NASA CR-1 79548 RI/RD 8 -214 94-08574 - DTICAD-A277 095 NASA CR-1 79548 RI/RD 8 -214 94-08574 FINAL REPORT Design and Experimental Performance of a Two Stage Partial Admission Turbine

AD-A277 095NASA CR-1 79548RI/RD 8 - 214

94-08574

FINAL REPORT

Design and Experimental Performance of aTwo Stage Partial Admission TurbineTask B.1 / B.4

byRocketdyne Engineering 4 QT1 -

•E LEC,_l I1

Rocketdyne Division SMAR.u1994 IRockwell International W

prepared for

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

December 1992

NASA-Lewis Research Center 00

Cleveland, Ohio 44135

Contract NAS3-23773

G. Paul Richter. Project Manager

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FOREWORD

The work presented herein was conducted from December 1984 to June 1986 by

personnel from Engineering functional units at Rocketdyne, a division of Rockwell

International, under Contract NAS3-23773. Mr. Dean Scheer, Lewis-Research Center,

was the NASA Project Manager. At Rocketdyne, Messrs. Anthony Zachary, Program

Manager, and Robert F. Sutton, Project Engineer, were responsible for the direction of

the program.

Important contributions to the conduct of the program, and to the preparation of the

report material were made by the following Rocketdyne personnel:

Advanced Turbomachinery Projects Mr. Dan Shea

Mr. Edmund Roschak

Hydrodynamics and Aerodynamics Mr. Jim Boynton

Mr. Lou Rojas

Structures Mr. Linsey Orr

Dynamics Ms. Linda Davis

Engineering Development Laboratory Mr. Brad King

Mr. Bill Bubel

Mr. Joe Lichatowich

Mr. John McPherson

Mr. Ted Davis

Ms. Dawn Elliott

Mr. Robert Herbig

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NASA CR-179548

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TABLE OF CONTENTS

FO R E W O R D ...................................................... ............................................................ i

U S T O F F IG U R E S .............................................................................................................. i i

UST OF TABLES ........................................................................................................ v

S U M M A R Y ......................................................................................................................... 1

INTRODUCTION ........................................................................................................ 4

O B JE C T IV E S ...................................................................................................................... 6

TECHNICAL DISCUSSION .............................................................................................. 7

MK49-F TURBINE DESIGN ............................................................................... 7

TESTER DESIGN AND DESCRIPTION ................................................................. 1 6

ROTORDYNAMIC ANALYSIS ............................................................................. 21

FACILITY DESCRIPTION .................................................................................. 30

INSTRUMENTATION ......................................................................................... 34

DATA ACQUISITION AND ANALYSIS .................................................................. 37

TEST MATRIX AND PROCEDURES .................................................................... 39

TEST RESULTS .................................................................................................................. 41

DESIGN ADMISSION CONFIGURATION ................................... 48

Design Configuration Efficiency Characteristic .................................. 48

Design Configuration Flow Characteristic ........................................... 53

Design Configuration Static Pressure Distribution ........................... 56

Design Configuration Turbine Rotor Axial Thrust ............................... 6 66

HALF DESIGN ADMISSION CONFIGURATION .................................................... 69

Half Configuration Efficiency Characteristic ...................................... 69

Half Configuration Flow Characteristic ............................................. 73

Half Configuration Static Pressure Distribution ............................... 73

Half Configuration Turbine Rotor Axial Thrust .................................. 73

QUARTER DESIGN ADMISSION CONFIGURATION ............................................. 80

Quarter Configuration Efficiency Characteristic ............................... 80

Quarter Configuration Flow Characteristic ........................................ 84

Quarter Configuration Static Pressure Distribution ......................... 87

Quarter Configuration Turbine Rotor Axial Thrust ............................ 87

CONCLUSIONS ................................................................................................................... 92

REFERENCES ..................................................................................................................... 94

APPENDIX A Standard Air Equivalent Conditions ....................................................... 95APPENDIX B Raw Data, Tests 8-10 (Pressure Distribution) .................................. 97APPENDIX C Data Reduction Technique ...................................................................... 131APPENDIX D Raw and Reduced Data, Tests 11-13 .......................................................... 140

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LIST OF FIGURES

1. Program S chedule ................................................................................................ . . 2

2. MK49-F Liquid Hydrogen Turbopump .................................................................. 5

3. Blade Path Mean State Conditions at Design .......................................................... 9

4. Velocity Vector Diagram ........................................................................................ 1 0

5. MK49-F Turbine Cross Section ............................................................................. 1 1

6. Blade Path Design Data ........................................................................................... 1 2

7. First and Second Stage Nozzle Angular Orientation ............................................... 1 3

8. Two Stage Partial Admission Turbine Blade Coordinates ...................................... 1 4

9. Performance Characteristics ..................................................................................... 1 5

10. Two stage, Partial Admission Turbine Tester ........................................................... 1 8

11. First and Second Stage Nozzle Turbine Wheels .................................................... 2 22

12. Tester Components ................................................................................................ 23

13. Second Stage Nozzle Angular Orientation Motor Drive Setup ............................... 24

14. Load Path Diagram ................................................................................................ 25

15. Rotordynamic Critical Speed Plot ........................................................................ 2 6

16. Rotor Group Mode Shape - 1st Rotor Mode ........................................................... 2 7

17. Rotor Group Mode Shape - 2nd Rotor Mode .......................................................... 28

18. Facility S chem atic .............................................................................................. . . 3 1

19. T ester Installation .............................................................................................. . . 3 2

20. Tester Disassembly ............................................................................................. 3 33

21. Instrumentation Parameter Locations .................................................................. 35

22. Static Pressure Tap Locations ............................................................................. 36

23. Data Acquisition Flow Chart .................................................................................. 3 8

24. Efficiency vs. U/CO, Zero Degrees, Three Arcs .................................................... 4 3

25. Equivalent Flowrate vs. PR (eq) for Arcs of Admission ...................................... 4 44

26. Interstage Static Pressure Distribution ............................................................. 4 6

27. Efficiency vs. U/CO, Design Configuration ......................................................... 4 9

28. Design Configuration Second Stage Nozzle Orientation ....................................... 5 0

29. Delta Temperature Efficiency vs. Angle Orientation, Design Configuration ..... 5 2

30. Equivalent Flowrate vs. PR (eq), Design Configuration ...................................... 54

31. Equiv. Flow Parameter vs. Angle Orientation, Design Configuration ................. 5 55

32. Interstage Static Pressure Distribution at Design ............................................. 58

33. Static Pressure Distribution, Design Configuration, 0 Degrees ........................ 59

34. Static Pressure Distribution, Design Configuration, +40 Degrees ................... 6 1

35. Static Pressure Distribution, Design Configuration, '30 Degrees ................... 6 2

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NASA CR-179548RI/RD 86-214

36. Static Pressure Distribution, Design Config., 0 Degrees, 10,052 RPM ................ 64

37. Static Pressure Distribution, Design Config., 0 Degrees, 25,003 RPM ................ 6 5

38. Axial Thrust From Pressure Loads, Design Configuration ................................. 6 7

39. Axial Thrust From Pressure Loads, Design Configuration, High Pressure ........ 6 8

40. Efficiency vs. U/CO, Half Design Configuration ................................................. 7 041. Delta Temperature Efficiency vs. Angle Orientation, Half Design ........................... 72

42. Equivalent Flowrate vs. PR (eq), Half Design Configuration .................................. 74

43. Equivalent Flowrate vs. Angle Orientation, Half Design Configuration ................... 7 544. Interstage Static Pressure Distribution at Half Design ...................................... 76

45. Axial Thrust From Pressure Loads, Half Design .................................................. 7 8

46. Axial Thrust From Pressure Loads, Half Design, High Pressure ....................... 7947. Efficiency vs. U/CO, Quarter Design Configuration ............................................. 8 1

48. Delta Temperature Efficiency vs. Angle Orientation, Quarter Design ................ 83

49. Equivalent Flowrate vs. PR (eq), Quarter Design Configuration ........................... 8 5

50. Equivalent Flow Parameter vs. Angle Orientation, Quarter Design ......................... 8 6

51. Interstage Static Pressure Distribution at Quarter Design ................................ 8 88

52. Axial Thrust From Pressure Loads, Quarter Design ........................................... 9 053. Axial Thrust From Pressure Loads, Quarter Design, High Pressure ................. 9 1

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LIST OF TABLES

1. Design Requirements for the MK49-F Turbine .................................................... 7

2. MK49-F Turbine/Tester Design Parameters ............................................................ 1 7

3. Nozzle Admission Fractions .................................................................................... 1 94. Turbine Tester Parts List ...................................................................................... 2 0

5. Test Peak Efficiency for Arc of Admission .............................................................. 4 1

6. Turbine Axial Thrust From Pressure Forces ........................................................ 4 7

7. Target N (eq) and PR (eq), Design Arc of Admission ........................................... 5 1

8. Interstage Static Pressure Distribution Ratios, Design Configuration ................ 5 7

9. Target N (eq) and PR (eq), Half Design Configuration ........................................ 7 1

10. Interstage Static Pressure Distribution Ratios, Half Design Configuration ..... 77

11. Target N (eq) and PR (eq), Quarter Design Configuration .................................. 82

12. Interstage Static Pressure Distribution Ratios, Quarter Design Configuration .. 89

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NASA CR-179548RI/RD 86-214

SUMMARY

The Rocketdyne Orbital Transfer Vehicle (OTV) rocket engine system's high pressure,

liquid hydrogen turbopump was designed with a two stage, partial-admission axial

turbine. The turbine is basically two single stage, partial-admission, subsonic impulse

stages designed so the kinetic energy leaving the first stage rotor is discharged directly

into the second stage nozzle at nominal operation to minimize staging losses. Very little

data was available in the literature for this type of turbine design. Therefore, the

decision was made to test a full-size model of this turbine design using ambient-

temperature gaseous nitrogen as the working fluid.

The effort conducted herein was sponsored by the Space Propulsion Technology Division,

NASA Lewis Research Center, Cleveland, Ohio, under Contract NAS3-23773, "Orbit

Transfer Rocket Engine Technology Program." The program began in December 1984

and continued through June 1986. Figure 1 presents the overall program schedule.

The tester design features a rotatable, remote-controlled, variable orientation second

stage nozzle system which changes the first to second stage nozzle angulation during a

single test period and allows adjustment to the optimum position for highest

performance. In addition, this feature minimizes test turnaround time and maintains

good test-to-test performance correlations. Nozzle arcs of admission were changed by

plugging, or unplugging, a discrete number of nozzles with silicon rubber. Problems

with the retention of these plugs during the test program were solved by epoxying the

plugs in place.

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NASA CR-179548RI/RD 86-214

A total of thirteen tests were conducted in the Rocketdyne Engineering Development

Laboratory starting in September 1985 and continuing in three phases until April

1986. Second stage nozzle orientation angles from +40 to -30 degrees from the

designed nozzle angular orientation (40 degrees from the center of the first stage nozzle,

in the direction of rotor rotation) were tested. Arc of admissinn variations for the first

stage nozzle were from 37.4 percent (10 nozzle passages - 5 per side) to a low of 6.9

percent (2 nozzle passages - 1 per side). The second stage arc of admission varied from

a high of 84.4 percent (26 nozzle passages - 13 per side) to a low of 12.9 percent (4

nozzle passages - 2 per side).

