CLE A RA N CE _1 PRO BE 1 | 0 ° PT & TT A IR ANGLE A NA LYSIS AN AL Y SIS REGION 270° 90 ° REGION QU A DRANT 3 Q UADRANT 1 PT & rT • AIR ANGLE 18 0 o LAS E R CLEARA N C E LASE R CLEARAN C E PRO BE 2 P RO BE 3 \ AIR ANGLE PT & TT AN A LYSIS REGION QUADRANT 2 Figure 5.3.1-7 Instrumentationand CircumferentialMeasurement L o cations (Looking forward at the rear of the rig) TABLE 5.3.1- V I DESIGN POINT EFFICIENCY Circumferential Efficiency at Measurement quadrant Location, deg Measurement C]earance, % Clearance, cm (in) l 42 0.020 (0.008) 80-95 89.13 2 132 0.033 (0.013) 170-185 89.17 3 2 22 0.081 (0.032) 260-275 87.0 4 4 312 0.078 (0.031) ' 355-20 88.12 Averag e 88.37 0.053 (0.021) 47 "_.i - . o OF PO_,?_ ' , " ', -"" ' -' ; ' 0 29 4 0 8 0 80 _ 00 0 20 40 00 _ ) 100 PERCENT SPAN i PERCENT S PAN Figure 5.3.1-8 Exit Spanwise Efficiency Figure 5.3.1-9 Av e rag e Spanwis e by Quadrant Efficiency

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C URVE LAB E L E FFIC IE NCY 2 78 8 9 2 3 79 4 3 0 4 79968 5 80.506 6 81 04 3 7 81 581 8 82 11 9 9 82.657 10 83.195 1 1 83.73 3 12 84.271 13 84.809 14 85348 VANE 15 85 884 W A KE 16 86 422 " LOCATION 17 86.960 18 87.498 19 88O36 _ 20 8 _ .574 21 8 _ 1112 22 89 649 2 3 9 0 187 24 90 725 ; _ 25 91.26 3 : 26 91 . 801 "_. _ 27 92 339 29 93 414 30 93 952 - =- - - '----'- " 15° TRAVERSE L _ 28 92 877 80 ° 9_o Figure5.3.1-I0 Efficiency ContourPlot of One Vane Gap in First Quadrant, Showing H aximumEfficiency Near the MidspanRegion 5.3.1.2.3 TurbineExit Aerodynamics A spanwiseprofileof the turbineexit total pressurefor al] four quadrants is presentedin Figure 5.3.1-II.All four quadrantshave the same general shapes.The total pressure Jr , creasesto approximately 70 percentspan and then decreases to 90 percentspanwhere the effectof tip clearancecausesthe total pressureto increasetoward the tip. Figure 5.3.1-12shows the total pressure cont o urover a 30-degreesectionof the circumference iN the first quadrant.

This figure exhibits the same spanwise profile seen in Figure 5.3.I-11 and also shows the effect of two upstreamvane wakes which cause th circumferen- tial pattern.

The turbineexit total temperature characteristics are presented as a spanwise profileof all four quadrants in Figure 5.3.1-13 and as a contGur plot for quadrantone in Figure 5.3.1-14.As shown in Figure 5.3.1-]3,a l l quadrants have consistent shapeswith temperature increasing from the inner to the outer wail. As with the total pressuredata, the temperaturc data show the effect of tip clearance.This is evidentby the data from the third quadrant,whe; -_. the largestclearance has a ll owedthe most hot gas to bypass the rotor and produce the highes_ temperatures. The contour plot of these temperaturesin Figure 5.3.1-14shows the same increasingtemperatures from the inner to the cuter wa l ls as th e " .p anwise plots. The c o nto u r p lo t a ls o s h o w s the e ffect o f th e coolingair from the upstreamvanes, This vane coolin§air di l utioncauses the _ re su l t ing circumferential patter n . \ ; " 4 9 '*_ , _ _ 3 F r _i , + , Q UADI _ A N T 2 Og_ C 4"_"" : r - _ ooAo , _ * ,

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t 32 4t b :. 132 0 ' ! _ ' _ 1 3 1 6 _ 2 t3 1 2 1 0 2 0 4 G 8 _ 8 0 I 00 f -;<" PER C ENT S PAN Figure5.3.]-!] SpanwiseProfileof TurbineExit Total Pressure BL A DE.r = xtT CJRVE TOTALP I IESSU rI E LABEL ¢m ll n l HG A 2 78 7 {31 Ol 3 711.9 {311) 4 795 I l l 3) 5 797 (31 41 (I 802 (3 _ _ 7 805 ( 31 71 Is ts _ o 1 3 1_l 9 812 1 32O i 10 II 1 7 1 3 22 1 I t 820 {323l 12 825 1 32 51 13 8 2 8 (326 1 ' 14 833 1 328 1 1 5 83S (3 29 1 16 d 4 0 QL_ 1 1 1 7 M 3 (3.1 21 18 1 4II (334 _ V N _ E _g 8 _ ,0 1 335 1 ,, WA K E 20 8 15 _t 1337 1 21 16II (33 8) 2 2 N31340 i 2 3 N 4 (34 1 1 24 87 ! 1 343) _ * 873 ( _ k m 41 2 _ 878 _ 46 1 27 M1 1 3471 211 Ill 3 _ 34 Ill 2 9 M ! ( 3 1, O I 3O I 1 1 3T _ I I I10w ]i= $0 Fig u re5.3.1- 7 2 TurbineE x it Tota l PressureC o ntourP l ot _ " i l ed I 1 3 101 C) QUADRAN T I O ( _ JADRAN T2 _ QUADRAN T 3 14 8 _ Q' U A D RANT4 _ 143 _ g 1 , , OF ; G' . :. : ' - r

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_ J 1 2 8 0) "_ ' ' ' | i I _ 1 21 no t2_l 1 1 0 _b PEFIC E N T SPA N I Fig u re 5.3. 1- 13 S panwi s e Pr o file of T u rbine Exit Total T e mperat u r e A sunlnary of turbine exit air angle data is presented in Figures 5.3.1-15 : through-17. Figure5.3.1-15shows a spanwiseplot of the four quadrants.The data trends indicatethat exit air angles in all quadrantsare generallycon- sistent, with the air angle becomingmore axial toward the outer wall. As with the pressureand L Emperature data, the effect of tip clearancecan be seen in " the air angle data. The largesttip clearancewould be expectedto have the air angle closest to the vane exit conditions.This is shown by the data, since quadrant three has the most axiai flow of the four quadrantsand the largesttip clearance.The effect o_ tip clearanceagain can be seen to per- sist toward the innerwa l l to approximately 6 C percentspan as the air angle plotsfor the four quadrants agreewithin4 degreesat this location.

The air anglecontourplot in Figure5.3. 1 -16,shows a 1 5-degree sectorof air angle data from a circumferential traverse.It shows the sans spanwisetrend as in the precedingfigure and a relativelyflat contour in the circumferen- tial direction.

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,t 52 _ - P E RCE N T SP AN F i gu r e 5.3.l -1 7 _v e ra ge S pan wi se _ir An gl e Tr ends Comp ar ed t o t he Des i gn P r ed i c ti on and P r eced i n g Unc o o l ed R i g Te st Resu lt s 53 . _k.4 O F PC ": I n F ig u r e 5 .3. 1 - 1 7 , t he ave r a ge spanwis e air angl e d a ta ar e co m p ared t o t he d e sign p redi c ti o n . As sh o wn , th e r e sults diff e r c onsid e r a bl y in sl o p e fr o m the pr e di c ti o n . In t e rms o f t he a e r o d y nami c im p a c t o n t he str u t f a iring in t he transition duct, this type of flowfield would c ause a more negative incidence from the root to approximately 40 percent span and a more positive incidence over the rest of the span. When comparing the design prediction to the results acquired from this test as well as the previous uncooled rig test (build 2), the -lopes at the midspan region are similar. However, the root and tip re- gions show different slopes than the prediction. At the root, the uncooled rig results show a steeper slope resulting from viscous three-dimensional effects.

The cooled rig results have a still steeper slope. This is attributed to the addition of cooling air, since the aerodynamics of the two rigs art similar.

In the tip region, the uncooled rig air angle starts to show a fall off simi- lar to the root region and then turns more axial because of the affect of the tip clearance. This trend is not duplicated by the results from the cooled rig test. The addition of cooling air and the tip clearance combine to cause the air to become progressivelymore axial from the midspan to the tip.

Figure 5.3.1-18 shows the t,,r_ineexit Mach number compared to the design pre- diction. Exit Mach number was calculated using the exit total pressure data and assuming a linear static pressure distribution between the inner and outer wall pressure taps at the instrumentationplane. The results, therefore, follow the exit total pressure distribution. The prediction has a considerably flat- ter profile, since it assumed a flat spanwise loss profile.

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0 20 4O 6 O S O 1 00 PER C E N T SPA N Figure 5.3.1-18 Blade Exit Mach Num b er Characteristics \ 54 wi_ i _" 5.3.2 Vane Cascade Performance Assessment Vane cascade testing was conducted at a series of Mach numbers that bracketed the vane exit Mach number in the full stage test. In addition, the cascade design point was run at three Reynolds number levels, with the highest at the stage test level. The overall results from this testing showed that the vane l oss was above the level for the uncoo l ed vane cascade and at the pre-test predicted level.

5 . 3.2.1 Pressure Loss Cascade data were acquired in two phases . The test conditions and results for both phases are presented in Table 5, 3 .2-I .

TABLE 5.3.2-I HIGH-PRESSURE TURBINE ANNULAR CASCADE TEST CONDITIONS AND RESULTS PHASE I - Mach Number Effects Exit Reyno]ds Pressure Test Mach Corrected PTin TTin Number Loss (%) : Point Number Flow MPa (psia) °C (°F) (105 ) Mass Avg ] 0.70 I5.58 0.146 (21.1 5 ) 316.0 (600.9) 1.97 3.06 2 0.85 16.67 0.173 (25.15) 314.0 (597.3) 2.56 4.49 3 0.91 16.94 0.187 (27.09) 314.6 (598.3) 2.82 5.16 4 1.00 17.05 0.212 (30.74 ) 313.5 (596.3 ) 3.2 4 6.54 5 1. 0 8 1 7 .22 0.23 7 134.36) 314.5 (598.1) 3.62 7.80 6 1.16 17.22 0.257 (37.25) 314.0 (597.2) 3.92 9.03 7 0.56 13.78 0.]29 (18.65) 3_$ ) {601.1) 1.52 2.06 PHASE I] - Reynolds Number Effects 8 0.92 17.06 0.310 (44.90) 315.9 (600.7) 4.66 5.49 9 0.92 16.94 0.427 (61.99) 316.2 (601.3) 6.44 5.40 w In the first phase, the cascade was tested at a series of vane exit Mach num- bers that bracketed the full stage test. These tests w_re performed with an : ambient discharge and, therefore, varying Reynolds numbers. Test data were processed to determine the mass averaged total presc . _reloss. The results presented in Figure 5.3.2-] indicate the typical vane pressure loss increase with Mach number.

In the se c ond phase of t e sting, the design point Reynolds number was elevated in two steps by backpressuring the.test rig so a match could be made with the stage test Reynolds number. Results showed no effect on the loss level within the accuracy of the pressure measurements. Figure 5.3.2-I also shows results from the uncooled cascade test (build 1), As expected, the uncooled vane loss is lower. However, adding the predictel cooling penalty to the uncooled cas- cade pressure loss results produced a pre-test level that is in a good agree- ment with the cooled rig results.

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0 0e (_ UN COOLE O RI G I__ CO S _ , i 00 2 _ t k o 0 6 0 7 08 0 9 I 0 I t2 V ANE EX I [ MA C H N UM B ER

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"I Figure 5 . 3.2-1 Van e Loss Trends I ; 5.3.2.2 Flow C a pacity an d V ane D e vi a ti on z . Vane cascade inlet f l ow parameter data are presented illFigure 5.3.2-2. Re- sults show flow capacity increasing as a function of t _ochnumber until the cascade becomes choked. The data also show an excellent agreement with the design predicted flow capacity.

Vane air angle deviation is presented in Figure 5.3.2-3 as a function of vane - exit Ma c h number . For this presentation, deviation is defined as the exit air angle minus the gage angle, where air angle is measured from tangential. The exit air angle for this analysis was calculated from continuity. The results in Figure 5.3.2-3 show both the to.al deviation, which includes the affect of cooling air injection and aerodynamic deviation.