Performance of the turbine at design conditions was approximately 6.7 percent higher

than originally predicted, which was probably attributable to a higher predicted turbine

windage loss. Effects and trends of nozzle arc of admission variation were generally as

expected with the lowest performance coincident with the lowest arc of admission. In

addition, large deviations from the design nozzle orientation produced the lowest

performance. Turbine pressure ratio variations (1.3 to 2.0) had little effect on the

overall turbine performance.

The data generated during the test program substantially verified the performance

prediction methods used at Rocketdyne. Minor nozzie to nozzle angular orientation

changes could be made to attain the highest performance of the MK49-F turbine.

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NASA CR-179548RI/RD 86-214

INTRODUCTION

Partial admission turbines are used in applications such as low thrust rocket engines

where low volume flow rates and flow areas in the turbines require very small blade

heights for full admission (360 degree) designs. In the 15,000 Ibf thrust expander

cycle engine concept defined by Rocketdyne under NASA/Marshall Space Flight Center

(MSFC) contract NAS8-33568, partial admission was used in the drive turbines for

both the oxygen pump and the hydrogen pump (Reference 1). The turbine for the oxygen

pump was a "conventional" single stage, impulse type, partial-admission design based

on experimentally verified analytical codes. On the hydrogen side, a two-stage turbine

with partial admission in each stage was shown by analysis to provide the highest

efficiency. However, a lack of substantiating empirical data to support the analyses

flagged the turbine as one of the technology issues in the hydrogen turbopump concept.

Subsequent to the engine point design effort conducted under the MSFC contract,

Rocketdyne initiated company-funded detail design, analysis, and fabrication work on the

major components required for the engine system. An isometric view of the high

pressure hydrogen turbopump (MK49-F) resulting from this work is shown in Figure

2. As noted in the figure, 1,704 horsepower was required to drive the three stage pump

which delivered 418 GPM at a discharge pressure of 4671 psia. In view of the lack of

empirical data for the turbine, the NASA Lewis Research Center (LeRC) issued a Task

Order under the Orbit Transfer Rocket Engine Technology Program (contract NAS3-

23773) to experimentally determine performance and establish a data base for future

designs.

Task Order scope consisted of design, fabrication and test of three two-stage partial

admission turbine configurations using ambient temperature nitrogen gas as the working

fluid. The three inch diameter turbine utilized the exhaust housing and rotors fabricated

under the company-funded effort. Flow passages in the first and second stage nozzles

aerodynamically reflected the MK49-F turbine design, but could be blocked to change the

degree of admission for a given test. In addition, the second stage nozzle could be

circumferentially rotated during a given test to determine the effects of second stage

nozzle angular orientation on turbine performance. Results of the Task Order effort are

reported herein.

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OBJECTIVES

The overall objectives of this program were to (1) verify the two stage partial-

admission turbine analytical predictions by conducting laboratory tests using ambient

(room) temperature gaseous nitrogen, (2) update analytical performance prediction

methods for future designs of similar low thrust engine turbines, and (3) provide

baseline data for comparison with the OTV MK49-F turbine for possible performance

enhancements with only minor hardware modifications.

The technical approach to reduce cost was to use all available MK49-F turbine

hardware, such as the turbine and exhaust housing, and design most of the other tester

hardware using aluminum material to minimize design and fabrication complexities.

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NASA CR-179548RI/RD 86-214

TECHNICAL DISCUSSION

MK49-F TURBINE DESIGN

Design requirements for the MK49-F turbine were derived from the cycle power

balance for the engine and are shown in Table 1. The turbine speed of 110,000 rpm

was set based on optimization of the hydrogen pump performance. The working fluid was

hydrogen gas at a flow rate of 3.73 Ibm/sec with a total pressure and total ,emperature

at the inlet flange of 3830 psia and 881 degrees Rankine, respectively. Total pressure

at the turbine outlet flange was 2173 psia.

Table 1. Design Requirements for the MK49-F Turbine

Working fluid Gaseous HydrogenPower output, hp 1,704

Speed, rpm 110,000Flowrate, lb/sec 3.73

Inlet flange total temperature, R 8 81

Inlet flange total pressure, psia 3,830

Outlet flange total pressure, psia 2,1 73

Pressure ratio 1.763

Specific work, btu/lb 323

Efficiency, percent 63.6

The MK49-F turbine was designed assuming an equal available energy split between

stages. Design point pressures and temperatures at the blade path mean diameter are

shown in Figure 3. Design point velocity diagrams at the mean diameter are shown in

Figure 4. The turbine inlet manifold total pressure loss was set at 5 percent of the

flange-to-flange total pressure drop resulting in a first stage nozzle inlet total pressure

of 3,747 psia. The first stage rotor exit absolute velocity pressure was 3 percent of the

overall total pressure drop and 25 percent of that was considered available to the second

stage. The outlet manifold total pressure loss was set equal to the second stage rotor

outlet absolute velocity pressure.

A cross-sectional view of the turbine is shown in Figure 5. The rotor blades were

shrouded to reduce tip clearance loss. Shroud operating radial clearance was 0.005 inch

for each stage. A close clearance interstage seal under the second stage nozzle provided

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NASA CR-179548RI/RD 86-214

low bypass nozzle leakage and also served as a rotordynamic stabilizing factor. Blade

path design data are given in Figure 6. The nozzle and rotor blade height for each row

was constant from inlet to outlet. Rotor blade heights overlapped the nozzle outlet

heights and were set to provide the flow area requirements for the gas at the outlet of the

respective nozzle arc of admission. The axial space between the nozzle trailing edge and

rotor leading edge for each stage was set at 0.05 inches to minimize nozzle outlet flow

spillage into the axial space with the associated energy loss at the end of the arc of

admission. The axial space between the first stage rotor trailing edge and the second

stage nozzle leading edge was set at 0.20 inches to minimize circumferential gradients

into the second stage. The benefit of a large axial space between stages was reported in

Reference 2 from tests of a two stage turbine with a partial admission first stage and a

full admission second stage. Two admission arcs spaced 180 degrees apart were used in

each nozzle. Total admission fractions were 0.345 (5 passages per arc) and 0.452 (7

passages per arc) for siage 1 and stage 2, respectively. The circumferential position of

the second stage nozzle center relative to the first stage nozzle center was set at 40

degrees in the direction of rotation (Figure 7) such that the absolute velocity vector

from the first stage rotor would discharge directly into the second stage nozzle.

Blade shapes and blade profile coordinates were defined at the mean diameter and are

presented in Figure 8. Non-twisted, constant section blading was used because the

blade heights were small and previous Rocketdyne designs with similar constant section

and small height demonstrated high performance. As noted in Figure 6, prime numbers

were selected for the number of rotor blades (53 first stage, 59 second stage) in order

to minimize blade frequency problems.

Predicted design and off-design performance and flow characteristics for the MK49-F

turbine are shown in Figure 9. At the design point velocity ratio of 0.286, an

efficiency of 63.6 percent was predicted.

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Figure 4

VELOCITY VECTOR DIAGRAMMK 49-F TURBINE

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First SecondStage SecondNozzle NozzleStage Stage

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Figure 7 First and Second Stage Nozzle Angular Orientation

MK 49-F TURBINE

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PERFORMANCE CHARACTERISTICSMARK 49-F TURBINE

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NASA CR-179548RI/RD 86-214

TESTER DESIGN AND DESCRIPTION

In order to provide a comprehensive data base for future designs, program scope

included gaseous nitrogen testing of a full-scale model of the MK49-F turbine plus two

additional configurations with smaller admission fractions. Design point parameters for

the MK49 -F turbine are listed in Table 2 along with standard air equivalent conditions

as defined in Appendix A. Also presented are speed, flow rate and pressure ratio

magnitudes that simulate MK49-F turbine design point operation in the tester with

nitrogen gas as the working fluid at representative test inlet conditions. Note in Table

2 that the Reynolds number of the MK49-F turbine was not duplicated in the tester. In

Reference 3, the effect of Reynolds number on turbine performance was increased

performance with increased Reynolds number and the effect was negligible for Reynolds5 5

numbers greater than 2 x 10 . Thus, the tester Reynolds number of 8.05 x 10 was

high enough to preclude Reynolds number effects on performance. Turbine performance

on the engine was predicted equal to or better than the GN2 test performance because of

the higher Reynolds number.

A cross sectional view of the tester, Rocketdyne P/N 7R0017780, is shown in Figure

10. Where possible, components from the Rocketdyne in-house MK49-F program were

used in the tester. These included the turbine rotors, the exhaust housing, and rotor-to-

shaft attachment hardware (the exhaust flow guide shown in the figure was also a

MK49-F component, but was not used in the testing reported herein). The first and

second stage nozzles for the tester were fabricated with larger admission fractions (i.e.

more nozzle passages) than the MK49-F turbine and the required admission for each test

configuration was obtained by blocking passages with Room Temperature Vulcanized

(RTV) silicon rubber. Stage 1 and stage 2 nozzle admission fractions for each of the test

configurations are given in Table 3. A key feature of the tester was the rotatable second

stage nozzle used for in-test changes in the angular orientation of the second stage nozzle

with respect to the first stage nozzle. This concept provided a cost effective method to

evaluate first to second stage nozzle circumferential relationships without excessive

down time between tests. A positive second stage nozzle angle change from the design

position was opposite the direction of rotor rotation. The tester was also built with two

quartz windows to provide capability for laser velocimeter viewing of the flow at the

inlet of the second stage nozzle. However, tests using the velocimeter were not conducted

because of a reduction in Task Order scope. Polyimide turbine tip shroud and interstage

seals were used in the tester to prevent damage to the MK49-F turbine rotors in the

event of a rub. The tester parts list is shown in Table 4.

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NASA CR-179548RI,,RD 86-214

Photographs of the first stage nozzle, the MK49-F turbine wheels, and the second stage

nozzle are presented in Figure 11. Note that Figure 11 is a composite of individual

photographs which were taken at different object distances (approximate scales are

shown above the component photos). Figure 12 is a photograph of all the tester

components. Figure 13 is a closeup photograph showing the second staga nozzle

angular orientation motor drive setup and the instrumentation tube bundles emanating

from the nozzle housing.

ROTORDYNAMIC ANALYSIS

The two stage partial admission turbine tester was analyzed from a rotordynamic

standpoint to determine the rotor critical speeds, stability, and response to unbalance

forces. Following an iterative process, analysis indicated that a tester design with a

1A5 inch shaft diameter between duplex bearing sets could adequately meet the design

criteria of operation outside of the 20% critical speed margin. Response analysis

indicated low steady-state bearing loads and rotor displacements during operation.

Stability analysis showed the rotor to be stable to speeds greater than the operating

speed range.

A detailed finite element model was developed which represented the stiffness, mass, and

geometric properties of the rotor. Figure 14 presents a diagram of the rotor

corresponding load path.