., 5.3.2.3 Design Point Performance Characteri z ation The following sections present the measured vane cascade rig inlet pr e ssure and temperature profiles; the exit pressure, temperature; and air angle pro- ; files; a n d the calculated total pressure loss and Mach number profiles. In addition, the total pressure l oss profile is compared to the uncooled annular cascade results, and the exit air angle and Mach number profiles are compared to the design prediction.

+ • 5.3.2.3.1 Inl e t Profiles

!

, The cascade inlet total pressure and ,oral temperature profiles are shown in Figu'es 5.3.2-4 and 5.3.2-5, respectively. Fig,re 5.3.2-4 presents the spaP- wise tutal pressure at three circumferential locations. These profiles show a spanwise variation of 0.7 cm {0.3 in) of mercury and a circumferentia! varia- tion of 0.2 cm (O.l in) of mercury.

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( O F r % , _ : , ........ . 4_ O CASCADE C T _ lib OESIG N I _E D I C TI ON 12 --_ I 0. 5 0 6 0 7 08, 011' t 0 1 1 11 13, I .

VAN E E X IT MACH NUM BER Figure5.3.2 - 2 Flow CapacityCharacteristics 1 II 0 AJ EA O IX V1AT: O N 14 _ _.__ O T O TA L O EV1ATI(. _ ._ ) Ow 10 _ f

.. J _ Io

' 1 1

OI 07 O | O g I0 I _2 MACH N O Figure5.3.2-3 Van e De v iatl o n V e rsusr4ach Number

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: ._ 3.2 2 0 n26 8) !-- I _- " < O Ou.OR_N_I ORIG t N ;, L F::_ : ;' : " _ " [] Qu . _.A , ,, T 2 3 2 1 0 _Q UA DR ANT 3 OF POO_ : _-_"_'_ " " • . _ . (126 4) , _ 1 1 2 6 2 1 o.

, .J _ 320 0 t-- Z , _ 3 1 9 5 ( , _ 1 1 25 8 ) .....

C O < 319 0 (125.6l :. 20 40 60 80 _ O0 I PERCENT SPAN ' J Figure 5.3.2-4 Van e Cascade Inlet Total Pressure i '_ r64 0 1 , 332 ( 630) [7 38 DEG ; ("_ 86 D C G _ Y # , 3 ,o,:G 185 DEG 32_ U _ D E G......

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. 321 _ 30 4 158 0)

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29 6 ( 570) 0 20 40 60 80 100 PE R CENT SPAN , r Figure 5 .3 .2-5 Va ne Cascad e Inlet T o tal Temp e rature | :

k

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IP ' " "_ _ " : O F PC'C_ _ ...... _._ Figure 5 . 3 . 2-5 presents the spanwise total temperature profile at six circum- ferential locations. These results show a midspan circu m ferenti a l variation at , five locations to be within lO degrees, with one location showing an additi o n- a! ]0 degrees of variation. The profiles also show the characteristic lower temperature at the inner and outer walls because of the heat transfer into the rig structure. Overa ;; , these profiles show a well documented inlet with con- : ' sistent pressure profiles for use in calculating the cascade total pressure loss.

5.3.2.3.2 Vane Exit Aerodynamics Pressure Loss Cascade exit total pressure characteristics are shown in Figures 5.3.2-6 and -7. Figure 5.3.2-6 shows the circumferential pressure variation as measured at twelve radial locations in the first quadrant. The profile shows two vane wakes with the lowest exit pressure, highest loss, at the inner and outer walls. A spanwise profile of the vane exit total pressure in quadrants one, two and three is presented in Figure 5.3.2-7. The profiles shown are a result _ of area averaging the circumferential data for each quadrant over one vane pitch, 15 degrees. Alnng with the inlet pressure profiles, these spanwise pro- files were used to determine the spa_wise vane loss characteristlcs shown in Figure 5.3.2-8.

(12B) " ,!

(124) " < \ (1 2 o ) CURVE DIAME TE R = I_ L AB EL ( cm, i n ) c z.u" 1 116) 0 2 7088 (2783 ) n .- + 3 71.75 (.28.25) ( . n E U 2 94 / • I (I Q .I W I (27 , 42 ) ( n x 4 72 79 (28 66) ,112 ) r'l B 7383 (2907) 0 6 74.90 ( 2 9.49 ) I - - 0 A 7 _ 5 9 4 I_ "Q g , 0 ) k- 274 I-- (10 6 ) • 8 76. 118 ( 30 3 1) X 9 78r _ ( 3 0 . 73) uJ Z 10 79 . 0 ,41 31 . 14 ) > (1 04 ) Y 11 80,13 131 ._1 X 12 8120 1 3197) (100) FJ6) l i v i I (92)80 D8 9 _ i 10,4 112 120 *, CI R C UMFERE N TIAL P O SITI ON , DEGREES ._ Figure 5.3.2-6 CircumferentialVane Exit Pressure Characteristics \ ;- , .

33 b .... ..... F 5 ,: , _L * ,' -

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, 12 0) l. ' - z 2 7 4 C ) Q UA D R A NT1 • [_ QUA DF_ NT 2 { ') QUA D RAN T 3 264( 1 0 8 _ "_ ' ' I_ o 4_I v J 2 0 40 60 80 1 0 0 PERC ENT SP AN L, _ Figure 5.3.2-7 Spa n wise Va ne Exit Pr e ssure Profile w (_ _ Q U A D R ANTt • Q UADRANT2 Q UADRAN T 3 ": 1 I _ _ - ---- g g • _ o _ o o ..... _

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, ooo o 0 20 dO 60 80 t OO PER C ENT S PAN Figure5.3.2-8 SpanwiseVane _oss Characteristics 60 ,_:._ JD ' .b ) I ' The results in this figure show high endwall loss in the inner and outer 20 percent span regions and lower midspan loss from 20 to 80 percent span. These trends show good agreement for the three quadrants measured.

Figure 5.3.2-9 pFesents a comparison of average spanwise pressure loss for ' similar conditions tested in both the cooled and previous uncooled rig tests.

llqetwo test points shown for the uncooled rig bracket the cooled cascade test . exit Mach number. Overall, the data trends are very similar. The three sets of data show the highest Ioss in the endwalls with comparatively lower losses in the midspan region. As expected, data from the Cooled Rig Test Program exhioit higher losses. Endwall losses are more pronounced because of platform c : _oling and leakage effects and midspan loss levels are increased over uncooled r_g test data because of the losses encountered with the addition of cooling flow.

01 6 0

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"-, I OF F'j :C ;' 2 Y PERCENT SPAN Figure 5.3.2-9 Vane Loss Profile Showing the Influence of Cooling on Performance } The contour plot in Figure 5.3.2-10 shows vane pressure loss measured in one quadrant. Typically, the areas of high loss are the endwalls and the vane wake, with the highest loss occurring at the endwalls. This profile also shows a nonsymmetrical pressure gradient emanating from the suction and pressure sur- faces. This nonsymmetry is primarily a function of the measurement technique as opposed to airfoil aerodynamics. In addition, Figure 5.3.2-I0 shows the , 2 zero loss core region occurring between the vanes. This flow region accounts for approximately 50 percent of the pitch and 60 percent of the span, Since the core region is expected to be loss free, this confirms the validity of the ,_ test measurements.

_2 61 Figure 5.3.2-10 Contour Plot of Vane Loss Characteristics A summary o f vane cascade loss characteristics for each exit Mach number tested is presented in Table 5.3.2-11 on a quadrant by quadrant basis. The results indicate a consistent difference in measured pressure levels in each quadrant, with the first quadrant having a nigher level of loss. These dif- ferences are m_.stly attributed to circumferential tolerance variations in the experimental engine hardware and not measurementerror.

TABLE5.3.2-11 CASCADE LOSSBY QUADRANT Test Average Quadrant Quadrant Quadrant Average Point N o. MachNo. l 2 3 APt /Pt 1 0 .7 0 3.41 2.97 2.81 3. 0 6 2 0.85 4.97 4.40 4.08 4.49 3 0.91 5.76 4.95 4.75 5.16 4 1.00 7.06 6.33 6.21 6.53 5 1.08 8.27 7.59 7.52 7.80 6 1.16 9.50 8.91 8.64 9.03 7 0.56 2.23 ! , 99 1.94 2.06 8 0.92 5.92 5.40 5.11 5.49 9 0.92 5.83 5.36 4.99 5.40 Temperature Profiles OF P ....._.... ;:_y Vane cascade exit temperatures, _s measured in the fourth quadrant of the rig, . are presented in Figures 5.3.2-II through - 1 3. Figure 5.3.2-11 shows the cir- cumferential temperature distribution and the influence of cooling flow across the vane gap. Results indicate that the coolant has mixed effectively with the main stream, spreading out to approximately 50 percent of the gap. Tempera- tures at the endwalls are consistently lower because of the _nfluence of gas path leakage as well as platform cooling.

The integration of circumferential data into a spanwise profile of vane exit total temperature is presented in Figure 5.3.2-12. The effect of leakage and platform cooling on endwall temperature levels is prominant. Temperatures in these regions are approximately 28°C (50°F) lower in comparison to the average midspan temperature.

The contour profile in Figure 5.3.2-13 shows the circumferential temperature gradients as measured at the vane exit plane. The wake region, as expected, ': has the lowest temperature level, followed by the endwalls. The nonsymmetry of the gradients on the pressure and suction surfaces is primarily a result of t the measurement technique ; ( 620 ) 3 15 C U R V E D IAMETER (_) LABEL (cm, i n ) • } 69 64 ( 27.42 ) (560) " + 3 71 75 (28.25) 293 t _ O 2 7068 5?7.03) . _ X 4 72 79 (28 e _) ( 54 0)" (_ 6 ?4_W ) (2949) _ _ _ Iz _ 7 75 . 94 (29 . 90) • _ 271 1 ,_ B 7698 (3031) ( 5 20 ) " X 9 78 06 (30.73) I-- Z 10 79,0 _ 1 (31 14) _ u J 260 Y 11 b01 3 (31 . 545) q_3 0 0 ) " X 1 2 81.20 (31 97) < :> 2 48 (480) 1460) 22 _ ( 440) , _ , , - T , 8 12 16 20 CiRCUMFERENTIAl. POSITION, DEGREES : Figure 5.3.2- 1 1 CircumferentialTrends of Vane Exit Temperature J , : / 63 _h ,_ " P I k -- _ o x I w ( r , _3 0 I Z i < > J _ 5 2 0 :

!.l

2_ L • t c 5 0 0 1 20 dl O 6 0 80 1 00 _ PERCEN T S PAN Figure 5.3.2-12 Span , ._ise Profile of VaneExitTemperature 1 Exit Air An_le A summary of vane exit air angle characteristics is presented in Figures 5.3.2-14 through -16..The data in Figure 5.3,2-14 show spanwise air angle trends, as integrated from a circumferential traverse of quadrants I, 3 and 4.

The general slopes of the curves are uniform, showing a high exit angle at the root and a low angle at the tip. The steep gradient in air angle is further indicated by the contour plot in Figure 5.3.2-15. This figure shows the cir- cumferential angle variation over the vane gap with the discontinuity across the vane trailing edge.

A comparison of these results to data acquired during the preceding Uncooled Rig Program is p_esented in Figure 5.3.2-16. Overall, the data are in close agreement. When compared to th _ design prediction, however, there is consider- able deviation, particularly ac the root and tip locations. This is a result of the streamline analysis used in the prediction.

Exit Mach Number A spanwise profile of vane exit Mach number, as calculated from the measured exit total pressure rrofile and a linear distribution of static pressure from the inner and outer wall pressure taps, is shown in Figure 5.3.2-17. Th; re- sults essentially duplicate the vane exit spanwise pressure distribution shown previously in Figure 5.3.2-7.

Figure 5.3.2-17 also compares the measured data to the design prediction. In i.

general, the prediction is in good agreement with test data throughout the 25 to 75 percent span region. Since the prediction does not _ - count for viscous effects, Mach numbers in the endwall r_gions show an expect dissimilarity.

Vane Pressure Distributions Vane pressure distributions were c, btained by normalizing the airfoil static pressure, acquired from surface pressure taps, to the inlet total pressure at the respective radial locations. A comparison of test data to the design intent for the vane root, mean and tip sections is presented in Figures 5.3.2-18, -19 and -20, respectively. In the evaluation of these distributions, it is note- worthy to point out that the measured span location at the root is II percent as opposed to 0 percent for the prediction. Similarly, the measured tip loca- tion is 89 percent span as opposed to lO0 percent for the prediction. Furth_r- more, the prediction is based on an inviscid flow analysis.

In taking these factors into consideration, there is good agreement between test data and the prediction for the root and tip sections (Fig',res5.3.2-18 and -20). Data in the midspan region more closely match the predic _3n, as shown in Figure 5.3.2-19. This reflects the capability of the analysi,.'o more accurately simulate the aerodynamic behavior in the 20 to 80 percent span region.

Vane Film Effectiveness The entire suction surface of the vane is film cooled only by the air dis- charged through holes near the l_ading edge. This design assumes that the c o oling film does not deteriorate over the long chord of the vane and main- tains acceptable metal surface temperatures.

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/_ QU A DR A NT I le _I QU A DRANT 3 14 ' Z 10 ---- < PERCENT SPAN Figure 5.3.2-|4 Spanwise Vane Exit Air Angle Trends Figure 5.3.2-15 Contour Plot of Vane Exit Air Angle 66 , ,. ._} OF p OL /

V

0 EL,':HPT C O OLED RiG I-] EEE H P T UN C O OLED R I G * > _ EEE HFT DESIGN PREDICTI O N 20 _ _ _ I _ P ERCENT SPAN Figure 5 . 3.2- 1 6 Comparison of Air Angle Trends . I G4

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( _E EE COO LED RIG DESIGN PREDICTI O N 0 EEE C OO LED RIG 0.7 B 0 2 O 4O 6O q 0 100 PE RCENT S P AN t Fig u re 5 .3 .2-17 Spanwise Profil e o f _n e E x it Mach N u m be r t

X

_ 67 :2 , _, SPAN 0 P RE SSU RE SIDE 11 % -- D E SIGN P R _"O IC 0 % L [] SUC T IO N SlO E I1 % [ 2O 4 _ 6 O 8O I 00 -_" AX I AL CHO R D , PER C ENT Figure 5.3.2-18 Vane Root Section (11 Percent Span) Pressure Distribution

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0 9 [

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o e ] D O _ 0 5 SP A N O P R E SS URE SJDE 50% [ _ SUC T I O N SIDE 50 % M DE S I GN P RE D I C S0 % o_ l j 2 0 4O _ 0 I O 1 00 AXI AL CHORD , PER C E N1 ' Figur e 5.3.2-1 9 Vane Mids p an (50 P e rc e nt Span) Pr e ssur e Distributi o n ,i,- 3 8 '_ -

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0 20 ' _ _ 80 1 00 AXIAL CHORD , PERCENT - i Figure 5.3._-20 Vane Tip Se c tion ( 89 Per c ent Span) Pressure Distribution ! To verify the effectivenes s of th e su c ti o n surfa c e c ooling, one van e was modi- fied by plugging the last two cooling supp l y cavities, thereby allowing air to - . enter only th e front cavity. ( As s h own in Figure 3.2.3-4, this c avity supplies ; cooling to th e first t hree rows of spanwise inje c tion holes near th e l e ading e dge and the second three ro_v s of spanwise inje c ti o n holes near t h e sucti o n surface.) Temperatures were measured by thermocoup l es positioned at four axial ' : lo c ations and two radial locations on the s uction side.

Results of the evaluation are presented in Figurc 5.3.2-21. The measurements (how that the film cooling effectiveness is far more superior than initial l y predicted durin g the vane design and in good agreement with a refinement of this prediction. Test results suggest that, because of the location of the stagnation point and the pressure distribution around the leading edge of the airfoil, the first three ro._;s of cooling enhance the suction surface effec- tiveness. This suggests that it may be possib l e to design a second-generation vane for the same application with a smaller percentage of l eading edge cool- ant _iow.

5.3.3 Turbine Blade Performance Analysis Data acquired from both the full stage and vane cascade tests were used in the analysis of b l ade performance. As part of the a alysis, a streamline model was generated to evaluate blade incidence effects and turning. The following sec- tions p r esent a discus s ion of pressure loss, deviation and internal aero- dynamics.

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e o uJ E E E H PT RIG D A TA E EE E x TRAP_,LATION Z

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_ 1 S - S URF A C E D I S T ANCE P -R OM ST A G NA T I ON POINT D - C O D 'NG H OLE D,AM E IE R I I i / I

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TG - TE M P E RAT U RE , CI A S ST RE A M TF _ T EMPER A TURE FI LM • ' I T C T EMPERAT U RE C O OLAN _ SU P PLY o 1 1 I +O _ 0 _in 1 _) 200 400 600 BOO " S, D FR O M S TA GN ATI ON PT Figure 5.3.2-21 Vane Suction Surfa c e Film Cooling Effectiveness 5.3.3.1 Blaue Pressure Loss Pressure loss is calculated from the measured stage efficiency and vane cas- cade pressure loss levels at Fach nambers evaluated during the stage test. A summary of blade performance trends is p r esented in Figure 5.3.3-I, showing t h e relative pressure loss as a function of exit Mach number.

The trends, which include tip clearance losses, show the expected relationship of increasing pressure loss as a function of Mach number. In addition, losses increase with positive incidence and decrease with negative incioence relative to the design incidence angle.

Of particular significance are the results at transonic Mach numbers. This blade was designed to minimize loss in the transonic region by the use of a refined curvature distribution on the suction ide downstream of th. _ throat, plus trailing edge coolant ejection to control tf • shock structure. T1ese con- cepts, which were demonscrated in the High-Pressure Turbine Superson+c Cascade Supporting Technology program (Ref. 3), are shown to be effective in providing a m_re linear as opposed to a highly skewed loss trenJ in thu transoi_i c region.

5._.3.2 Deviation Blade deviation, as calculated from the air angle at the exit plane minus the throat plane air angle, is shown in Figure 5.3.3-2. The results in t h is figure show both the total deviation, which includes the affect of cooling air injec- : tion and aerodynamic deviation. As expected, deviation increases with Mach !

number and the addition of cooling air.

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, / " _ O i S _ J j . / 0 1 4 J AIleal ) / 0 2_ . ,7 - ; o, 1 -- ..m 3 _ I_ I 1 2 I I(% I20 i 2 4 i 28 I12 I_ I " _i BL AD E E'," R EL ATI V E M A CH NU M B E R . , Figure 5.3.3-I Calculated Blade Pressure Loss as a Function of Exit Mach : Number I SO

i /

<_ f ' O TAL Df _ VIATION _ 1 AIERO D f V I ATI O N 4 0 -- I . '!

I- I , 10 11 12 13 4 MACH NO .

Figure 5. 3 . 3 -2 B l ade Deviation Versus Mach Number 5.3.3.3 Internal Aerody n amics A streamline simulation was generated and used to determine the stage internal aerodynamics, includin_ incidence trends and blade turning characteristics.

• The simulation was made for the design point con6ition evaluated during the full stage test. The stage inlet and exit total tenperature and pressure pro- files were input values. Vane exit conditions f_r the design point were : ascertained from the results of the cascade test. In s pecific, the vane total pressure loss profile was adjusted for stage exit Mach number level, the cas- cade exit total temperature profile was adjusted for stage inlet _emperature level and the vane exit air angle profile was adjusted so that the streamline ' simulation matched the vane endwa,l static pressures in the stage test. Also, the simulation accounted for the measured stage inlet flow and cooling flows.

Table5.3.3-Ipresentsa comparison of the simulation predicted endwallstatic pressure to the test data.Again, there is a good correlation betweenthe sim- ulation and the measured data.

WALLSTATICPRESSURE MPa (psia) : : Measured Streaml ire Vane Leading Edge Inner Diameter 0.45 (65.5) 0.45 (65.6) ' .

Vane l_eading Edge Outer Diameter 0._5 (65.5) 0.45 (65.7) Vane Trailing Edge Irner Diameter 0.23 (33.5) 0.23 (33.5) Va-o Trailing Edge Outer Diameter 0.27 (38.6) 0.27 38.6) _, BI , .e Trailing Zdge Inner Diameter 0.09 (13.2) 0.09 (13.1) Blade Trailing Edge Outer Diameter O.IO (14.0) 0.I0 (14.5) _mulation-predicted blade exit air angle and exit Mach number profiles are presenteo in Figures 5.3.3-3 and -4, r=spectively, and compared to measured data. In both cases, the simulation r qlts are in good correlation to the nJeasureddata, verifying the validity of ne simulation.

i • w [] DES ' GN P O I NT DATA MODE L I 0 MEASURED DATA 70 _ ,.....

"= O F PO0 / _ ' "'" _''_'; _c , , .....

_ U A L j_ < PERCENT SPAN Figure 5.3.3-3 Spanwise BladeExitAir Angle O _ -

- J Y

= -° - \ /

0 . 44 _ I 0 RIGSIMULATION 0.42 _ M EA S URED DATA 0.40 ,_ 20 40 e 0 iN ) 1 0 0 % SPAN Figure 5.3.3-4 Spanwise BladeExitMachNumber 73 ' _'I" A summary of vane and blade internal aerodynamics is presented in Table 5.3.3-11. This table compares predicted integrated core / low spool aerodynamics to the rig simulation. In general, simulation data for the vane are close to the design intent, particularly at the mean section. The differencP in the root and tip values is expectpd hp r euse of the inherent limitations of the j design analysis. Also, as expected, the aerodynamic differences produced by , the design are more pronounced in the relative frame, as shown by the simula- i tion data for the blade inlet root and tip sections. Of particular interest is this appreciable difference in blade incidence and its effect on stage effi- ciency. This is shown in Figures 5.3.3-5 and -6. Figure 5.3.3-5 shows the spanwise profiles of the blade inlet air angle for the design pressure ratio '! of 4.0 and varying corrected speed using the design point simulation for the ,, base case (corrected speed of 244). Also showr is the blade design metal i angle. It is noteworthy to point out that the design metal angle accounts for i vane loss and combustor exit temperature profiles of an engine rather than rig I environment. Consequently, the calculated blade inlet air angle profile- ,n I Figure 5.3.3-5 are flatter than tLe metal angle profile. As shown, negative ti o n t o t he d es i gn metal a ng l e , Fr o m t h i s poi n t, po s itiv e i n cid en c e , wit h t he i incidence dominates in the 0 to approximately 70 percent span region in rela- att e nda n t h i gh pr essu r e l oss , i s pr e val en t, A t t he h ig he r c o rr e ct e d s p ee d_ less of th e blad e l e adi ng edge spa n operat e s at positive incidence. . Th e oppo- 4 site e ffect occurs at lower corr e cted sp ee ds.

TABLE 5.3.3-11 i INTERNAL AERODYNAMICS* Vane ROOT MEAN TIP c_in (deg.) (90.0) 90.0 (90.0) 90.0 (9n_O) 90.0 _out (deg.) (If.6) 15.9 (I0.6) I0.5 (9.1) 4.0 Mn in (0.09) 0.09 (0.08) 0.08 (0.07) 0.06 Mt_out (l.Ol) 0.78 (0.93) 0.95 (0.87) 0.75 egas (deg.) (78.4) 74.1 (79.7) 79.5 (80.9) 86,0 Blade : _ (deg.) (32.7) 67.9 (41.5) 38.89 (59.2) 105.9 Pout ( deg.) (16.3) 8.4 (17.2) 16.5 (18.1) 27,3 Mn in (0.38) 0.23 (0.25) 0 ._.8 ( 0 .15) 0.05 Mn out (I.23) 1.22 (I.25) 1.25 (I.30) 1.04 _)Gas (deg.) (131.O)r103.7 (122.7) 128.5 (I02,7) 46.8 _out (deg.) (38.3) 21.2 (43.5) 42.5 (48.6) 81.3 Mn out (0.56) 0.49 (0.54) 0.53 (0.54) 0.48 ' *Values in parentheses are for the restaggered integrated core / low spool design. Values not in parentheses are simulation data at the design point.