Damped natural frequencies and corresponding mode shapes were calculated. Results are

shown in Figures 15 through 17. These curves represent the rotor critical speeds

with damping and cross-coupling forces present as well as turbine Alford forces.

Observing Figure 15, at a speed of 31,000 RPM, the nearest rotor critical speed

occurred at 47,747 RPM. This resulted in a 35% separation of the maximum operating

speed from the rotor first critical.

Since the seals on the turbine tester produced destabilizing cross-coupling forces, the

turbine tip Alford forces were expected to be present, therefore the possibility for

unstable rotor whirl to develop existed. For this reason, a stability analysis was

conducted. The results of the analysis showed that the rotor was stable for all speeds

analyzed. This was the expected result since the rotor operated below its lowest critical

speed.

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Figure 14

Load Path DiagramMK49-F Turbine Tester

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Qui11 Shaft NodaL PLocement. Turbine Tester

R•.aL Coord.natee

-25-

NASA CR-i 79548RI/RD 86-214

Figure 15ROTOROYNAMIC CRITICAL SPEED PLOT

rmc 49-r TURB8INE TESTER - DAMPED CRLTICR. SPEEDSCIESIGN -DUPLEX SLRVC BAGS. 1.45 IN SIW'T DIR BETWEEN BEFIW35G

*~ ~ . .. .. , . . . . . .

.............C

- -~. .14 .. . . .*

ROO SPNSCE P

* C C-2 6 -

NASA CR-179548RI/RD 86-214

Figure 16

Rotor Group Mode ShapeM~ 19-f rLIRINr =6t - M¶WM CRITICkL iPMS

OESICN OLX &UME ML~ . 1.45 IN WOO"lF CIA ncibcr M

1ST ROTOR MODE47747. RPMI

QUILL SHAFT

BEARING 1,2

•BEARINGS 3. 4

ifTURBINE END

STRUION I ORB!T - FrORw9R3 PRECCSSION

-27-

NASA CR-179548RI/RD 86-214

Figure 17

Rotor Group Mode Shapeit 49-r TLumiNt 1TECR - CN1 ITICk. if=s

IDLSICN - ItPo .RC SLV , 1.415 IN 354" CIR aBcrW MS

2ND ROTOR MODE61227. RPM

QUILL SHAFT

EARINGS 1, 2

\ -. BEARINGS 3, 4

TURBINE ENDSTATION I MRIT - -ORWNARD PRCCCSSION

-28-

NASA CR-179548RI/RD 86-214

A linear response analysis was also performed to assess the bearing loads and rotor

displacements as a function of rotor spin speed. For an assumed 1 .0 gram-inch residual

unbalance, the maximum radial load over the tester's operating range was 250 lbs. at

the turbine end bearings. The maximum rotor displacement was .0025 inches peak and

occurred at the second stage turbine disk. A more realistic unbalance of 0.20 gram-

inches scaled these values to 50 lbs. and .0005 inches peak. Note that these results were

for steady-state operation.

The possible addition of squeeze film dampers to the rotor assembly were found to lower

the overall response, but also shift the peak response into the operating range of the

tester. Thus, the maximum response of the rotor with squeeze film dampers produced

higher loads than without the dampers over the tester's operating range, and therefore

were not included in the design.

-29-

NASA CR-179548RI/RD 86-214

FACILITY DESCRIPTION

Testing was conducted at Rocketdyne's Engineering Development Laboratory (EDL)

Rotary 1 Test Facility, which is shown schematically in Figure 18. Nitrogen gas for

the test turbine was supplied from high pressure storage through a servo controlled

valve which regulated turbine upstream pressure during testing. After expanding

through the turbine, the gas flowed through the exhaust control valve and was discharged

to atmosphere. Power generated by the turbine was absorbed by an electric

dynamometer that was coupled to the turbine through a 10:1 speed increasing gea1; •ox,

quill shafts, and a brushless torque meter. In order to accelerate the turbine up lu test

speed without an excessive amount of time or turbine drive gas, the dynamometer was

used as the prime mover. Figure 19 is a photograph of the tester installed in the

facility. Figure 20 shows the start of disassembly in the test cell for nozzle arc of

admission changes between tests.

-30-

NASA CR-179548>- RI/RD 86-214

-i I-

a: CLw u

00

U I

U)U

LU 0 LU 4

-31 CC-O


0t

-32

NASA CR-179548RI/RD 86-214

0

CD-3 3

NASA CR-179548RI/RD 86-214

INSTRUMENTATION

In addition to the instrumentation necessary to determine overall turbine performance,

interstage static pressures were measured at selected axial and circumferential

locations to determine pressure distributions within the test turbines. Location of the

instrumentation is shown in Figures 21 and 22. The number designations shown

adjacent to the static tap locations refer to the instrumentation channel numbers listed

in Appendix B.

Shaft speed was measured using an eddy current proximity probe that sensed the rotation

of a 0.005 inch, 40 degree arc depression that was machined on the tester shaft. Turbine

shaft torque was measured with a torque meter as shown in Figure 21. However,

torque data was suspect because of problems with the torque meter. Overall turbine

performance reported herein was based on first stage nozzle inlet total-to-turbine

discharge flange static conditions. Inlet total pressure was measured in the stagnation

region of the manifold 180 degrees around the inlet manifold from the inlet pipe as

shown in Figure 22. Turbine discharge pressure was measured with a four hole static

pressure piezometer ring attached at the outlet flange parting line. Inlet and discharge

temperatures were measured at the respective pressure measurement locations.

Nitrogen gas flow rates were measured upstream of the inlet flange with a sonic flow

nozzle. Because of the wide flow range, different size flow nozzles were necessary for

each of the three test configurations. Efficiency characteristics of the test turbines were

calculated from the measured temperature drops across the turbine.

Static pressure taps within the turbine, which inciuded first and second stage nozzle

flow channel hub and tip taps at the inlet and discharge locations, are shown in Figure

22. The channel taps were positioned at the midpoint of the channel in the plane of the

respective nozzle blade leading or trailing edge. Taps were also located on the upstream

and downstream surfaces of the nozzle block at a three inch diameter, approximately 90

degrees from the centerline of the arcs of admission as shown. The second stage nozzle

inlet and outlet cavity pressure taps shown in Figure 22 were located on a two inch

diameter.

Additional data parameters necessary for safely conducting the tests were monitored in

real time in the facility control room. These included the redline parameters, facility

status, and test condition monitors.

-34 -

NASA CR-179548RI/RD 86-214

CC,0 D mO~

-J 0>I.

LZUCC~ z m

U) CL C1

LU 8

ui W -jC, CCZC

0 ~~~ CL-u0NzuCC NP -ý-i CL0 -0 Z ( Lac

mi UJz z LL -35--


04

-- o

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U, U,

N U,

- U

0 -U-~ U~ --

.0-m. IWO

SN 0 IL =(nS ..

-~ 5- SmSN z

0)-36-

NASA CR-179548RI/RD 86-214

DATA ACQUISITION AND ANALYSIS

A flow chart showing the steps involved in the acquisition and analysis of the test data is

presented in Figure 23. Only steady-state operation data were recorded. The

individual data parameters were recorded on a measurement group System 4000. The

System 4000 recorded the transducer output voltage data on floppy disks which were

later used as the database for a scaling program. The scaling program converted the

voltage data onto another set of floppy disks in engineering units (pressure,

temperature, torque, etc.). Since the System 4000 utilized a Hewlett Packard computer

which was not IBM compatible, the scaled data was transmitted to the IBM directly via an

RS-232 port line.

The Lotus 1-2-3 spreadsheet program was used to reduce the scaled data to performance

parameters. It was also used to produce a majority of the performance plots. An

engineering generated macro program was written that automatically read the scaled

data, averaged it, and then stored this averaged test data in the format required for the

performance reduction program. The test data required averaging because each test slice

consisted of 10 to 20 scans at one steady state test condition.

The performance program was also an engineering generated macro program that

automatically read the averaged data and performed iterative calculations for

performance parameters. These performance parameters were used to produce

performance plots.

Real gas properties were used in the analysis of the test data. The curve fit equations of

the property data along with the equations used in the data analysis are given in

Appendix C.

The flowrates were obtained by applying the ideal isentropic critical flow equation

across the nozzle. The specific heat ratio used in the analysis was obtained by a curve fit

equation !hat characterized the real property gamma as a function of temperature and

pressure. The total pressure and temperature were obtained by iterating the Mach

number at the nozzle inlet. The mass flowrate was calculated, from the ideal isentropic

critical flow equation, with the Mach number equal to 1.0 at the nozzle throat. The total

pressure was initially assumed equal to the measured static pressure. Applying this

initial assumption, a Mach number at the nozzle inlet was calculated. A new total

-37-


~~dc

Ozz

LIE...0 2 cc ~r -

WIL ~I

00.

0~ * 4 IAI

-I I4ow C

CLw

z 1 0

U-A

W(J UUOd. Uj

LL x a -K L'c

'a occ-38-

NASA CR-179548RI/RD 86-214

pressure was then calculated. This process was repeated about 3 times before any

further change occurred.

The turbine efficiency was obtained by dividing the actual enthalpy drop across the

turbine by the isentropic enthalpy drop.

hin(act) - h out(act)Efficiency =

hin(act) - h out(isen)

Two curve fit equations of real gas properties were used to calculate enthalpy; one as a

function of temperature (T) and pressure (P), and the other as a function of pressure

(P) and entropy (S). A curve fit equation was also used for the entropy calculation, as a

function of pressure and temperature. The pressure and temperature measured at the

inlet were used to obtain the turbine inlet enthalpy. The pressure and temperature

measured at the outlet were used to obtain the actual outlet enthalpy. The turbine outlet

pressure and calculated inlet entropy were used to obtain the turbine outlet isentropic

enthalpy.

TEST MATRIX AND PROCEDURES

Testing was conducted with nitrogen gas at turbine inlet manifold total pressure and

temperature of approximately 215 psia and 480 degrees Rankine, respectively. Target

test speeds were 50, 70, 100, and 120 percent of the design equivalent speed. At each

test speed, data was recorded at target pressure ratios of 1.3, 1.6, 1.74 (design) and

2.0 with the second stage-to-first stage nozzle angular orientation set at the design angle

(+40 degrees as shown in Figure 22) and at -30, -10, +10, and +40 degrees (+30

degrees for the "quarter design" configuration) from the design angle. Positive second

nozzle angle rotation was opposite the rotor rotation direction.

The turbines operated by setting inlet conditions and motoring the turbine to the test

speed with the electric dynamometer. Test pressure ratio was set by adjustment of the

exhaust control valve with the turbine output power absorbed by the dynamometer. At

each test pressure ratio, the second stage nozzle was remotely rotated to each of the angle

settings noted above.

-39-

NASA CR-179548RI/RD 86-214

Static pressure checks were performed before and/or after the main portion of each test.

The system was pressurized to approximately turbine inlet pressure (200 psig) with

the discharge valve closed and the shaft at or near zero speed. The pressure transducers

were checked to verify that all read nearly the same value. Significantly lower pressure

values indicated leaks in the instrumentation lines that would have caused erroneous

readings during the test.

Three test series involving a total of thirteen tests were conducted accumulating

approximately 36 hours of run time on the rotor assembly. The first series (seven

tests) included checkout and what was intended to be performance testing of the design

arc of admission configuration. However, a problem with nozzle plug retention caused

by inadvertent reverse motoring of the turbine, along with instrumentation problems,

precluded the use of the first series test data. In the second series (tests 8, 9, and 10),

each of the turbine configurations was tested at the target speeds, pressure ratios, and

nozzle angle settings noted above. After completion of test 10, the thermocouples used to

measure gas temperatures at the sonic flow nozzle and at the turbine inlet manifold were

found to be malfunctioning. Subsequent flow checks failed to provide temperature data

that could be used with confidence, and therefore a third test series (tests 11, 12, and

13) was conducted to determine turbine performance. Since the pressure data from the

second series was valid and relatable to the performance tests because of similar inlet

conditions, the level of instrumentation in the third test series was reduced to only what

was necessary to determine turbine overall isentropic efficiency.

The second and third test series were run without the exhaust flow guide installed. The

flow guide performance effects were to have been evaluated in the first test series, but

the plug retention and instrumentation problems noted above made evaluation

impossible.

-40-

NASA CR-179548RI/RD 86-214

TEST RESULTS

Efficiency test data for the design admission configuration, the half design admission

configuration, and the quarter design admission configuration are shown for the 0 degree

second nozzle orientation angle in Figure 24. Smaller admission resulted in lower

peak efficiency, and peak efficiency occurred at lower velocity ratios. The peak

efficiency and velocity ratio values are listed in Table 5.

Table 5. Test Peak Efficiency for Arc of Admission

(0 Degree, Second Stage Nozzle Orientation)

% Admission Peak Velocity Ratio (U/CO)

Configuration 1st stage 2nd stage Efficiency O%1 •Peak Efficiency

Design 34.5 45.2 68 0.30

(10 of 29) (14 of 31)

Half Design 13.8 19.4 47 0.24

(4 of 29) (6 of 31)

Quarter 6.9 12.9 36 0.18

Design (2 of 29) (4 of 31)

The equivalent flow test data for the three nozzle arcs is shown in Figure 25. Similar

trends of increasing flow with increased pressure ratio are shown for each admission.

The spread with equivalent speed was small for design admission and became smaller as

admission was reduced.

The test data analyses included determination of the turbine efficiency characteristics,

flow characteristics, internal pressure distributions, and turbine axial thrusts. All

admission configurations and second stage nozzle orientation tests were conducted

without the exhaust duct flow guide installed. Comparisons were made of the test results

and the design predictions. The results are presented in terms of standard air inlet

condition equivalent parameters: equivalent flow (Weq), equivalent speed (Neq),

equivalent pressure ratio (PReq), and efficiency (TI). These parameters are described

-41 -

NASA CR-179548RI/RD 86-214

in Appendix A. The equivalent pressure ratios were determined from the first stage

nozzle inlet total pressures and the elbow duct outlet flange static absolute pressures.

The performance test results included the efficiencies and flow characteristics. These

characteristics for the three arcs of admission found in Table 5 were determined from

test 11 for the design admission, from test 12 for the half design admission, and from

test 13 for the quarter design admission. The test data and performance parameters for

these tests are listed in Appendix D. The efficiency characteristics were calculated

from the measured temperature drop across the turbine because of problems with the

torque meter data.

-42-

NASA CR-179548RI/RD 86-214

z

Po

0

0 0

00

00,•z o

• r- 1

Sacoo

zraz000

Il . .j L |, , . ..... , .,. l l ~ , , l, , , , i.. , . |, i . . I , ,. . l 0

Z w

3•Ifl±V•IHdN3. Vnaa0 AJN-3IIZ A.H-

-43.

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NA:SA CR-179548RI/RD 86-214

on 0- -Lfl

oZ -NJ

Lii -

LLAJ

LU -

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U) U

%00

-44-

NASA CR-179548RI/RD 86-214

The turbine pressure distribution data and the turbine axial thrust calculated from the

pressure distribution data were obtained from test 8 for the design admission

configuration, from test 9 for the half design admission configuration, and from test 10

for the quarter design admission configuration. These tests were found to have erroneous

temperature measurements and, therefore, were not usable for the performance

determinations. Test data for tests 8, 9, and 10 are listed in Appendix B.

The interstage static pressure distributions for the three configurations are shown in

Figure 26 for the design equivalent conditions. Higher overall pressure ratios were

used for the half and quarter configurations because of the higher second-to-first stage

nozzle area ratios. The inactive arc pressures (triangle symbols) were equal across the

first rotor as expected with the inactive rotor blade flow passage areas equalizing the

pressures. Positive pressure drops are shown in the active arcs for the both tip and hub

pressures across the first rotor and across the tip of the second rotor. Slight positive

rotor blade reaction is shown for good performance in the active flow region. The higher

pressure differences between tip and hub at the exit of the first stage nozzle and second

stage nozzle reflect the high nozzle outlet swirl compared to the nearly axial rotor outlet

swirl and nearly equal hub and tip pressures. Higher first stage pressure ratios are

shown for the half and quarter configurations reflecting the higher second-to-first

nozzle area ratios. These configurations require higher overall pressure ratios than

were tested in order to produce more equal stage pressure ratios and power splits.

Turbine rotor axial thrust values were calculated using measured static pressure

averages from each interstage location (Figure 26) for the three configurations at the

design equivalent conditions. The turbine axial thrust values are shown in Table 6,

which includes the turbine axial thrust ratios (turbine thrust divided by turbine inlet

pressure). The turbine axial thrust for each engine condition was determined by scaling

the test pressure ratios by the engine turbine inlet pressures and calculating the thrust.

Turbine thrust was defined as positive in the direction of turbine through-flow.

-45-


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X~ -z ZU2~< E- 0 D >rz

cr Wr'-o1 Z<

zu DECC -. U z:

0< E L

co C-2 u u

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0 z~ w - E)E--0 0 0 t0 04

-46

NASA CR-179548RI/RD 86-214

Table 6. Turbine Axial Thrust From Pressure Forces

(POSITIVE AXIAL THRUST IN DIRECTION OF TURBINE THROUGH-FLOW)

CONFIGURATION

PARAMETER DESIGN HALF QUARTER

TEST AXIAL THRUST, LBf -120 -109 -122

AXIAL THRUST/INLET PRES., LBf/PSIA -0.56 -0.51 -0.57

ENGINE AXIAL THRUST, LBf -2109 -1912 -2118

In the sections that follow for each configuration, detail figures are included that show

the locations, pressures, areas, and forces for the two-stage partial admission turbine

axial thrusts that are summarized in Table 6. The application of the pressure forces

and the general trends are discussed below.

First stage rotor disk and blade axial thrust was calculated from pressure forces acting

on the area from the shaft seal and interstage seal diameters to the rotor tip diameter.

Second stage rotor axial thrust pressure forces acted on the areas from the interstage

seal diameter to the rotor tip diameter on the upstream side and from the tip diameter to

the shaft centerline on the downstream side. The blade annulus area was defined from the

blade hub diameter to the rotor tip seal diameter. The diameters were taken from the

rotor drawings. The average of the inactive arc pressures was applied over the inactive

arc rotor blade faces and over the disk faces when the chamber pressure was not

measured. The average active arc pressures were applied over the active arc rotor blade

faces.

The disk and blade pressure forces were nearly balanced for the rotors with partial

admission. Most of the turbine axial thrust resulted from the unbalanced pressure on

the downstream side of the second rotor over the area from the interstage seal diameter

to the shaft centerline. This pressure force was directed upstream, opposite the turbine

through-flow, and is listed with a negative sign in Table 6. These test results show

that partial admission stages provided the lowest axial thrust across blade and disk

surfaces of any staging option.

-47-

NASA CR-1 79548RI/RD 86-214

DESIGN ADMISSION CONFIGURATION

The design admission configuration consisted of a first stage with 34.5 percent admission

and 10 of 29 full ring nozzles flowing, 5 per half 180 degrees apart, and a second stage

nozzle with 45.2 percent admission and 14 of 31 nozzles flowing, 7 per half. The second

stage to first stage nozzle admission ratio was 1.310.

Design Configuration Efficiency Characteristic

The efficiency characteristic for each of the 5 second stage nozzle orientation angles is

shown in Figure 27 versus velocity ratio. Highest efficiency is shown over most of the

range for the +10 degree orientation, and the lowest efficiency for the -30 degree

orientation.

The design second stage nozzle orientations with test rotations for the maximum angles

are shown schematically in Figure 28. The +10 degree nozzle orientation, which

related to a 30 degree angle between the center of the first nozzle inlet arc of admission

and the center of the second stage nozzle inlet arc of admi3sion in the direction of

rotation, provided the highest efficiency (not the design 40 degrees as expected). This

indicated that the flow velocity through the first stage was higher than predicted.

The efficiencies and equivalent flowrates for the targeted design equivalent speed -f

21,219 RPM and pressure ratio of 1.765 for the 5 second stage nozzle orientation

angles are listed in Table 7 with the design equivalent prediction. The test efficiencies

at design conditions were plotted in Figure 29 versus second nozzle orientation, along

with the design prediction. The design orientation test efficiency was 68.6 percent

compared with 63.6 percent for the design prediction. The test value was 6.7 percent,

4.3 percentage points higher than predicted, for the test measurement locations of first

stage nozzle inlet total and outlet flange static pressure. The test efficiency adjusted to

the design locations of inlet flange total and outlet flange total pressure was 67.9

percent. The efficiencies for the +10 degree and +40 degree second nozzle orientations

were 68.8 and 68.6 percent, respectively. High test efficiency was shown over a wide

range of second nozzle orientation angles.

-48-

NASA CR-179548Ri/RD 86-214

I- -4

L= .

cmJ

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~iii

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NASA CR-i /9548RI/RD 86-214

W W

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ata

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-50-

NASA CR-179548RI/RD 86-214

Table 7

TARGET N (eq) AND PR (eq)DESIGN ARC OF ADMISSION

2-N LINE N (eq) PR (eq) FLOW (eq) EFFICY Um/CoANGLE, NUMBER (ETA)

DEGREES TEST 11 RPM LB/SEC

0 PREDICTED 21219 1.7651 0.0744 0.636 0.286

-30 45 21246 1.780 0.0751 0.638 0.289

-10 51 21599 1.767 0.0760 0.677 0.295

0 57 21690 1.787 0.0767 0.686 0.294

+10 63 21526 1.766 0.0769 0.688 0.294

+40 71 21320 1.761 0.0766 0.686 0.292

-51 -

NASA CR-179548

RI/RD 86-214

[11111 0 0n

H W

F-~

-i LL, m

S~r5

HLL N

____ LU±IJJ±LWIU

cmJ

-52-

NASA CR-179548RI/RD 86-214

The slight efficiency decrease for the +40 degree orientation indicated that +40 degrees

was beyond the optimum, which was apparently between the +10 and +40 angles for the

design admission configuration. Nozzle orientation was shown to significantly affect

efficiency. An increase of 5.0 percentage points or 7.8 percent was shown in efficiency

from -30 to +10 degree orientations.