' _ / _ 74 _ , 7. 11 o O DES,GN M_ ' A L ANGLE , _ > D ATA MAT _ H " R'ESU L TS ) T_ RIG SI M ULA T ION = . , ,...: .... _ / ' _ . _, , IS N / V r T A INCIDENCE OF POO R Q U A L ITY l og O 244 BASE ; O 253 - 3.9 _ ? 34 + 3.7 9 0 _ _ 222 +8. 1 o3 8O UJ : _ LU U.J > --J Z <_ p- gJ . r t e t

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30 _ _ _ _ .. _ 0 20 40 60 80 i oo PE RC ENT SPAN Figure 5.3.3-5 Spanwise Profile o f Blade Inlet Air A_gle lllead v antage o f minimi z ing the am o unt o f p o siti v e incidence is ap p arent in Figure 5.3.3-6, which shows stage efficiency for the data in the preceding f igure. At the root section, where incidence is negative for the entire range of speeds, the level of efficiency is constant. However, efficiency decreases with iower levels of corrected speed since me, re of the airfoil span is _ubjec- ted to positive incidence. This explains the increase in stage efficiency at corrected speeds higher than the design level (Figure 5.3.1-1).

A spanwise comparison of design and streamline simulation data for blade turn- ing is sh o wn in Figure 5.3.3-7. In c om parison t o the design, the si_ilati o n data sh_w a red u ction in turning at the root, indicating that less work is • b ei n g perf o rmed in this area. The midspan regi o n is in agreement with the de- sign, whi l e the tip sh o ws mo re w o rk b e i n_ acc om p l ished than p redict e d b v the _ design. This contrast i n design an d simu . ati o n data is ldrge l y attributed t o ; _i the difference i F ,blade inlet angle.

- pi_G tt ti

O F pOO_ Q_L _ B L AD E N / _"_ T I N CJD_ N C E [ _ ] 2 44 B AS E 0 253 - 39 o /% , 234 +3 7 ° N U

g

?

7 5 i 50 ......

t 4 O ; 20 4 0 6 0 80 IO(I P E R CENT SP A N 4 76 Figur e 5.3.3-7 B lade and Vane Tu rni n g Charact e ristics wi ._ 1 1 1 _i , 5.3_4 Secondary Flow System Performance CharacterizationStudies C The evaluation of secondary flow system performance involved a series of sepa- ' rate investigations to examine the effect of different component flow rate changes on overall turbine performance. The investigation included the main TOBI (tangential on.board injection) flow sensitivity test, mini TOBI flow sensitivity test and compressor end sea l l eakage test. Test results are pre- sented it,the followiqg sections.

In addition, a related rig test was completed to determine the blade inlet supply pressure. Pertinent results from this test are summarized in Section 5.3.4. 1 , while additional information such as the test rig description, in- t strumentationand test sequence is contained in Appendix A.

5.3.4.1 Turbine Blade Supply Pressure Test The turbine blade coe]ing supply system is based cfla pressure-balanced tan- .

9enti_l on-board injection system. This design eliminates the r_quirement for Inner and outer TOBI seals.

As part of the high-pressure turbine performance program, a flow test of the ._ tangential on-board injection system was conducted to verify the l evel of _I blade cooling air supply pressure. For this test, an existing ro;ating disk -, rig was modified to the Energy Efficient Engine configuration. Figure 5.3.4-I shows how a scaled Energy Efficient Engine tangential on-board injection cas- cade and vortex plate were added to the existing rig to simuiate the Energy Efficient Engine system. Pressure and temperature instrumentationwas install- ed as indicated in Figure 5.3.4- L . Rotating pressure taps were ]o_ated on the ' disk to determine pressure rise. ' .

A summary of the test results is shown in Figure 5.3.4-3. Extrapolated to in- tegrated core / low spool engine sea level takeoff performance, the data indi- cate a blade coo]ant supply pressure of 53 percent of the compressor discharge total pressure versus a predicted pressure of 54 percert of the compressor " discharge total pressure. Since the test program was successful in that the J - minimum desi g n assumptions were verified, n o further analysis or testing was conducted. The minimum allowabl e supp l y pressure is 51 percent of compressor discharge total pressure.

5.3.4.2 Tangential On-Board Injection (TO_I) Nozzle Flow Cnaracterization Studies A series of tests was performed to assess the aerodynamic behavior of the ii blade and mini TOBI nozzles in maintaining a positive blade supply pressure and acceptable rim cavity temperatures. Testing was conducted by varying the nozzle flow rates a percentage above and below the design rate. The conditions are l isted in Table 4 l-I under Phase II.

The results of the tests are summarized in Tab l e 5.3.4-I. Overa l l, these :_- ' sults show that temperatures and pressures in the front rim cavity are insen- _i sitive to TOBI flow rate changes.

I

OF POOf_, Q,.laLf'_"_ I • 25 40(10O O ) R 1 757 (0 692) ; 34 29 (13.50) I - 1 O 838 (0.330) -- -_ 2 212 (O 871) ,2 .

i -

E 3 DESIGNED PARTS cm hn) Figure 5.3.4-1 Tangential On-Board Injection System Rig Geometry : ( 0.481 kg / sec 0 PRESSURE ( ) (1 O621b / s ec) • • TEMPERAT U RE ( _ //// . I_ _. _ • - ; / Figure 5.3.4-2 Tangential On-Board Injection Rig InstrumentationLocations }-.f d O F?;C " _ : ., 1.2o + / 1 1 . 18 , , ,,,k ,,/ I, Z,, / __ 1,14 _ / .

5 11o _ _ 1 08 1 06

// I

1 02 1.o I1_" 1 1 I I 1 1 1 1 1 I I I 2 7 0 28 9 29 9 30 9 32.0 _3 0 34 0 (110) (114) (11. 8 ) (122) (126) Il _) ( 1 3 . 4) RADIUS,cm (in) I I I 1 J 1 . 00 1 05 1.1 0 1.15 1 20 RADIUSRATIO Figure 5.3.4-3 Tangential On-Board Injection Rig Radial Pressures Figure 5.3.4-4 presents a summary of rim cavity temperature tre n ds at the dif- ferent test conditions in Table 5.3.4-I. The trends indicate that overflowing the nozzles lowered the cavity temperatures, as expected. However, the temper- ature gradient did not increase significantly with either a 20 percent reduc- tion in blade TOBI flcw or a 33 percent reduction in mini TOBI flow. Moreover, complete elimination of flow from the mini TOBI did not cause a dramatic in- crease in temperdtures. This result suggests that the cavity flowfield con - tained a much higher level of induced swirl than anti c ipated and this suffi- ciently reduces the windage shear effects. Thus, for the type of design in the Energy Efficient Engine high-pressure turbine, the loss of preswirled air from the mini TOBI does not produce a detrimental temperature increase in the front rim cavity.

The effect of main TOBI flow variations on front rim pressure ; evels is shown in Figure 5.3.4-5. As indicated, neither increasing nor decreasing the flow J rate imparts any substantial effect in the cavity flowfie!d. Consequently, a , positive blade supply pressure level is assured. Changes to the exit velocity i did not produce a noticeable change in the swirl level of the TOBI air to sub- stantially change the front rim swirl characteristics.

: The small decrease in static pressure measured with a lower flow rate is cor- roborated by the corresponding small increase in the rim temperature. This is clearly shown in Figure 5.3.4-6, where the rim temperature rise is only 16.6 ° C (30°F) at the minimum flow condition compared to the design condition.

.®

I ,_ , r OR!Gi:" ' t i" OF F : .,..':, L . , ",, TABLE 5.3.4-I HIGH-PRESSURE TURBINE FRONTRIM CAVITY DATA Rim Cav t TOBI Exit Radius Pressure Temps Solid body Ptcalc cm (in) MPa (psia) % PT IN Pr / Pr 1 R / RI Mach l °C (°F} factor MPa (ps_a) Tr I / TEC DeslgnPolnt -I0..........................................................................................................

22.0 (8.7) 0.223 (32.36) 49.1047 1 I 35 (95) 1.01277 25 (I0) 0,227 (32.86) 49.8634 1.01545 1.14943 .261431 37 (100) .4 % _32 0.238 (34.4591) 1.02190 27.6 (10.93 0.229 (33.173 50.3338 ; .02503 1.25287 .24942 3 42 (1083 .479982 0.239 (34.63713 1.03650 30 ( 1 23 0.230 {33._23 50.7132 1.03276 1.37931 .226390 43 (]113 .437282 0.239 (34.6344) 1.04197 32.0 (12.63 0.230 (33.29) 50.5159 1.028 7 4 1.44828 .192429 51 (124) .376422 0.236 (34._609) 1.06569 PT IN 0.454 (65.9) 100 TTIN 315 (599) PT-30.4 (12.03 0.389 (56.4B) 85.7056 TEC 31 (88) Zero Mini Point - 19........................................................................................................

22.0 (8.7) 0.225 (32._1) 49.1262 l 1 45 ll3) 1.04753 25 (lO) 0.227 (32.87) 49.5179 1.00797 1.14943 . 1 88045 48 120) .366645 0.232 (33.6908) 1.06033 27.6 (10.9) 0.228 (33.143 49.9247 1.01625 1,25287 .201294 5l 125) .393963 0.235 (34.0895) 1.06947 30 (123 0.230 (33.37} 50.2712 1.02331 1.37931 .191275 52 1273 .3 7 5140 0.236 (34.2325) 1.07313 32.0 (12.6) 0.228 (33.06) 49.B042 1.01380 1.44828 .133691 60 140) .265586 0.231 (33.4755) 1.09689 PT IN 0.458 (66.38) lO0 TT IN 316 (602) Z PT-30.4(12.0) 0,225 (32.57) 49.0660 TEC 30 (8 7 ) Mln Mini Point - 20.........................................................................................................

,. 22.0 (8.7) 0.227 (32.86) 49.5925 1 1 36 (97) 1.02015 25 (lO) 0.229 (33.22) 50.1358 1.01096 1.14943 .220312 43 (110) .425283 0.23 7 (34.3F25) 1.04396 27.6 (10.93 0.230 (33.4) 50.4075 1.01643 1.25287 .202403 46 (116) .393058 0.237 (34.3677) 1.05495 30 (12) 0.2_2 (33.64) 50.7697 1.02374 1.3793l .193022 47 (118) .375629 0.238 (34.5255) 1.05861 32.0 (12.6) 0.230 (33.34) 50.3169 1.01461 1.44828 .137530 53 (129) ,27066) 0.233 (33.7835) ].07875 PT IN n.457 (6626) 100 TT IN 316 (601 PT-30.4(12,0) 0.304 (44.11) 66.5711 TEC 30 (86) Max Mini Point - 21..........................................................................................................

22.0 (8.7) 0 . 224 (32.42) 48.9506 I 1 32 (90 ) 1 . 0110 3 25 (lO) 0.227 (32.98) 49.7962 1.01727 1.14943 .276327 33 (92) .523483 0.240 (34.7767) 1.01471 7 7 ,6 (10.93 0.230 (33.33) 50.3246 1.02R07 1.25287 .263986 3 7 (lOO) .504045 0.241 (34.9844) 1.02941 30 (12) 0.232 (33.67) 50.8380 1.03 o 56 1.3 1 931 .245369 ?q (103) ,470192 O. L i2 (35.1105) 1.03493 32.0 (]2.6) 0.231 (33.52) 50.6115 1.03353 1.44828 .208_96 4b (1"4) .404855 0.2 3 8 (34.5551) 1.05515 PTIN 0.457 (66.23) I00 TT IN 314 (598) PT-30.4I12.03 0.451 (65.36) 98.6864 TEC 28 (84) Mln TOBI Point - 15.........................................................................................................

2_._ (8.7i 0.223 (32,29 48.5345 1 I 38 (101) 1.01630 25 (I0) 0.226 (32.74 49.2109 1.01394 1.14943 .248350 43 (110) .478784 0.237 (34.1755) 1.03261 27.6 (10.93 0.227 t3_.99 49.5867 ].02168 1.2 E 28 7 .232256 49 (121) .452403 0.236 (34.2526) 1.05254 30 (123 0.23G _33.34 50.1127 1.03252 1.37931 .225573 51 (124) .440652 0,238 (34.5427) 1.05797 3 2.0 (12.6) 0 .22_ ( 33 . 0 2 4 9.63 1 7 1 . 02261 1.44828 .1 7 0855 5 8 ( 1 38 ) . 3 _84 6 5 0.2 3 2 ( 33 . 699 7 ) 1 .08 3 33 PT IN 0 . 459 (66 .53 100 T T IN 31 0 (590) PI-30.4(12.0) 0.391 (56. 7 4 85.2848 TEC 33£ (92) Max TOBI Point - 18.........................................................................................................

22.0 (8. 7 ) 0.226 (32.83) 49.3981 1 1 32 (90) 1,01289 25 (1 0) 0.23 0 ( 3 3. 38) 5 0 .22 57 1. 0 1 6 7 5 1.149 43 . 27 2 1 5 9 3 4 (94) .5 16 6 3 5 0. 2 4 2 (3 5. 143 0 ) 1 .0 2 0 2 6 2 7 .6 ( 10 . 9) 0. 232 (3 5 .69 ) 50.692 1 1.0 262 0 1 . 2528 7 . 2 55 1 0 7 3 7 ( 100) .4 8731 4 0 . 2 4 3 (35 . 249 9) 1 . 03131 3 0 ( 12 1 0.2 34 ( 3 _.98 ) 5 1 .1285 1 .0 3 50 3 1 . 3 79 3 1 . 2340 19 3 9 (103) . 4 4 8683 0 .2 43 (3 5. 3006 ) 1 . 03683 32.0 (1 2. 61 0 .2 3 3 ( 33 . 86) 50 . 9479 ]. 03137 1 . 44 928 . Z 3 0963 45 (113) . 389 2 67 0 .2 40 (34 . 8 2 69) 1 . 055 2 5 P T IN 0 . 4 5 8 (66.4 6) lO 0 TT IN 3 1 5 (6 00 ) PT-3O.4(12.0) 0.39 1 ( 5 6.7) 85. 31 45 TEC 2 8 (8 3 ) !

i 8O \ .

1 10-- _ TOSI WMi NI r - _ I % _ RIG) 1 % _ R I GI _ J _ DE S IG N 102 _ / PT TOB I M I N P T 12 3 I f_ 0 ,5 7 MI N I MA X PT 2 1 I_ I _ 068 ,- P _ U u . I 0 2 I I I I I 2 0 22 25 30 3 " I ; FRONT RIM CAVITY R AD I US , cm hn) -_ Figure 5.3.4-4 Front Rim Cavity Sensitivity to I,_BIFlo w R ate Variations 02 3 4 _ 13 4 Ol 0 23 3 -- 1 33B b 0 23 1 _ _ MAI N - 4 J _ % w RIG (33 6 1 02 3 0 -- O E N GN POt N T 43 3 4 _ w * w 13 3 2_ > t, 3 3 ot 0Z._ ( 1 2 II 1 2Q% POI N T t_ 0 _ 3 -- 13 2 41 o _2 4O 45 @ 0 f _ @_ ) II { 14 _ J1 6_ 1 1 9 1 _ (2 _) [ 24) i _ i _) D I AMET ER , crn (tn _ : _ ; Figure 5.3.4-5 Rim Cavity Pressures withBladeTOBIFlowRateVariations N "_ 81 ,_ . T , h ' .

| " - OR : .....

OF p OOl, - (1 5 @ "' F _ _ RIMC A VI_ Y T EMP E R A T URES ( _ 20 % POI NT 15 54 / ' - MAIN 3 15% _ ' R f G (130t _ c

/

_ ( 12Gl I- > / J ( _ . 22 % POIN T 1B 4 3 _ j ,_ _ ,, . , ,. , ._ / WM A I N 4 8 5% _RIG 111 0 } o_ 3 7 _ ( l oo) L 32 _ POINTS ( 151_ ? 19 0 1 DESIGN P O INT 0 11 8 1 0 2 6 ' 8 0) L T O 81DI L ,CHARG E " I - TE MPERATURES 2_ -- . I 1 I I I i ! I (7°13s 4 0 , s so _ e o _ 71 76 /1 4 ) { 1 6) 1 18_ 1 2 ( I) _ . 2 ) (2 4 1 ( 2 1_ ) i2_ ) { 3 1_ 1) D'AI A ET E R cm (m) Figure 5.3.4-6 Rim Cavity Temperatures with Blade TOBI Flow Rate Variations Similar results are pres e nted in Figure 5.3. 4 -7 for variations in the mini TOBI flow rate. Again, the pressure ]eve] differences between the design and off-design conditions are smaI]. The correspo n ding temperature ]eve]s are a;so small, as indicated ir,Figure 5.3.4-8.

The information in Frgure 5.3.4-9 shows the calculated swirl factor in the front cavity. As indicted, there is a gradual loss in swir] strength as the flow approaches the rim. These results are consistent with previous data trends as wel] as the pressure gradient results in Figure 5.3.4-]0.

5.3.4.3 Compressor End Seal Study Ibis test evaluated the influence of compressor discharge sea] flow rates on the pressure distribution in the front disk cavity and the blade TOBI supply plenum as well _, the resultant effects on blade coolant flow. Testing showed that the sea], which is a 9-knife edge geometry, performed as designed. Changes in flow rates had litt]e, if any, affect on the performance of the secondary flow system. The blad.:supply pressure is insensitive to changes in the sea] . clearance. In additio as sho_ n in Figure 5.3. 4 -11 , t h e desl_ prediction is f in good agree m ent with the cr._,flow and rig measurements.

t

0 222 ( 3 2 2 1 0 221 _3 2 o l I [ I I l l ! DIA M E ' r z_ c m { in ) , i i

i

Figure5.3.4-7 Rim CavityPressures with Mini TOBI Flow Rate Variations _1501 !

5 4 _ J J% _ MA IN P OIN T 20 11 ] 0 1 _MI N I o ]a ' _ _I I G UE 51( ; N POI N f _J , 19% _ M A I N POIN T 2l ¢ 3 "C -M IN I _ o 6 t 1% _IG 0:: ,_I O l Ol S IGN _ I _ 0 _ O _C M A R _ E I _ T$ , 2, ,_11 TEM I_II ATU _ S " ' _ I I ! | I I, I I , l 4 _ 41 I I ( 1I I 1 2 _1 (2 _i { 24 _ i . . - _l l { ; N_I 1 3 0 _ D I AP AE TER, C ,' T _(ru r al I 1 Figure5.3.4-8 Rim Ca v ityTemper a ture s with Mini TOBI Flow Rate Varietio , \

., . # ; , ' /

r O F P0,";, ; ", ; J , , . ., i 0 6 I e_ _ 04 -- _ 0 3 -- o er P " / W T O BI _ M I NI 9 2 T OO I M AX P T 11 485 5 7 Z _ (% _ R IG ) % _ R IG, _ T C ;B I MINP T 1 2 3 1 _ 57 • _ 4 1N I M A . _,F ' r 2 1 3 _ _ -] MIN I MINP; 2 0 3 _i 38 " I -- I M I Nt Z ERO PT 19 39 6 O 0 / .

I I I J • 20 2 2 25 2 1 3 3 LS, t l C l_ rl 1, (12 l ,131 FRONT RIM C AVI T Y RADIUS , c m fin) Figure 5.3.4-9 Front Rim Cavity Sensitivity to Swirl Level Difference 10 4 -- F R ONT RI_,! CAV I TY PP ESSUR.: GR - _ DIE N T DUE T O VA R IATION IN MA i N T O B I & MffN,TOBI ) 0 3 -- > Z U C I _, L _ _ , T OR I ¢ _ M I N I (% w R I G I 1 % _ R IG ) 1 0 t F DiS' ,N PT _--_ TOBI t _A X PT I 4 1S 05_ I _ T O P IMIN P T 1 2 3 1 5 057 • MIP I IMAX P" 21 :)_I d _ 0 8 L _ MINI M I P IP T 20 ,"19(_ 0 _m -- M I N I ZI ROI_ T °9 _ 000 , o I l I 2 0 3 2 2 e 2 5 4 2 1 9 J 0 4 3 3 0 (lt O I 19 0 1 l lOO , I ll OI 11 2 0 I II J OI FR O'IT R IM C A VI T Y RADIU S , cm hn) H, 2 2 0 c , _ (_1 l , nl I Figure 5.3.4-10 Fr c ,t F,'(, Ca v ity Scr s itivity to Pr e s s .me Variation \ I I 01 0 - T f 1 c _ ' _E -T E ST PREOIC % I O N U _E S . EXISTING LAII S E A , . C. ) RIm_IL AT IO N j, o _ - --- O F Fu , ..._ (_ , L . , , ., r"_' I t 00 3 4 c m |0 01 _ 1 1 ml 1 o m ,t

' i

10 11 12 13 14 5 1 6 17 18 9 PR E SSURE RATIO , Pu / Pd Figure 5.3.4-]] Hi n h-Pressure Compressor Discharge End Seal Leakage Study R u su]ts ,_ 5.3 . 4.4. Van e L=dkage Study ] Leakage rates in the vane attachment area were ca l cula_ea and : or,. ' ared to " levels predicted during the evolution of the turbine d_sign. Attachment leak- : ag e rates were c alculated by de4uc+ing pretest vane fo_I a,_dblade outer seal f l ows and vane p l a tform e stimates fr o m th e me asured c ote_ f lo w o f th e vane i assembly. The results presented in Table 5.3.4-II indicate that the effects of thermals an_!vane loadings reduced attachment leakage to v_,ues close to the f l ight pr o pulsion syste m preliminary definition.

"ABLE 5.3.4- . I VANE A T TACHmeNTLEAKAGES FLIGHT EXISTING dPDATE BA S ED UPD_FE BASFD HP_ WARM PROPULSION FEATHERSEAL ON LEAKAGE ON HPT RIG RIG SYSTEM EXPERIENCE TEST PROGRAM COLD FLOW TESTS RESULTS Vane Inner Diameter Fr o nt O .Ob O.OS O . Z2 0 . 0 4 C . IS Mld 0.11 0.22 0.35 0.78 0.23 Rear 0 . 4u 0. 4 _.__ 6 6 0.44 1. 58 0. 3 8 Inner Diameter _ _ T ot a l 0.56 0.7 3 1. 0 1 2. 40 0 .7 6 Va n e Out er Diameter F ron t 0.1 3 0.25 0 .18 0 . 2 0 0 .1 2 M r .' 0. 1 3 0. ? 5 0. 34 0. 47 O .lt _ Re a r 0.2 1 0 . 26 0 . 30 0 .4 0 0 . 25 •I Outer Diameter .......

! T o ta l 0.47 0. 77 0 , 82 1. 07 0 . 53 I Vane A tta ch T o ta l ( _ Ri g Flo w ) 1. 03 1. _0 1 .8 3 3.47 1. 2 9 q | 5.4 POST-TEST INSPECTION RESULTS All rotatinghardware,the outer case assemblyand outer airsealasse_ly of the high-pressure turbinerig were disassembled and inspected.In general,all partswere in excellent condition.

Post-testvisual i_ - pectionof rig instrumentation showed that the pressure and t emperature sensors,alongwith associated hardwareaccessories, m intained their pre-testcondition.In addition,post-testsystem validity checks for temperature and pressuredata acquired during rig testing indicatedmeasure- ment accuracies cons,stent with pretesterror / accuracy predictions and provid- e_ an excellent correlation betweenestinw_ted and actualdata.

I _o distre_ was observedon the blades,vanes and cuter air seals. Tne excel- 'entcon e . '.'on of these c_mpo n entsis shown in Figure 5.4-I through5.4-3. No evidence of coolingair hole plugging was notedwith thesecomponents.

Also, there was no indication of seal rubbingwith the high-compressor dis- charge seal, rear tb' - ust balance seal and the No. 4-5 bearing compartment buffer air seal. Dark line marks on the dischargeseal land at the 3 o'clock positionindicatedthat the knife edges may have been very close to the land at this location. However,therewas no appreciable rub.

i rig was in excellent condition,the No. 4 bearingshowed some minor distress.

ill While the majorityof componentpartsassociatedwith the aerodynamics of the Scoringmarkswere notedon the innersurfacesof the No. 4 bearingoil scoop, i the No. 4 bearingfront carbon seal plate, and th? outer diametersurface of the No. 4 bearing h ub. Thesewere the result of the oil scoop and the carbon seal plate spinningon the No. 4 bearinghub. Most rig-associated components • : were also in excellentcondition.Teardown inspectionshowed that the reason • for bindingof the traverseinstrumentation ring was insufficient axial clear- ance betweenthe rotatingring and stationaryparts.Galling and metal pickup were visibleinboardof the front 'W' seal on these parts, as shown in Figure 5.4-4.

5.5 SUMMARYO F RESULTS : T he Energy EfficientEngine high-pressure turbine componentwas successfully testedover a range of conditionsthat simulatedoperationof the integrated core / lowspoolat both intermediate and high power levels.Therewere no major difficultiesthat preventedthe acquisitionof performancedata. The good m e c h anical o p e ration o f the test rig was verifiedby the excellentpost-test condition of the test components.