Design Configuration Flow Characteristic

The equivalent flowrate versus pressure ratio for the targeted design equivalent speed

test points are shown in Figure 30 for each of the five second stage nozzle angles, along

with the predicted values. The equivalent flow for the test points targeted at the design

equivalent speed and pressure ratio from Table 7 are shown in Figure 31 versus

second stage nozzle orientation, along with the design prediction. The equivalent flow at

the 0 degree design orientation was 3.1 percent higher than predicted. At +10 degrees,

the equivalent flow was 3.4 percent higher than the design prediction. As with the

efficiency characteristics, all the test equivalent flow values were higher than the design

predictions. The equivalent flow decrease from the +10 to +40 degree orientation

indicated that +40 degrees was beyond the optimum which was between +10 and +40

degrees, as with the efficiency characteristic. The increased flow from the -30 to +10

degree orientation was 2.4 percent, which was less than the efficiency increases.

A test value for equivalent flow within three percent of predicted is considered close

agreement for this small-size turbine with a 3-inch mean diameter, considering the

difficulties involved in fabrication and inspection of the small nozzle, rotor passages,

and throat openings.

-53-

NASA CR-1 79548RI/RD 86-214

*-4 R

Iin~

Cf)C

U),

= 0-CDJ LD

<i LUO

U, U

Li 0 .

0J/2 2LCiAiN1~fO

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-55-

NASA CR-179548RI/RD 86-214

Deslan Configuration Static Pressure Distribution

The measured static pressures were averaged at each of the locations listed in Table 8.

The local absolute static pressures were divided by the average inlet manifold test

pressure to form pressure ratios. The test pressure ratios were compared with the

predicted pressure ratios in Table 8. The test and predicted pressure ratios were also

plotted in Figure 32. Good agreement was shown between the test data and the

predictions. The largest delta (difference) ratio in Table 8 was 0.039, or about 5

percent, for the first stage nozzle outlet hub pressure, which showed good agreement

with predictions.

The test pressure distribution results confirmed that the design staging achieved a near

equal available energy split between the two stages for high overall performance.

Similar magnitude effects of the high nozzle outlet tangential swirl (tip pressures

higher than hub pressures) are shown in Figure 32 for both first and second stage

nozzle active arcs and for both test and prediction. Positive test pressure drops shown

for tip pressures across both rotor blade rows and for hub pressures across the first

stage rotor indicated positive reaction in the active arcs for good performance. Near

equal pressures were present in the inactive arcs across both rotors. This resulted in

low blade and disk axial thrust loads, characteristic of partial admission staging.

The circumferential variations of a large number of measured static pressures were

analyzed using absolute pressure ratios of the local static pressure divided by the inlet

manifold average pressure to normalize the data. A plot of the test pressure ratios for

the design configuration at the zero degree second stage nozzle angle is found in Figure

33, which shows the distributions for the first and second stage nozzle inlets and

outlets. Data shown in Figure 33 were for the instrumentation that survived the

fabrication, build, and installation phases of the program.

The first stage nozzle inlet pressure ratios were shown with a larger scale than the

others in Figure 33. The inlet manifold pressure ratio wa,. 1.00, but the inactive arc

pressure ratio at 180 degrees from the inlet pipe indicated a slightly higher value

(1.015) due to the stagnation effect of being opposite the inlet pipe location (0 degrees).

The inactive arc pressure ratio at the location where the inlet pipe entered the manifold

was a value of 0.976. The hub values were slightly higher (1.3 percent) than the tip

values due to the curvature of the inlet manifold into the first stage nozzle. These

differences increased with flowrate (overall pressure ratio) from 1.3 to 2.0. Pressure

-56-

NASA CR-179548RI/RD 86-214

Table 8. Interstage Static Pressure Distribution Ratios, Design

Configuration

COMPARISON OF TEST (TEST 8, LINE 10) AND PREDICTION

LOCATION PRESSURE RATIO TEST PREDICTED DELTA

(PRED.-TEST)

PT1 INLET TOTAL PRESSURE, PSIA= 212.6 3747.3 . . .

N-1 INLET. - TIP/PT1 0.977 0.989 +0.012

N-1 INLET. - HUB/PT1 0.990 0.989 -0.001

N-1 INLET. - INACTIVE ARC/PT1 0.996 1.000 +0.004

N-1 OUTLET. - TIP/PT1 0.807 0.792 -0.015

N-1 OUTLET. - HUB/PT1 0.768 0.729 -0.039

N-1 OUTLET. - INACTIVE ARC/PT1 0.738 0.767 (1) +0.029

N-2 INLET. - TIP/PT1 0.746 0.762 +0.016

N-2 INLET.- HUB/PT1 0.743 0.762 +0.019

N-2 INLET. - INACTIVE ARC/PT1 0.740 0.762 (1) +0.022

N-2 INLET. - CAVITY/PT1 0.743 0.762 (1) +0.019

N-2 OUTLET. - TIP/PT1 0.602 0.606 +0.004

N-2 OUTLET. - HUB/PTI 0.576 0.547 -0.029

N-2 OUTLET. - INACTIVE ARC/PT1 0.568 0.583 (1) +0.015

R-2 OUTLET./PT1 0.571 0.579 +0.008

N-1 :FIRST STAGE NOZZLE

N-2: SECOND STAGE NOZZLE

(1) MEAN GAS PATH PRESSURE

-57-

NASA CR-179548

RI/RD 86-214

z E

E-CC

CE

CE- -c E-~

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C-58

NASA CR-179548

Figure 33 RI/RD 86-214STATIC PRESSURE DISTRIBUTIONS

DESIGN CONFIGURATION, 0 DEGREE SECOND NOZZLETEST 8, LINE 10, 21086 RPM, 1.751 PRESSURE RATIO1.02 -O0HUB

FIRST STAGE NOZZLE INLET TIP

AINACTIVE ARCz 1.00 o CHAMBER

c/) 0.9913. 0

0.98 -

0.97 L

0.84 -. FIRST STAGE NOZZLE OUTLET

0.82

0.80 U.-.z•. 0.78 6

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0.74

0.72 L0.80

SECOND STAGE NOZZLE INLET0.78

z 0.76

rn 0.74 A

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0.70

0.62SECOND STAGE NOZZLE OUTLET

0.60

z 0.58a.cf 0.56a.

0.54

0.52 I ,, I . I0 45 90 135 180 225 270 315 360

CIRCUMFERENTIAL ANGLE, DEGREES

-59-

NASA CR-179548RI/RD 86-214

Symmetry was shown for the two arcs of admission by comparing the hub pressure at 90

degrees with the hub pressure at 270 degrees. The distribution proportions for the

first stage nozzle inlet in Figure 33 were typical of all tests.

The first stage nozzle outlet pressure ratios in Figure 33 showed the pressure

distribution for the tangential swirl out of the nozzle with the tip pressures (squares)

higher than the hub pressures (circles) by values similar to those predicted in Table

8. The active arc pressures were higher than the inactive arc pressures indicating

slight nozzle underexpansion for good performance rather than overexpansion with

potential for recompression and separation. Symmetry was shown in comparing the hub

and tip pressures at 90 and 270 degrees and the inactive pressures at 0 and 180

degrees. The distributions of the hub and tip pressures between 250 and 270 degrees

changed with the second stage nozzle orientation position.

The second stage nozzle outlet pressure ratios showed a uniform circumferential

distribution for the near axial, low velocity head flow out of the active arcs of the first

rotor. The hub and tip pressures were nearly equal at the 310 degree location. A

difference was shown at the 345 degree location which changed with second stage nozzle

orientation.

The second stage nozzle outlet pressure ratios showed the nozzle outlet swirl in the

active arcs with the tip pressures higher than the hub pressures. Symmetry was shown

around the annulus for the pressures available. Underexpansion was shown comparing

the average active arc pressures with the inactive arc pressures. The second stage

nozzle orientation angles had a greater effect on the upstream pressures than on the

second stage nozzle outlet pressures.

Circumferential variations for various second stage nozzle orientations for design

operating conditions were determined by comparing the pressure ratios in Figures 34

for the +40 degree second nozzle position and the ratios in Figure 35 for the -30

degree position with the ratios in Figure 33 for the design 0 degree second nozzle

position.

The +40 degree position in Figure 28 included the center of the first stage nozzle inlet

and the center of the second stage nozzle inlet in the same axial plane with no

circumferential offset. In the -30 degree position at the bottom of Figure 28,

significant circumferential offset was present (70 degrees in the direction of rotation)

-60-

NASA CR-179548RI/RD 86-214

Figure 34

STATIC PRESSURE DISTRIBUTIONSDESIGN CONFIGURATION, +40 DEGREE SECOND NOZZLE

TEST 8, LINE 9, 21096 RPM, 1.754 PRESSURE RATIO1.02 - 0OHUBr FIRST STAGE NOZZLE INLET A M TIP1.01 A INACTIVE ARC

* CHAMBER

0.99 -

0.9L

0.9-'

FIRST STAGE NOZZLE OUTLET

0.8Cz.E 0.78 -

p- 0.76

0.74

0.72

0.8007SECOND STAGE NOZZLE INLET

0.78

z 0.76

rn 0.741

0.72

0.70

0.620 SECOND STAGE NOZZLE OUTLET0.60f -

0 0.58

cn 0.56

0.54

0.52 . . . .

0 45 90 135 180 225 270 315 360


-61-

NASA CR-179548RI/RD 86-214

Figure 35STATIC PRESSURE DISTRIBUTIONS

DESIGN CONFIGURATION, -30 DEGREE SECOND NOZZLETEST 8, LINE 14, 21101 RPM, 1.777 PRESSURE RATIO

1.02 - 0 HUBFIRST STAGE NOZZLE INLET A U TIP

A INACTIVE ARC

z 1.00 * CHAMBER"NkcO 0.99 ! 0c-

0.98 L0.97

0.84 -

0.82

-

0.80

D- 0.78"FIRST STAGE NOZZLE OUTLET

•- 0.76

0.74A A

0.72 L

0.30 -SECOND STAGE NOZZLE INLET

0.78

z 0.76E_C-1cr 0.74 •

0.72

0.70


0.60

Z 0.58

co 0.56 -

0.54 0

0 45 90 135 180 225 270 315 360

CIRCUMFERENTIAL ANGI E, DEGREES

-62-

NASA CR-179548

RI/RD 86-214

with almost no overlap of the first stage nozzle inlet arc with the second stage nozzle

inlet arc.

The design second stage nozzle position at 40 degree circumferential offset was between

the two extremes tested. Since a flow molecule traverses circumferentially in the

direction of rotation in passing through the turbine, highest performance is achieved

when the second stage nozzle inlet is positioned to most efficiently capture the kinetic

energy leaving the first stage active arcs of admission. Highest test efficiency was shown

for the +10 degree second stage nozzle orientation with only a slight reduction for the

+40 degree orientation.