The main result is that the measured turbinerig efficiencyof 88.5 percent surpassed both the rig designgoal of 86.5 percentand the estimatedcomponent efficiencyfor the integratedcore / low spool of 87.1 percent. Furthermore, this efficiencymeasurementwas repeatable, indicatingthat no performance deterioration occurredduri_ig the test.When comparing this measuredefficiency to thatof the uncooledris, thereis only a 2.6 percentpenaltj in efficiency from coolingand leakageflows.

I I I Figure 5.4-] High-PressureTurbine Rotating Disk Assembly | i i f igure 5.4-2 Po st- Te st C o n d ition of Turbine Vanes !

I

F igu r e 5 . 4- 3 H i gh - P r essu r e Tu r bi n e Ou t e r Air S ea l S eg m e nt s V a ne pe rf o rmanc e , a s do c u ment ed by c a sc a de t es tin g, wa s ge n e ra l l y in goo d a gr eement with resu l t s from th e p re c edin g Unco ole d Ri g Te s t p ro g ram an d the des i gn p r ed i ct i o n. M e a sured p r ess ur e lo s s w as at pr e d i c t ed ] e ve]s, an d sho w e d th e e xp e ct ed incr e asein l oss at h igh e re x it Mach n u mb e r s . T y p ically,t he en d - wa il s exhibitedthe highest l oss, whi l e the midspan showed tee l owest l oss.

Operation at high Reyno l ds numbersproducedn o chan g ein l oss characte_-istics.

Of particu l arimportance,cascade testing demonstratedt h e effectivenessof I the vane coo l ingscheme.Resu l t s indicatethat the coo l ingfi l m from the vane l eadingedge remain s attached over t he entire suction surface t o provide a high e r than p r e dict e d cool il ;g e ffectiv e ne ss . Thi s sug g es t s t h at t he d e s i g n o f a second-generation vane for this turbine configurat io nc o u l d require l es s c o o l ant.

Th e hi gh levelo f s ta ge e ffici e ncyal o n g with t h e goo d van e pe rf o rmanc e indi- t cat es that the b l ade performedat higher than expected l eve l s. Perf o rmance trend s showed the typica l increa s e in loss with higher exit Mach number throughout the tran so nic regi o n. Thi s i s a re sul t o f the lo w l o ss b l ade design and trai l ing e dge c ool ing ejecti o n.The ana l y s i s o f b l ade perf o rmanc e a l s o s howed that testing en g ine hardware in a rig c o nfigurati o nimpart s sl i g ht c o mprGmises in p erformance.

The turbine s ec o ndaryf l ow s y s tem perf o rmedacc o rdin g to the de s ign intent.

Characterizati o n s tudie s showed that s y s tem perf o rmance,in particu l ar the p re ss ure-ba l anced tangentia l o n-b o ard injecti o n s y s tem,i s in s en s itive to f lo w and p re ss urevariati o n s .

88 ,_ i _ I F i g ur e 5.4 - 4 H i g h- P r essu r e T u r b i ne F u ll S t a g e R i g E x it P r obe In s tr um entati o n Ring Showing Gal l ing an d Metal Pickup o

"I

SECTION6.O CONCLUDING REMARKS : A sing]e-stage turbine configuration is attractivefrom the aspects of cost, ; weight an d maintaina b ility. Howev e r, in comparisonto a two-stage turbine, I th esebene fitsca n be c o m p r o mised b y a low e r le v el o f perf o rmance. T he res u lts fr o m t h is t es t h av e dem o nstrated tha_ it is possibleto minimize the perfor- mancedecrement with a highly-loaded, _ransonic, single-stage design.

This test has providedan opportunity to evaluatethe numerous technology fea- tures in the Energy EfficientEngine _ingle stage high-pressure turbine com- po n e nt.Be s i desh i gh rim speed operation,which is a prerequisite for higher pressureratio engines in the future,testinghas a l so shown that substantia l improvementsin p erf o rmanceca n be achie v ed with advances in airfoi l aero- dynamicsand sealingtechno l ogy. In severa l instances, resu l ts have indicated where additiona l performance gains cou l d be attained,shou l d a design update be contemp l ated.

The resu l tsof this test have wide application. Many of the advancedconcepts, suc h a s _irf o il d e s ign, are direct l y applicableto commercia l and mi l itary e ngines p lan ned in t he n ear f u ture. Th e se concepts also establishthe tech- no l ogybase for the designof far-termfue l efficientturbofanengines.

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APPENDIX A BLADETANGENTIAL ON-BOARD INJECTIONSYSTEM -- RELATED TESTING INTRODUCTION In a supporting test, the flow and pressure characteristlcs of the blade tan- _, gential on-board injection ( TOBI ) system were assessed to verify the design assumptionsand overall efficiency of the pressure-balanced system in the EnergyEfficientEnginehigh-pressure turbinecomponent.Testingwas conducted with an availableturbinedevelopmenttest rig at the Middletown, Connecticut Test Facility.

_ t TEST CONFIGURATION To maintain the integrityof the rotating instrumentation and s!ip ring, no c h ang es were made to the existingrig disk. The disk hole entrancesof the rig disk were at a radius of 34.29 cm (13.5Uin), and the remainingvortex chamber " .

dimensions were scaled from this point holding radius ratios constant.Thus, _i the TOBI radius in the rig was 28.605 cm (11.262in), as indicatedin Figure L_ A-I. A limitationimp o s e dby t h e di s k,which does not fit the modeling,occurs :i at th e vo rte x p l ate b olt circle.The net effect was a 40 percentreductionin :_ available flow area.

TEST PROGRAM To effectively model the TOBI conditionsin the integratedcore / low spool at !

se a le v eltake o ff, three parameterswere held constant.These includeda TOBI exit Mach number o f 1.05,an air-to-diskvelocityratio of 1.68 and a radial- to-tangential air velocityratio of 0.027. In order to satisfy these condi- ti o ns, th e rig sp e edwas m aintainedat 6580 rpm with the TOBI n o zz l e pressure ratio set at 2.0. In addition,the TOBI supply air temF.erature was established to minimizeheat transfereffectson disk temperatures.

An investigation of off-designconditions was l imitedto variousTOBI pressure ratios with a constant di s k speed. Pressure ratios ranging from 1 .2 to 2.5 were tested.During these tests, TOBI inlet temperature was varied to maintain minimumdisk temperature gradients.Table A-I shows the test conditionsestab- lishedf o r t h e rlg p rogra m . I i v O RIG I NAL P AG E _S i O F P OOR QU AL I TY i DIMENS IO N i cm (in) f f _ " i 25 . 4 0 _'_ i ( 10. 00) - , ! 27. 5 4 8 -_! (10.846) a : . 2 8 6 05 25 .9 0 8 • (11.262) (10.200) ( ) (10.825)_ ";_ r" 1 • I 29 . 083 (11.450) 2.212

_°:i_ 1 27.495

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'4 TABL E A-I TEST RIG PR O GRAM PressureRatio Acro ssTOBI RPM TOBI InletTemperature °C (°F) 1.2 0 37 (lO0) 1. 4 0 3 7 (I 0 0 ) 1.6 0 37 (lO 0) 2.0 0 37 (lO0) 2.5 0 37 (lO0) 2.0 0 37 (lO0) 1.2 65 80 lOl (214) 1.4 6580 ll3 (236) 1.6 6580 12l (250) 2.0 658 0 1 3 2 ( 270 ) 2.5 6580 132 (270) 2.0 6580 132 (270)

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9 3 _, '1 T ABLE B-I F ULL S T A G E T URBINEWARM RIG T ES T RESULIS PHASE I TEST* PRESS SPEED POINT RATIO PARA.

1 4.Of 244.8 2 4.03 234.1 3 3.9 1 222.3 4 3.9_ 253.1 i 5 3.4 9 24 4 .4 6 3 .48 2 3 5.5 7 3.4_ 223.0 I 9 4.43 245.3 i i PHASE 3 i 1 8 4.03 244. 6 21 4.02 24 4 .2 28 4 .Of 244.8 i PHASE 4 22 4. 30 2 3 4.4 2 3 3.4 7 25 3 .8 2 4 4. 0 4 2 44 .8 2 9 4. 7 5 24 3 .8 i * Fo r each te s t po int, e x it t o ta l p res s ure ( in HgA), s tatic pre ss ure (in HgA), air angle ( s upp l ementary ang l es),and t o ta l temperature ( ° F),and efficienc y are pre s ente d .

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3 6 S i w _ w • 0 20 40 60 80 LO0

%8PQN

Figure B-7I Aver a ge Tota l Pr e ssur e V e rsus S p an (Point 2 Q)

i 1 66 J ' t • , (, : 28.B O { _B t

/

Z 8. 4 0 '1 _ 8. o0 Z' / .6 0 _ , CO ( 1 - Z " / ,2 0i 26.8 0 1 _

, /

!

2 5 -B O! , , ', , , , 0 2Q 40 60 80 LO O XSPA N F i§ u r e B-7 2 Av e r_ : geStatic Pr essu r e V e rs us Span (P o int 29) 167 i _ !

I \ Z 6 0 ?,50 a.

24 0 l e i _ : g30 L ' 2 _0 F-- _ " ' " 2 1.0 !

i , 200 i t g o tS O " # . ' t O , , , , , ¢ 0 20 4 0 6 0 e O lO0 : %SPRN FigureB-7 3 A ve rag e Total Te mpera tu r e V e rsu s Span (P o int29) i

" X

._ 168 _'$ .... , , .. _ - _i . .,'',, L B O @ t S O 1 4 0 _ ' 13 0 W 2 " 1 - 2 0 • { r io 1 0 0 J

I

90 i, _0 w ! ! w ! ! IL 0 2 . 0 4 0 80 8 0 L O 0 %SPAN Figure B-74 Av e rage Air Angle Versus Span (Point 29 ) --" 169 "_ "

t '

( D ,, 7e I 0 20 40 60 iX ) 100 ?

PERCENT SPAN Figur e B-75 E ffici e ncy Ve rsus Span (P o int 29 ) - 1 7 0 ' _ ' _t _y2 -- _ _.

. . _ _ i l " l , I Y

| TABLE B - II HI r -H PRESSURETURBINECASCADETFST RESULTS L PHASE I TEST* MACH . : POINT _UMBER • o 1 0 , 70 2 0 .85 3 0.91 _ 4 1.0 5 l.08 := 6 I. 19 _. 7 0,56 :' PHASE 2 ; 1 _: B 0 , g 2 9 0.92 ', *For each test point, the exit total pressure (in ' HgA), static pressure (in HgA), air angle, vane loss (APT / PT), and pressure distributionare presented, I .

i x

.__ 171

-" i_. / J : 4Z .BO_ 4 1.6 0_ I 3 9 .6 0! 1 3 9.20 ( , • _ , , , 0 ; t O 4 0 6 0 BO t OO

XSPRN

Fig u re B-76 A ve rag e T o tal P r e ss u re V e rsus Span (P o int 1)

, .k

172 ,_

I

% _ r ? 32,00_ _ r _ : 3 1 .6 Ol 3 1. 20 ( 3 0 .8 01 i_l 3 o . 4o , A ;_ 3 0 . 00 1 i 2 9.6 0 . ; ; 2 9. 2 0 ' 2 8.8 0 v 2 8,4 0 , .....

0 ZO 4 0 6 0 6 0 100 XSPF:IN ¢ F i gu r e B- 7 7 Av e rag e StaticPr e ssureV e rsu s S p an (P o int 1) A

173 _ X

........ ,P)J

k

?

J II = I0 x b. J C_ rr i ?

0 £0 40 60 eO 1 00 XSPRN F i gu r e B -7 8 A ve rag e Air A ngl eVer s u s Span (Point 1) C , ,t • i _ 174

'I , }

Fig u r e B- 7 9 Va ne L oss V e rsusSpan (P o intI) t , 175 ' _,_ c, r -- . .......... L" / J; Ioi00" 1.000" 0.900 ® I B 0 o 8 00 " • , _ ¢ 0 . 7 00" ® ® " • _ ®®® o .. ® ®® E l • _ .

0 3 £ n 0 . 8 00 " = .

} _L " 0. 5 00 " !

0 ° 400 - 0= 3 00" t 0 . _00 " _ .....

I 0 2 0 4 0 BO e O t O O

% X I BX

Fi gu r e B-80 Va ne Pr ess ure D istributicn at II Pe r c ent Span ( Point I)

i

i

i76

% ' !.1 00 _ , 1 .000 " E] 09 ® E] • ® 0 . 9 00 " ® 5] 0 . 8 00 ":" ® ® ® N ® " i _ .._ 0 . 7 0 0 - _ ® n 3 : , i m :- _ a -0. 8 00.

, 0 . 500- i _. 0 .4 00 " 0 .3 0 0 0 . 2 00 .....

0 £ 0 4 0 fl O fl O 10 0

X X / BX

F igur e B - 81 V ane P res s ur e Di s trib u ti o n at 50 Pe rc e nt S p an (P oint 1) $ J ,

i X

'1 177 _ : _ !

I' I .1 00 ' 1 .000" B ] [] ® ® o ® 0. 9 00 ' [] ® O, e O 0- ,. ® i ® [] [] . _ 0 . 7 30. _ ® ® ® ® G o mO , BOO - 0 .5 00- 0 . 400 " 0. 3 00 " 0.g 00 .....

0 g O 4 0 60 8 0 1 00

i % X / BX

F ig u r e B-82 Van e P ressur e Di s tri b ution at 89 Pe rcent Spa n ( Point I) ' 1 7 8

_ N

, J , ' !

• 5 2 ¢ 5t X +J I L l ; ; 47 :. : • : 4 6 4 5 0 Z O 4 0 6 0 80 L O0 _ $ PAN F igur e B- 8 3 Van e A ve rag eTo talPressur eVe rsusS p an ( P oint2) +:I i

-

I 3 2 o 40 1 3 ?. . 001 i

//

31 . 6 0 1 / ( _ 3], , Z OI i _i . 3 0 .8 01 , L co . - 3 0 .4 0 1 30.001 ¢ 2 9.601 29.20 1 0 Z O 40 60 80 t OO % $P RN Figure B-84 Vane AverageStatic PressureVersusSpan (Point2) 1 2 X L _ r r - C _ 9 i I '7 i i v m . y : 0 20 40 60 B O tO 0 Z S PRN F i gure B - 8S Vane A v e c a ge Air Angle Versus Sp-:n (Po i nt Z ) 181 _ k _

i r . ]

0 . 180 " " b 0. 1 6 0 - 0 .140" t 0. 1 20- ( 1 3 cO O. l O0 c3 _ J (_ 0.080" 0 . 0 6 0 " , i o.o4 o , i 0. 0 20" ' : t 0.000 ...... i I 0 20 40 6 0 80 t O0 7. SP RN Fi gu r e B-B6 Van e L oss V e r s usSp a n (Po int2 ) r_ 182 ( ,. .- ; I I.I00° I . 000 " [] r_ ® B i 0 .9 00 " _, 0 .8 00 "

I _

I .-- ,, n t .,O 0 . ?00 " ® ® ¢L 0"800 " ® ® ® t 0 . 5 00 J 0 . 400' 0 . 300 " I 0 . 200 .....

0 2 0 4 0 6 0 8 0 10 0

% X / BX

F i gu r e B- 8 7 V an e Pr e s s ur e Distributi o n at I I Perc e nt Span ( P o int 2 ) i 183 |a'_,_l I.I00 " I ,000 " El E l 0 . 9 00 " _ ) i ® [] , z 0 .8 00 " !

0 . ' 7 00 - ® I .- - \ ® } n o .6 00 " _ ) ® _ ) ® , 0.50 0" - I 0. 4 0 0" 0. 3 00 " 0 . _00 .....

0 _0 4 0 6 0 8 0 1 00

% X I BX

Figure B-8 8 Van e P r e ssur e Distributi o n at 50 P e rcent Span (Po int 2 ) 184 ' i , I°I00 " I. 000' E] E] ®

I ® ®

. _ o . s o o .

E] ® I 0 . 70 0' I-- , I 0 . 800- i O_ E _ ! : CO ® ® El ® El : °- 0.600" _ ® 0. 4 00' _ o.3oo- i

• i

0. 2 0 1 .....

0 20 4 0 60 80 10 0 % X I BX Figur e B- 8 9 Vane Pr es sur e Distribution at 89 P e rcentSpan ( Point2 1 i

i

185 i . , _ _ 5B 5 9 ° 4 " 7 I 4 8 ......

0 _0 40 6 0 B O I00 : %SPRN , i FigureB- 9 0 Van e AverageTotalPressureVersusSpan (Point3) J 1 8 6 , , ,j 3 4 f.

f ": 3 0 ¢ 1 2 9 _-_ " 28 i 27 26 , , , _ , , 0 20 4 0 6 0 B O 1. 00

%S P QN

Fi gu r e B- 91 V an e A v erag e Static Pr e ssur e V e rs u s Span ( P o int 3)

j, i, i_ 187 _!

I ,!

t8 t6 Z , . ' t O ' 6 2 ¼ v w v v v 0 2 0 4 0 6 0 8 0 tO0

%$PF N

Figu r e B - 92 Vane Ave ra ge A _r Angle Ve r s u s Span (Point 3) ' 188 _' 0.000 , , ;_ i i T !

0 2 0 4 0 6 0 8 0 100 %SPFIN Fi g ur e B-93 Va n e Los s V e r sus Span ( Po int 3) 18 9 _i

@

I 1. 1 00" 1 . 000 " E] ® B 0 .9 00" ® 0 .B O O " El , = . _ 0 .? 00- ® no .6 0 0 - _ - ®

®e e®®_ i I

0 .5 00 " ® I

I

0. 4 00" 0 , 3 00" i .

0 .2 00 ..... i 0 20 40 60 8 0 10 0

% X l BX

F ig u r e B - 9 4 V ane P r e s s ure Di s tributi on at 11 Pe rc en t S p a n ( P o int 3) 1 90 ' , _ ' ,J ,

l i

I.I00" 1 .0 0 0' "El [] _ . _ ® [] 0 .9 00 " ® ® [] , ii 0 .80 0 - 0 .? 0 0 " ® O- ® C O n o .6 0 0' ®

®[] i

t 0 .5 00 " _ , ' 0 . 4 00 " . , 0 . 300 " 0. 2 00 .....

0 2 0 4 0 6 0 80 I O0 % X / BX , :

Fi gu r e B - 95 Va neP r essu r e D i s tri bu ti on at 50 Perc e n t S p an(Po tnt 3) i

i 191 , k,

i

1.1 0 0 " 1.000" E] ® ® 0 , 900" [] ® 0 .8 00" ® 0. 7 0 0" 0 3 nO . 6 0 0' El ® El -- ® ®®® :i ® I 0.5 0 0" , I 0 .4 00 " 0 .30 0 " !

0 . 2 00 ..... i 0 2 0 40 80 BO 1 0 0

% X / BX

i

Fi gure B-96 Vane Pr essu r e Distributi o n a t 89 P e r ce nt Sp an( P o i n t 3)

6 4

i

t_ 6 O

i

I t

56 /

Q-- : • !

5 2 _ j ; g ' 5 0 !

,8 !

46 _ -t ....

0 2 0 4 0 6 0 80 t O0 %S P R N Figur e B- 9 7 V a n e Av era g e Tota lP res s ur e V e rsu s Span (Polnt 4 ) Q 1 93 | 3 4 3 3 J 2 " ¢1 3 I i !

i," 30 !

2 0 _ .

l _ 7 p ; • ! ; I v 0 20 40 60 80 tOO X S P R N Ftgure B-98 Vane Ave r a g e Statt c P r e s su r e Versus S pan (Po i n t 4) ' % e e . _ k 4 1 2O f I r L s J r r : , , _-

{

t 8 - I , 4 _ w ! ! ! T 0 2 0 4 0 6 0 8 0 100 %SPAN • F i gu r e B - 99 VaneAve r age A ir Angle Ve r sus S pan ( P o i n t 4) i 195 L_ I _ J tJ.

0.1 8 0 - 0 . 1 6 0 " 0.1 4 0 ' 0.120] _ o3 o30.I00 - ¢ _0.080 "

° l '

!

!

0.0 6 0 "

i

0.0 4 0 " !

0.020 - I.

0.000 t .....

0 20 40 80 8 0 100 %SPFIN F i gur e B-IO0 Vane L oss Versus Span ( P o tnt 4 )

_!

19 6

. %

...... __. ____ , ,.},, 0 ,300 " I

I o . _ oo ' ' ' ' ' I : '

0 2 0 4 0 8 0 O 0 L O0 _

X X / B X !

Fi g ure B - 101 Va ne P r essu re Di s trib u ti o n at 11 P e r ce nt S p a n ( P o int k ) 1 9 7 _ • il m .

"_ ! . I 0 0 - I . 0 00- B [] ® [] 0 .9 00 - ® ® E] : 0. 8 00 " 0 . 7 0 0" ® O_ ® U 3 °-O .B O0 - ® _0 0. 5 00 " ® ® @ ®® : 0.40 0' 0,300 " , 0 . 2 00 .....

0 _0 4 0 B O S O 1 00 , I

% X / BX

Figure B-102 VanePressure Distribution at 50 Percent Span (Point 4)

| , It

l

198 " ,., !

I I i ! . I 00 - :*" ' " 0 .9 00" T ® . 0. 8 00 - c .

-_ ® ; F __ 0" ? 00 " _ !

_ °"0. 6 0 0 " ®

•_ o. soo- • • i

® 0. 4 00 " 1 i ,; 0 ° 3 00" i 0.20 0 .....

0 2 0 4 0 ri o 8 0 1 0 0

. . X / BX

Figure B-]O_ Vane Pressure Distribution at 89 Percent Span (Point 4) _ _ 199 ......

, L 70 6 8 : : 6' 2 x _ . ILl I'--- Q,..

6O 5_ w i v ! ! v

o zo 40 60 n o z oo

XSPQN

F ig u r e B - 104 V a ne A ve ra geTo ta l P r e s su r e Ve rs us S p a n ( Po i n t 5 )

.. . -I";_, f 3 S I 3 3 L 3 2 co 11- !

3 1 !

i ' . !

3o ! !

2S i i,

z8 i

L 27 v , , , , , 0 2 0 40 6 0 BO I00 I XSPAN i t FigureB - lO5 Vane AverageStaticPressureVersusSpan ( Point 5) 2 0 1 _ _ 4 r ! ! ! I !

0 20 40 8 0 O O t O 0

7 .SPRN

Fi g ure B-f06 Van e Av e rage Air Angle Versus Span (Point 5) " 20 2 ' ,'$,'_ A I

. r_2

t

I .

_ . o.36o

i o.32o I 0 .28 0 " T " J 0 . 2 4 0-

J _

c oo . 2 00 ' C D

I

._J I L l _o . 1_o .

,t 0 . 1 2 0 '_ 0.0 8 0 " 0.0 4 0 ' _ C fr • 0.000 ....

I i 0 2 0 4 0 6 0 _ : 0 IO0 Z$P R N Fig u r e B - 1 0 7 Va ne Lo _ .sVe r su s S pan (Po int 5 )

i

"i

i k 1 I . I 0 0 ' | ' I . 000 " [] B ® 0. 9 00 " ® 0 .BOO ' [] S 0 . 7 00 " _-- ® 13_ C o ® °-0 .6 00 " ® 0. 5 00 " ® ®® [] O . a O0 " ® ® !

0 .3 00 " 0.200 .....

0 20 40 6 0 8 0 t O O

% X I BX

Figur e B-I0 8 Van e Pr es sur e Distributi o n at II P e rc en t Span (Po int 5 )

- ._ 2 04

I.I00- J 1.000- _ B ® B '_ 0 . 900 " ® O . BO0" 0.?00' ® k -- c o °-0.80 0 ' ® J 0 . 5 00 ' ® B ® ®®® 0 . aO 0" !

0 . 3 00 " • t i 0 . 2 0 0 .....

0 20 4 0 80 80 I00 t . % X / BX ' F i gu r eB - I09 Van e P r ess ur e Di s tributi o n at 5 0 Pe rc e ntSpan (Po int5 )

i

' I . I00" I.000 " I _ [] ® 0 .9 00 ] _) B 0.8001 i , _ , | L , ® 0. 70 0- o ...

° '- 0 .6 00" ® i ] !