The first stage nozzle inlet and second stage nozzle outlet pressure ratios did not show

significant changes for the second stage nozzle orientations. However, the active arc

distributions at the first stage nozzle outlet and at the second stage nozzle inlet did show

the effects of the second stage nozzle orientation changes.

The first stage nozzle outlet tip pressure ratios increased from the 258 degree to the

282 degree angles due to the effective blockage with no offset for the + 40 degree second

nozzle position in Figure 34 compared with Figure 33 for the design 0 degree

position. The hub pressures at 310 degree and 345 degree angles were also higher for

the same reason. The second stage nozzle inlet pressure ratios at both the hub and tip

exhibited an increasing trend from 235 to 305 degrees for the +40 second stage nozzle

position in Figure 34. The higher velocity head and lower static pressures are shown

for the +40 second stage nozzle in Figure 34, lower than the inactive arc pressures.

Comparing the -30 degree, high offset, second stage nozzle orientation in Figure 35

with the design in Figure 33, less increase was shown for the tip pressures at the first

stage nozzle outlet from 258 degrees to 282 degrees for the - 30 degree second stage

nozzle. The lower value at 282 degrees indicated a lower resistance with the higher

offset. The second stage nozzle inlet pressure ratios were lower at the 305 degree and

340 degree angles indicating lower pressures required to pass the flow beyond the range

of kinetic energy from the first stage.

The circumferential variations with shaft speed at the design pressure ratio and second

nozzle orientation are shown in Figure 36 for the test at 10,052 RPM and in Figure

37 for the test at 25,003 RPM. The 25,003 RPM test compared closely with

-63-

NASA CR-179548

RI/RD 86-214

Figure 36STATIC PRESSURE DISTRIBUTIONS

DESIGN CONFIGURATION, 0 DEGREE SECOND NOZZLETEST 8, LINE 82, 10052 RPM, 1.759 PRESSURE RATIO1.02- U

FIRST STAGE NOZZLE INLET A HUB

1.01 -TIP

A INACTIVE ARCz 1.00 * CHAMBER

sAL

m 0.99 -

0.98 -

0.97 L

0.84 -. FIRST STAGE NOZZLE OUTLET

0.82

0.80 ,z U

•- 0.78 -N0

o 0.76

0.74 A

0.72


0.78

z 0.76 A

Cn 0.74

0.72

0.70


0.60

Z 0.58

Cn 0.56

0.54 -

0.52 ' .. I ..... I.. .. , ..0 45 90 135 180 225 270 315 360


-64-

NASA CR-179548

Figure 37 RI/RD 86-214

STATIC PRESSURE DISTRIBUTIONSDESIGN CONFIGURATION, 0 DEGREE SECOND NOZZLETEST 8, LINE 57, 25003 RPM, 1.741 PRESSURE RATIO1.02- U

FIRST STAGE NOZZLE INLET A HUB

1.01 A INACTIVE ARC

S1.00 - * CHAMBER

c/) 0.99 -

0.98 -

0.97 -

0.84 -FIRST STAGE NOZZLE OUTLET

0.82

0.80z

0.78 -

a 0.76

0.74

0.72


0.78

Z 0.76

W 0.74 A

0.72

0.70


0.60

Z 0.58

1-1 AV 0.56

0.54

0 45 90 135 180 225 270 315 360

CIRCUMFERENTIAL ANGLE. DEGREES

-65-

NASA CR-179548RI/RD 86-214

the design speed test in Figure 33. The overall veiocity change from design was less

than 20 percent. The velocity ratio change for the 10,052 RPM test was down 60

percent from design. Slightly higher first stage nozzle outlet pressures were shown

along with higher second stage nozzle inlet pressures indicating reduced active arc

reactions (lower rotor pressure drop), characteristic of lower velocity ratio operation.

A similar indication was shown with lower second stage nozzle outlet pressures

compared with the outlet flange pressure. Much greater pressure variations would have

existed for a full admission turbine for these operating condition changes along with

much greater turbine axial thrust changes.

Deslin Configuration Turbine Rotor Axial Thrust

The design configuration detail results for the test at the design equivalent operating

conditions are shown in Figure 38. The average active and inactive arc pressures are

shown upstream and downstream of each blade row along with the disk cavity pressures

and turbine downstream pressures. The active arc pressures were slightly higher than

the inactive arc pressures. The areas over which the pressures apply and the resulting

axial force component are shown. The summation of blade and disk forces for each rotor

are listed at the top along with the total turbine axial force of 119.7 pounds listed in

Table 6. The blade and disk forces were balanced down to the shaft seal diameters and

the total resulting turbine force was mainly from the unbalanced pressure on the

downstream side of the second stage rotor disk.

The test pressures were ratioed for the design inlet pressure of 3747 psia and the blade

and disk thrust loads were calculated and shown in Figure 39. The total turbine force

was 2,109 pounds which resulted from the high pressure level during operation. The

total force would have been much higher for a full admission reaction turbine and would

have been a significant function of operating shaft speed.

-66-


C-)

0U-z IVtoc

z . T , uiI

Cu 0

C6 l

cocoCV) 0 0 Z wZ-

10 - Cr

C-)

4) n i - ''d(

Pu, (A Z 0d.P

IL Cý LU L

00 cn4O wUJz-I-ic

L

> C40-. p Zl

0 2!(j 4

0W~

x CY -67-

NASA CR-179548

RI/RD 86-214

C.)

0

z W

0U

0 I0

0)~ 0 0 w -JD4) z C, r C-, 0XC

U) P

wI.-(D U) -Jl

LL4C O ow.l.1

ImI =Z >0 c',I-,- I 0.- N Co

ro-J

Cc~(I 0)C.)us4.LL

c68

NASA CR-179548RI/RD 86-214

HALF DESIGN ADMISSION CONFIGURATION

The half design admission configuration included a first stage admission of 13.8 percent

with 4 of 29 nozzles flowing, two per half 180 degrees apart, and a second stage nozzle

admission of 19.4 percent with 6 of 31 nozzles flowing, three per half 180 degrees

apart. The ratio of second stage nozzle admission to first stage nozzle admission was

1.406, compared with 1.310 for the design admission configuration. The half design

configuration was the ideal design for a turbine application with a lower specific speed

than the design admission configuration. Reasons for this are lower flowrate, lower

admission, higher pressure and higher second to first stage nozzle admission ratios for a

turbine with a near equal stage power split, and potentially lower speed for the lower

peak efficiency from reduced admission.

Half Configuration Efficiency Characteristic

The best fit efficiency curves for the 5 second stage nozzle orientation angles are shown

in Figure 40. A lower peak efficiency was shown for the smaller arc of admission with

a lower peak efficiency velocity ratio compared with the design admission configuration.

A greater variation in efficiency existed for the range of second stage nozzle orientation

angles. Relatively high efficiencies were shown in these figures for the tests with

pressure ratios of 2.0. A pressure ratio of 2.0 was used as the reference for analyses of

the half design admission configuration, rather than the design pressure ratio of 1.765,

because of the higher second to first stage nozzle admission ratio. Parameters for the

prediction, test points for the targeted design equivalent speed of 21,219 RPM, and the

equivalent pressure ratio of 2.0 are listed in Table 9 for the 5 second stage nozzle

angles. These efficiencies were plotted versus second stage nozzle angle in Figure 41,

along with the gas path program prediction. A much more significant effect from the

second stage nozzle orientation was shown for the half design admission compared to the

design admission configuration. An efficiency increase of 21 percent, 8.5 percentage

points, was shown from the -30 degree orientation to the +30 degree orientation. The

predicted efficiency was plotted at the design second stage nozzle 0 degree position, and

was 7.3 percent higher than the test value. The efficiency for the +30 degree

orientation was 3.7 percent lower than the prediction.

-69-


o~LL

~cn:

U).e

<LU-'-4

Ege- R Q<

UJM 0

-70-

NASA CR-179548RI/RD 86-214

Table 9

TARGET N (eq) AND PR (eq)i DESIGN ARC OF ADMISSION

2-N LINE N (eq) PR (eq) FLOW (eq) EFFICIENCY Um/CoANGLE, NUMBER (ETA)


0 PREDICTED 21219 2.000 0.0313 0.513 0.2662

-30 50 21563 1.985 0.0318 0.409 0.272

-10 54 21553 2.009 0.0321 0.437 0.270

0 61 21375 1.999 0.0325 0.478 0.268

+10 41 21259 1.996 0.0326 0.491 0.267

+30 46 21403 1.999 0.0329 0.494 0.269

-71-


1111117 -T~lri 1111 n1111 -T-Ti lIi jI *r1 11'* 11T

<U

wnLi !ý

q~CJ!

00-<i 0'

0 ) fir ic

w L(NJ LLi

00

C~i rn

EELPJ-72-

NASA CR-179548RI/RD 86-214

Half Configuration Flow Characteristic

The equivalent flow versus pressure ratio for the targeted design equivalent speed test

points are shown in Figure 42 for each of the 5 second stage nozzle orientation angles

along with the predicted values. The equivalent flow for the test points targeted at the

design equivalent speed of 21,219 RPM and pressure ratio of 2.0 from Table 9 are

shown in Figure 43 versus second stage nozzle orientation along with the prediction.

The equivalent flow at the 0 degree design orientation was 3.8 percent higher than

predicted. An increased equivalent flow of 3.1 percent was shown from -30 degrees to

+30 degrees. The efficiency also increased with increasing nozzle orientation angle,

with the optimum (beyond which efficiency and flow decrease) appearing higher than the

+30 degrees tested.

Half Configuration Static Pressure Distribution

The average static pressures measured at each interstage location were averaged and

compared with the predictions for the design equivalent test conditions in Figure 44

and in Table 10. Good agreement of the test data with the prediction was shown. The

largest deviation was for the first stage nozzle outlet hub pressure with a delta ratio of

0.076, or about 11 percent. The higher first stage pressure ratio was consistent with

the higher second-to-first-stage nozzle area ratio.

Half Confiouration Turbine Rotor Axial Thrust

The average active and inactive arc test pressures are shown in Figure 45. The areas

of active arc pressure were smaller for the half configuration. The areas and forces are

also shown in Figure 45. The total turbine force was less than the design configuration

because of the lower turbine outlet pressure. The pressures were scaled by the design

inlet pressure in Figure 46. The total turbine force was 1912 pounds for the high

design pressures.