B r ;1 0.5 00 " ® ®® 0 .4 0 0 - ' , 0.300 " { L 0.2 0 0 .....

0 2 0 4 0 80 8 0 1 0 0 ..

% X l BX t I F t gur e B-IIO Va n e P r essu r e Dis tr i b u ti on a t 89 Pe r cent . Sp an (Poi n t 5) 206 x ' !

! f ?2 CL

e o !

t 56 _ 5 Z i 4 B ......

0 20 40 $0 BO 100

' X SPR N I

F t g u re B-]|" V a ne A v era g eTo t a] P re ss ur e V ersus Span( P oint 6)

I

_ . 35 - 33 ¢ 3 0 -. _.

r 29 _ 21) of .

29 _ _ 6 , • v , • " ' ,' 0 20 40 60 3 0 I OU

_. S PA N : _

F i gure B-1 12 Van e Av er a g e St a tic Pre,_sure V er sus S p an (P o int 6)

. 208

L

w_ r 2 4 'i i 2 0 _ .

18 • Z !

' ° i

8 • ! T ! ! ! !

0 2 0 4 0 6 0 8 0 IO0

%SPRN _

Fi g ur e B -11 3 V a neAve r a g e Air A n g l e Ver sus Span ( P o int 6 )

, 209 _ '

. _ .... - ------ -, • ])a i

i 1 0. 36 0 0.320" !

0 .28 0 " I . 0.240 " ,_ co o . 2 D O " i A _ • 0. 16 0 " ?

_ g 0.1 2 0 " J !

l

oo oo i i

_j 0 . 0 4 0 ....

n n N¢l 0 2 0 4 0 6 0 80 tO0 !

%SPAN _L F i gu r e B-II4 Vane Lo s s V e rsusS p an ( Po int 6) i ' w r ' 2 1 0 , , 0. 2 00 , , • T r / 0 2 0 40 6 0 80 t O0

Z X / BX

F ig u r e B- II5 Van eP r ess ur e D i s tributi o n a t II P e r ce nt Span ( P o int 6 ) _ ; 1.10 0 " 1 . 0 0 0 " E] E] ® E] 0 .9 00 ® 0 .8 00 - ; 0 . ? 0 0 ' ® CO _ _ ' no . 60 0 _ 0 .50 0 " E] _ .

® ; .

0 . .t OO " _ ) ® _ 0. 3 00 ' 0.200 .....

0 2 0 4 0 8 0 8 0 t O 0 ;_ % X I BX ,' F ig u r e B- ]] 6 Van e Pr ess ur eD istri b uti o n at 5 0 P e rc e ntS p an ( P o int6 ) '_ 212

|

I.I00 I.000 _0 {3 ® ® ® 0 . 9 00 - ® E] 0 . BOO - !

® 0.70 0" O ....

co no . 6 00 " : " L @ ) ) 0.500 - @ ) _ I _ t 0. 4 00 " ® j _: 0. 3 00 " -_ 0 .201 .....

0 20 40 6 0 B O t O0 t

% X / BX

'i - Fi g ur e B- 11 7 Van e P r e s s ur e D istributi o n at 89 P erc e nt Span (P o int 6 ) !

3 8 .4 0 1 3 B .OOf 37. 6 01 37. 2 0{ !

3 8 . 8 01 X b J O _ 36. 4 0 ( I 36.00I i

1 '

3 5 . 6 01 35. 2 01 3 4 .8 0 1 ......

0 Z O 4 0 60 8 0 100 %SPQN Ft gu r e B - 118 V a ne A ve r a g e T o t al P r essu r e Ve r su s Sp a n ( Po i n t 7) 214 _ ' i I 3 1,0 0 _ ?

i 3 0 . 8 0t 3 0 ,6 0t 3 0 . 40{ : 3 0. 20 t _ _ o"- _ 3 0 . 00 (

,i

• 29.8 0 ( 29.6 0 ( _,.

29.4 0 1 2 9. 20_ ...... i 0 2 0 4 0 6 0 8 0 I00 ZSPRN F ig ure B -1 1g Vane Average S tati c P ressure Versus Sp a n (Po i n t 7) '_ i 21 5 t3 I [ t2

i

i l

" d'

t O X h.J ¢'r '7 4 ! ! • ! ! 1 0 EO 4 0 60 8 0 100 XSPAN Fi gu r e B - 12 0 Vane Ave r age A ir Ang l e Ve r sus Span ( P o i n t 7) !

I

I

2 1 6 0 . 0 9 0" :, 0 . 080 " - _ " " 0 . 070 - 0.0 6 0 "

_o . oso-

_d u.l Z o .040 " 0.0 3 0 " !

O.OIO" 0.0 2 0 i 0.000 ' , .....

0 20 4 0 6 0 8 0 t O0

%SPRN

F i g ur e B-1 2 1Van e L oss V e r s u s Span(P o i n t 7 )

( I.I 00 " ' 1 . 000 - B B ® 0 .9 00_ ® 0 .8 0 0" ® ® 0 . 70 0 " - I--- O _ C O no .6 0O " h 0.500 " 1

(

0 , 400 " i

i

[

0 .3 00" I

t

o .z oo ..... {

0 Z O 40 6 0 8 0 lO 0 I % X / BX ',

i '

Fi gu r e B -1 22 Vane Pr es s u r e D i s tri bu ti on at ll P e r cen t Span ( P o i n t 7 ) 21 8 I.I00 0.800 ® 0 ._ 0 0 ....

r 0 2 0 4 0 3 0 B O 100

_. X / BX f

ii Figure B-12 3 Van e Press u r e Distributi o n at 5 0 Per ce ntSpan ( P o int7 ) f

j

1.100" 1 , 000" [] [] ® ® ® ® 0 .9 00 " 0. 8 00"_ ® ® El ®®® E] 0 .'7 00 " 0 3 0"0. 6 00 " 0 .5 00 " 0 . 400 " 0 .3 00 " O . ZO0 .....

0 2 0 40 ri o 8 0 tO0 % X / BX Figur e B-124 Va ne P r essu r e Dis t rib ut i on a t 89 Pe rc en tSpan ( P o in t 7 ) . , 220 k j 9O 8B x n 8 2 7 8 i 7 6 '_ ' 7 4 , " .... , ,_ 0 2 0 4 0 60 BO t OO " _ XSPRN . _ F ig u r e B-125 Van e A ve rag e To tal P ressure Ver s us S pan (; - o int 8) ' i 4 5 5 5 4 5 , cO 6_ 4 8 4 ?

_ 6 w • v • • v 0 20 40 6 0 8 0 t00 XSPRN Fig u reB- ] 2 6 Vane AverageStaticPressureVersusSpan (Point8) 22 2 ,.2 4, 0 2 0 ,_0 6 0 80 100 '_

Y.SPAN {

,i

,!

Fig u r e B- 1 27 Van e A v era ge Air Angle Versus Span (Point 8)

l

L 0.180" l 0.160 0.140' 030.100- 0 30"120" I C :3 ._.J U.J c_:O .080- > 0.060" 0.0 4 0" 0.0201 '-'; 0 . 000! ......

0 20 40 60 80 I00 XSPAN Fi gu r e B- 1 28 Van e L oss V e r s u s S p an (P o int 8) 524 , F i gu reB- 1 29 Va n ePr ess ur e Distrib u ti on at 11 P e rc e nt S p an ( P o int 8) i' q l 'i' Ioi00- 0 .9 0_ ® O.BO0 0 . 700 " b--- eL ® mo . 6 oo- ® ®® N ® ® ® 0 .5 0 0 " 0.400- 0 .3 00 ' 0o_00 t .... t 0 _0 40 6 0 80 I00 i

% X / BX

t Figure B-1 30 Vane Pressure Distribution at 50 P e rcent Span (Point 8 ) i , I.I00" I.000" E] 50 ® ® ® 0 .9 00 " rzl ® 0. 8 00_

®

0 . '7 00 " I- -- rt 0 3 ® , n o. 6 00 " O0 ® ® 0 .5 0 0 " i b.1 0 ,4 0 0 " 0 .3 00 " 0 . 2 00 .....

0 20 4 0 60 fl O 1 00 % X / BX Fig u r e B-1 3 I Va ne Pr e ssure D istributi o n at 89 Perc e nt Span ( Po int 8 ) i t 1 1 2

to8

10 6 ......

0 2 0 4 0 6 0 8 0 tO0

XSPAN

Fi gure B -13 2 VaneAve ra geT ota l P r e s s u r e Ve r s u s S pa n ( P oi nt 9)

22 8 X

I _ _ _ _ , * " _. : ..- , ,.

_ 75 Q..

_o _i 68 " , - 2 ,

/ '

66 , _ 0 2 0 4 0 6 0 BO t OO %SPAN Figure B- 133 Va ne A ve rag e Static PressureVersusSpan (Point9) 2O t8 1 6 1. 4 t 2 > < b J O_ Q _ tO B 1 v t 1 i 1 0 2 0 4 0 6 0 B O t OO %S P AN F ig u re B - 13 4 Van e Av e rag e Air Angl e V ers u s Span (Po int9) 2 3 0 -_ 0. 16 =" F igur e B-1 3 5 Van e Lo s s Ve r s usSpan (P o i n t9) : l ] 231 e , 1.100" 1.000" ® ® 0 ,9 00 - ® O.BO0 " [] 0.700 - • Q_ \ 0 0 3 0 -0. 8 00 • • 0 o i_ 0. 5 00 - 0. 4 00 " 0. 3 001 • 0 .£ 00 .....

0 2 0 4 0 6 0 8 0 1 00

% X I BX

Figur e b-136 Van e Pressure Distributi o n at II P e rc e nt Span(P o int 9 )

2 3 2 m ......... I . ,':ij, , _,

t

I . I00 I .O 00 - El ® E] 0 .9 00 - ® ® 0.800 I ___ 0. 70 0 ® i ® a"o . 8 oo- ® , p ® ®® F]

. I

®® i , 0 . 5 00" i O . aO 0" I 0.300- . _ 0 . 2 0; , , , I -

o _' o 4 0 6 0 s o t oo

X X / BX

- 1

Fi g ure B-1 3 7 Van e Pre s sur e Distributi o n at 5U Percer_t Span(Point 9 )

J

23 3 I L i ! _ _.

; o- ID D

O

' O

O

0.9 -- [ _ ,

O

0.8 --

o 0 . 7- I

0 0

0,6 -- C] ,

,, •

0.5-- + t

o 4 - I-I i

0.3 -- O + : "

1 I _ I i , I

020 20 40 60 80 100 120 PERCENT SPAN

i+

Figure B- 138 Van e P re s sure D i s tribution at 8 9 P e rc e r0tSpan ( Po int 9 ) 2 3 4 LIST OFSYMBOLS ANDABBREVIATIONS A annulusarea AA air angle ACC activeclearance control Bx axialchord BOAS bladeouter air seal Cx axialflow velocity CET combustor exit temperature EEE EnergyEfficientEngine EGT exhaustgas t_nperature FP flow parameter, W _ / TT / P T FPS flightpropulsion system HPT high-pressure turbine ICLS integrated cere / lowspool ID innerdiameter LE leadingedge LPT low-pressure turbine m flow . , M Mach number ;{_ Mn Mach number N nmchanical speed,rpm OD ou ter diameter P / PT staticto total ? ressureratio . _ PR totalto totalpressureratio _ Ps staticpr e ssure . .

PT totalpressure . -_ R radius RIT rotor inlettemperature Tc coolanttemperature TF film te mp erature TG gas temperature _' TTET trailingedge TEC Tangential On Board Injection Exit Cavity total temperature _ _ TOBI tangential on-boardinjection TR temperature ratio ._ Tm metal surfacetemperature W flow Wc / a totalcoolingair flow Wae engine a irflow ab s ab s o l ute U tangenti al wheel speed - - i p d e ga s tu rning absolu teair ang l e relatlveair angle Z_ d e l ta n efficiency I : REFERENCES , 1 Gardner.W. B. : "EnergyEfficient EngineHigh-Pressure T urbineUncooled Rig T e chnology Report",(CR-]65149); October]979.

2 Thu l in,R. D., Howe D. C., and SingerI. D. : "EnergyEfficient Engine High-Pressure TurbineDetailDesignReport",(CR i 65608); January]982.

3 Kopper.F. C. eta]: "EnergyEfficient EnginpHigh-Pressure Turbine Supersonic Casca(eTechnology Report",(CR-165557); November]981.

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