-73 -


CNJj

Lc\J

D a

I-.j

cn~

LL- N

C)0

<UE LL1

Lo9

U0

CeD

(NJ

(-Da

< +x

F-74

NASA CR-179548RI/RD 86-214

L(\J

ED LLO.n

TqLi Fii

00 <

&-0 1 11

< F-ý-OoU)

0=J U)LJLL

<w (DL

-75

NASA CR-179548RI/RD 86-214

C)

ZZ)

E-Cx

0 (Y E

E- M E-

E-E-.f

U) Z -cc Z

M L

ýDa w00 ))

E-O

E-UE)

I I I14

-~~~ 0 -:Dor

-76-

NASA CR-179548RI'RD 86-214

Table 10. Interstage Static Pressure Distribution Ratios, Half Design

Configuration


LOCATION PRESSURE RATIO TEST PREDICTED DELTA(PRED. - TEST)

INLET TOTAL PRESSURE (PT1), PSIA 213.4 3747.3 - -

N-1 INLET - TIP/PT1 0.974 0.989 +0.015

N-1 INLET - HUB/PT1 0.987 0.989 -0.002

N-1 INLET - INACTIVE ARC/PT1 1.002 1.000 -0.002

N-1 OUTLET - TIP/PT1 0.774 0.735 -0.039

N-1 OUTLET- HUB/PT1 0.725 0.649 -0.076

N-1 OUTLET-INACTIVE ARC/PT1 0.692 0.697 (1) +0.005

N-2 INLET - TIP/PT1 0.679 0.700 +0.021

N-2 INLET - HUB/PT1 0.679 0.700 +0.021

N-2 INLET - INACTIVE ARC/PT1 0.690 0.700 (1) +0.010

N-2 INLET - CAVITY/PT1 0.690 0.700 (1) +0.010

N-2 OUTLET - TIP/PT1 0.532 0.515 -0.017

N-2 OUTLET - HUB/PT1 0.505 0.456 -0.049

N-2 OUTLET- INACTIVE ARC/PT1 0.497 0.489 (1) -0.008

N-2 OUTLET - CAVITY/PT1 0.490 0.489 (1) -0.001

R-2 OUTLET/PT1 0.500 0.488 -0.012




-77-


-j

0

zLU C4~

9- IM 9 0

CI U

0C jO uz o$, M0

(P 0s

o V ) c o cCt

0 >

o 0 Z

3m co)C.-c)0.0

0L >t. 0

0 00

C-4-

b. 1Z

CF Z V0 ~ .uFm-78-


w VIý

co 0!C)cI-

C4 T- L)

q~4.C*j wJY 0ww

L.j 0>

0 0

U,

pC >

w-l

00..

Q 4neJ. Oc2. LLJ

Ln(f -TwC4 V; a

-JJ );Co.. o.ch

I-I- FZ.Ln0

0(

cc 0) (nC.

4 N. o L

i79

NASA CR-179548RI/RD 86-214

QUARTER DESIGN ADMISSION CONFIGURATION

The quarter design admission configuration consisted of a first stage admission of 6.9

percent with 2 of 29 nozzles flowing, one per half, 180 degrees apart and a second stage

nozzle admission of 12.9 percent with 4 of 31 nozzles flowing, two per half, 180

degrees apart. The second stage to first stage nozzle admission ratio was 1.871 compared

with 1.406 for the half design admission configuration and 1.310 for the design

admission configuration. The quarter configuration design is applicable to a turbine

design with an even lower specific speed and a higher design pressure ratio for a near-

equal power split for the two stages for low flow losses.

Quarter Configuration Efficiency Characteristic

The best fit efficiency curves for the 5 second stage nozzle orientation angles are shown

in Figure 47. A lower peak efficiency was shown for the smaller arc of admission with

a lower peak efficiency velocity ratio compared with the larger admission configurations

tested. A more significant variation in efficiency was shown for the range of second stage

nozzle orientation angles. Relatively high efficiencies were shown in these Figures for

the test pressure ratio of 2.0 which was used as the reference for comparison of the

effects of second stage nozzle orientation angle for the quarter design admission

configuration. Parameters for these test points were tabulated in Table 11 for the

targeted design equivalent speed of 21,219 RPM and an equivalent pressure ratio of 2.0

for the 5 second stage nozzle orientation angles. The efficiencies were plotted versus

second nozzle angle in Figure 48 along with the gas path program prediction for the

quarter design admission configuration. The test values were significantly lower than

predicted by 26.5 percent, 11.2 percentage points. A 43.4 percent, 9.5 percentage

points, increase was shown in efficiency from the -30 degree to +30 degree second stage

nozzle orientation angles. A reduction in efficiency was shown for the +10 degree nozzle

angle compared with the 0 and +30 degree angles. This reduction was possibly due to a

shift of one of the second stage nozzle plugs during test, which was found upon

disassembly.

-80-

NASA CR-179548RI/RD 86-214

C)

jFTT~rnr-rF-T~rmr i i iTTTTT '0

U,

LU.,

wLCin

Lo<

N..J

-- -l

/L>*

K0

0-

wfI -I LUj j -

Nw -jUj c

EDt NJ 00U4

rnL~C 13 LJL

IN ~~~C\j x+ 4 0 l

InVO13 Vi3 ý,- 0U

NASA CR-179548RI/RD 86-214

Table 11

TARGET N (eq) AND PR (eq)i DESIGN ARC OF ADMISSION

2-N LINE N (eq) PR (eq) FLOW (eq) EFFICIENCY Um/CoANGLE, NUMBER (ETA)


O PREDICTED 21219 2.0001 0.01725 0.4218 0.2662

-30 44 20970 1.997 0.0164 0.219 0.255

-10 48 21282 1.979 0.0163 0.248 0.262

0 50 21144 1.992 0.0167 0.310 0.259

+10 56 21256 1.989 0.0163 0.277 0.260

+30 59 21143 1.978 0.0167 0.314 0.261

-82-

NASA CR-179548RI/RD 86-214

LI,

00

F-(d< I j

U) -- !

(NJF- (n

a-83

NASA CR-179548RI/RD 86-214

Ouarter Confiauration Flow Characteristic

The equivalent flow test points versus pressure ratio for the targeted design equivalent

speed of 21,219 RPM are shown in Figure 49 for each of the 5 second stage nozzle

orientation angles along with the predicted values. The equivalent flow for the test

points targeted at the design equivalent speed of 21,219 RPM and a pressure ratio of 2.0

from Table 11 are shown in Figure 50 versus second stage nozzle orientation along

with the design prediction. The equivalent flow at the 0 degree design orientation was

3.3 percent lower than predicted. An increasing equivalent flow characteristic was

shown from -30 to +30 degrees with a 1.8 percent increase in equivalent flow.

-84-

NASA CR-179548RI/RD 86-214

wN

x

(f) U

LLL.

L a-9

<-_L&J

F--Cr)

c\J UJ

p- .c 0-

a .4 0 > x

0 C\J

339/21 '3LVgMO-U i-N3-hAIflO3

-85-

NASA CR-179548RI/RD 86-214

Lfl

LLjJ

co

Cl.R

000-4

I....

g~i=

-'

N

it--

r'-

NASA CR-179548RI/RD 86-214

Quarter Configuration Static Pressure Distribution

The average static pressures measured at each inlerstage location were averaged and

compared with the predictions for the design equivalent test conditions in Figure 51

and in Table 12. Good agreement of the test data with the prediction was shown. The

largest deviation was for the first stage nozzle outlet hub pressure with a delta ratio of

0.076, or about 11 percent. The higher first stage pressure ratio was consistent with

the higher second-to-first stage nozzle area ratio.

Quarter Configuration Turbine Rotor Axial Thrust

The average active and inactive arc test pressures are shown in Figure 52. The areas

of active arc pressure were smaller for the half configuration. The areas and forces are

atso shown in Figure 52. The total turbine force was less for the design configuration

because of the lower turbine outlet pressure. The pressures were scaled by the design

inlet pressure in Figure 53. The total turbine force was 2,118 pounds for the high

design pressures.

-87-


z 1

E- Z E--

o2 a:0 E-E-

E- > >-

E- CL

LOO<

;,z, p

2. >z

0-I3

CC 0 0 C

G1OJJKVK 13IN1 I~ d d

-88-

NASA CR-179548RI/RD 86-214

TABLE 12. Interstage Static Pressure Distribution Ratios, Quarter

Design Configuration


LOCATION PRESSURE RATIO TEST PREDICTED DELTA

(PRED. - TEST)

INLET TOTAL PRESSURE (PT1), PSIA 215.6 3747.3 - - -

N-1 INLET - TIP/PT1 0.972 0.987 +0.015

N-1 INLET - HUB/PT' 0.982 0.987 +0.005

N-1 INLET - INACTIVE AriC/PT1 1.005 1.000 -0.005

N-1 OUTLET - TIP/PTI 0.667 0.656 -0.011

N-1 OUTLET- HUB/PT1 0.628 0.547 -0.081

N-1 OUTLET-INACTIVE ARC/PT1 0.612 0.608 (1) -0.004

N-2 INLET - TIP/PT1 0.607 0.623 +0.016

N-2 INLET- HUB/PT1 0.606 0.620 +0.014

N-2 INLET- INACTIVE ARC/PT1 0.611 0.622 (1) +0.011

N-2 INLET- CAVITY/PT1 0.618 0.622 (1) +0.004

N-2 OUTLET - TIP/PT1 0.509 0.521 -0.012

N-2 OUTLET - HUB/PT1 0.488 0.482 -0.006

N-2 OUTLET - INACTIVE ARC/PT1 0.498 0.504 (1) +0.006

N-2 OUTLET - CAVITY/PT1 0.491 0.504 (1) +0.013

R-2 OUTLET/PT1 0.498 0.492 -0.006




-89-


-jui

U.cooL

vi-

1*T)

.4

CYc

0 an o t

0 )o 1'(Oc i i 4

in'I t

* ulCM -J z

toU

Lu LU

0 4->cr,

-o Cl

L q '

CF p

4~CO,.2- 4o

Pcn U4 WLAJ

0~U~LLiCD4 -.

0wWLu -l> ITu

Ln t-90-

NASA CR-179548RI/RD 86-214

0

m 1-6t.'

U z

U)Cw

~1 13V ca

0

C 0

0 0 .0 )V- G

U)

0)

z

c U 'U

0. CC>

m~P -0

2) LLC- * C

LL C4 0 w

04 (n0

(1)4

OWWCV

40..4u

-91 >.

NASA CR-179548RI/RD 86-214

CONCLUSIONS

A review of the data was made and significant results relating to the performance of the

Two Stage Partial Admission Turbine test during this program were determined and

listed below.

1. The design two stage partial admission turbine test results exceeded the predicted

efficiency by 8 percent and matched the predicted flow within 3 percent.

2. The peak efficiency decreased as the arcs of admission decreased.

3. The overall velocity ratio at peak efficiency decreased as the arcs of admission

decreased.

4. Both efficiency and flow characteristics were sensitive to the second stage nozzle

orientation angle.

5. The second stage nozzle orientation angle sensitivity was greater at lower

admissions.

6. The highest efficiency and flow appeared between second nozzle orientation angles

of +10 degrees and +40 degrees for the design admission configuration, and

beyond +30 degrees for the half and quarter design admission configurations.

7. Inactive arc pressures across rotor blades (upstream to downstream) for all

configurations and operating conditions were nearly equal with partial admission

configurations.

8. First stage nozzle inlet and second rotor outlet pressures were not significantly

affected by second stage nozzle angular position variation.

9. First stage nozzle outlet and second stage nozzle inlet pressures increased due to

effective second stage nozzle inactive arc blockage at the ends of the active arcs

for the extreme second stage nozzle angular positions.

-92-

NASA CR-179548RI/RD 86-214

10. Only slight changes in interstage pressure distributions were shown for large

speed changes with partial admission compared to a full admission reaction

turbine.

11. Turbine axial thrust was not significantly affected by second stage nozzle angular

position.

12. Partial admission staging provided the lowest turbine axial thrust across the

blades and disks of any staging option.

-93-

NASA CR-179548RI/RD 86-214

REFERENCES

1. Anon: Orbit Transfer Vehicle Advanced Expander Cycle Engine Point Design Study

(Volume I1: Study Results). Report No. RI/RD80-218-2, Contract NAS8-33568,

Rocketdyne, December 1980.

2. Nusbaum, William J., and Wong, Robert Y., Effect of Stage Spacing on Performance of

3.75-Inch-Mean-Diameter Two-Stage Turbine Having Partial Admission in the

First Stage. NASA TN D-2335, 1964.

3. Holeski, D. E., and Stewart, W. L., Study of NASA and NACA Single-Stage Axial Flow

Turbine Performance as Related to Reynolds Number and Geometry. J. Eng. Power,

Vol. 86, No. 3, July 1964, pp. 296-298.

4. NASA TN D-4700, Cold-Air Investigation of Effects of Partial Admission on

Performance of 3.75-Inch Mean-Diameter Single-Stage Axial-Flow Turbine. Hugh

A. Klassen, NASA/LeRC, Cleveland, OH, Aug. 1968.

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NASA CR-1 79548RI/RD 86-214

APPENDIX A Standard Air Equivalent Conditions

Standard Air Conditions

Pressure a 14.696 PSIA

Temperature a 518.7 °R

Gas Constant - 53.345 FT LBf/LBm 0RRatio of Specific Heats . 1.4

Critical Velocity = 1019.5 FT/SEC

Standard Air Constants

2ygRTIZ 1 26CR = /(1019.5) , critical velocity ratio based on inlet

total temp., test-to-standard air, squared.

6 - P1 /14.696 , inlet pressure ratio, test-to-standard air

0.739594 + 1Y 2 , gamma function ratio, test-to-standard air

Where:

S - Ratio of Specific Heats - Test Gas

R - Gas Constant - Test Gas

T1 - Inlet Total Temp. - Test Gas

Z1 - Inlet Compressibility - Test Gas

P1 - Inlet Total Pressure - Test Gas

g - 32.174 FT/SEC2

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NASA CR-1 79548RI/RD 86-214

Standard Air Equivalent Parameters

W EQ a

N

N EQ a -

PR EQ - I

1 - 0.1666667 ( ) P )(PR) /

Where:

W - Test Flow Rate - Pounds/Second

N - Test Speed - RPM

PR - Test Pressure Ratio

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NASA CR-179548RI/RD 86-214

APPENDIX B Raw Data, Tests 8-10 (Pressure Distribution)

This appendix consists of three test data sets for the following three configurations:

Test Number Configuration

8 Design

9 Half Design

10 Quarter Design

The data was not reduced because two thermocouples measuring the turbine inlet

temperature and the flow nozzle inlet temperature were faulty during testing. The

channel numbers associated with these measurements were 3 and 4.

The data set represents the pressure distribution between stages, both radially (hub and

tip) and circumferentially. Channel numbers correspond to pressure tap locations

shown in Figure 2 2. Note that the pressure magnitudes are gage pressures.

-97-


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C.~ .- 130-

NASA CR-179548RI/RD 86-214

APPENDIX C Data Reduction Technique

The following describes the equations and methods used to

determine the overall turbine performance. All the pressures and

temperatures are absolute values.

Many of the properties equation were obtained by curve fitting

real property data. The real property data were obtained from

running a nitrogen property routine, ONTHPR, which was written,

in Rocketdyne's main-frame UNIVAC computer, by J. K. Jakobsen in

1972. The program is located in the Engineering Subroutine

Library.

The curvefit equations used will be included in this section,

along with their coefficients. The curvefits are only valid for

the range of test operation. A discussion of how the curvefits

coefficients were calculated will be discussed at the end of this

section.

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NASA CR-179548RI/RD 86-214

The flowrates were obtained by using the ideal isentropic

critical flow equation across a nozzle. The specific heat ratio

(gam) used in the equation has obtained by a curvefit equation

that characterized the real property gamma as a function of

temperature and pressure.

1 -(gam+l)

2 2 (gain-i)

2mdot :=g A M PC ga I1 + gm M][g -R TO] 12

where : P0 = Total Pressure, psia

TO = Total Temperature, deg R

M = Mach number

A = nozzle throat area, inA2

R = gas constant, ft-lbf/lbm-deg R

g = 32.174 ibm-ft/lbf-secA2

mdot= flowrate, lbm/sec

gam = specific heat ratio

The total pressure and temperature were obtained by iterating

between the Mach number at the nozzle inlet, due to the mass

flowrate. The mass flowrate was calculated, by the above

equation, where M = 1. The total pressure was initially assumed

to equal the static pressure measured. Using this initial

assumption a Mach number at the nozzle inlet was calculated. A

new total pressure was then calculated. This process was

repeated about 3 times before no further change occurred.

The nozzle inlet Mach number was calculated by dividing the mass

flow by the inlet area and sonic velocity. The total pressure

and temperature were calculated by the following equations

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NASA CR-179548RI/RD 86-214

gamn

r a 1 2e gain-i gar 1 2 ]P0 := P[1+ M TO := T 1 2 Miniei

Where T and P are the nozzle inlet static temperature and pressure.

The equation for gamma is : where kO = 1 .. 9

i := 1 .. 3 j := 1 .. 3 Cko

01.38718101309 10

11.33421229506 10

3i-I -3.36768173119 10

P -5gam >=C 7.24069630487 10

S 3 ( -l + j-I -i

j T -1.02536990231 10

6.55360232959 10-7

2.40239920704 10-4

-2.83856529863 10-2

8.13424970141 10

The specific volume of the gas was calculated by the following

curvefit equation :

i := 1 .. 3 ;j := 1 .. 3 where ko0 1 .. 9 Dko

-2-9.73934911339 10

-3-5.387717631225 10

i-i -2T -4.36782615969 10

V (i-l)+j j-i 2.9426818990a 10

i j P -13.82360778543 10

-22.27130816879 10

-7-2.25608326345 10

-79.8782949984 10

-5-3.06153388554 10

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NASA CR-179548RI/RD 86-214

The velocity, V, at the nozzle inlet was calculated by mass

continuity equation :

vV = 144@mdot V = inlet velocity, ft/sec

areainl area = inlet area, inA2

v = specific volume, ft 3/lbThe nozzle inlet Mach number :

VMinlet Sg gain R T

The turbine pressure ratio across the turbine was assumed to be

total-to-total since the two areas of measurement were relatively

large, that is, the Mach number was less than 0.1.

Pin

P :=-r P

out

The turbine efficiency was obtained by dividing the actual

enthalpy, across the turbine, by the isentropic enthalpy.

h -hin-act out-act

turb h -hin-act out isen

Two curvefit equations were used for enthalpy. One as a function

of temperature T and pressure P. The other as a function of

pressure P and entropy s. A curvefit equation was used for

entropy, as a function of pressure and temperature. The pressure

and temperature measured at the inlet were use to obtain the

turbine inlet enthalpy and entropy. The turbine outlet pressure

and temperature measured were used to obtain the actual outlet

enthalpy. The turbine outlet pressure and the inlet calculated

entropy were used to obtain the turbine outlet isentropic

enthalpy. The equations used are :

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NASA CR-179548RI/RD 86-214

i := 1 .. 4 j := 1 .. 4

i-i

S :=ZE 4 ,btu/(lm*deg R)4 (-l)+ji-i

i j p

where : for kla 1 .. 4

E a E Ekl kl+4

-1 -3

6.754475257t 10 2.14457854498 101 -3

6.4345893938 10 -5.35320318401 103 1

-3.23568313559 10 -2.17601576899 104 3

-1.66276892738 10 1.82707334974 10

E i E -

kl+8 kl+12S-6 -9

-2.40290241705 10 1.0863782209 10-4 -7

-1.35426963699 10 1.61878278367 10-2 -5

6.4683199835? 10 -5.38690472998 100 -3

-4.67810491676 10 3.6511832911 10

i := 1 .. 4 j := 1 .. 4

1-1

Th :=? .F; btu/lbmPT F 4 (i-l)+j J-i

i j P

where : for k2a 1 .. 4

F a Fk2 k2+41 -1

-5.22655244231 10 4.11793038505 103 0

9.464278572589 10 -7.04911193625 105 3

-3.40829266077 10 -4.56884504042 107 5

-1.7935734637 10 4.07625294607 10

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NASA CR-179548RI/RD 86-214

F Fk2+8 k2+12

-4 -8

-1.69214857206 10 4.75648465964 10-2 -5

-4.17603347662 10 5.24523782643 101 -2

1.6882782708 10 -1.48071153388 103 -1

-1.1773032577 10 9.55472887408 10

i 1 .. 3 j := 1 .. 6i-i

Ps 6 (i-l) +j j-1i j P

where : for k3 = 1 .. 6

G r G * Gk3 k3+6 k3+12

2 3 29.03482161609 10 -1.70554732153 10 8.63431991984 10

4 5 47.7392940758 10 -1.08062748155 10 1.14082068107 10

7 6 61.63942252554 10 -6.44250655923 10 2.43746386678 10

9 9 8-4.83895335393 10 2.69374376999 10 -3.136026318 10

11 10 103.27377498877 10 -5.40664991925 10 -6.60879065646- 10

12 13 12-3.052261796 10 -1.01307802656 10 6.96976678636 10

The mean velocity to isentropic velocity ratio is :

D N

U 12 60 where :

C U = ft/seco I2g J[hnrct- houie

.F in-actu tis ejj Co = ft/sec

N = rmp

D = dia, inch

J = 778.17 ft*lbf/btu

This parameter is the loading function of the turbine.

It is useful to convert the turbine performance to equivalent

standard ambient state conditions. The following equations were

used to obtain equivalent speed and flow.

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NASA CR-179548RI/RD 86-214

mdot F e

m := ibm/sec : (Equivalent flow)eq

NN :=- rpm : (Equivalent Speed)

eq vrwhere

2 gam RT zTote

cr 2(gain + l). (1019.5)

PTot

$ :and14.696

0.7396

gam.6 :--

gam.

gam-i

The following discussion will be on curvefit coefficient

determination. The curvefit equations were derived by intuitive

interpretation of real property curve trends. If a dependent

variable, such as specific volume, was a function of two

independent variables, temperature and pressure, then the

equation would be a matrix of these two variables. The matrix

would be a cross product of two polynomial of the two independent

variables. The polynomial power would be positive if the

dependent variable increased with the independent variable and

negitive for the reverse.

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NASA CR-179548RI/RD 86-214

To illustrate; specific volume increases with temperature and

decreases with pressure, therefore, the equation would look like

the following:

2] b2 b3vol := [al + a2 T + a3 T [bl +-- +

P

or multiplying across and re-substituting new dummy coefficients

the equation would be as the folowing.

21 1 T T T

vol := c + c-- + c'-- + c, T + c - + c'- .. c'-1 2 p 3 2 4 5P 6 2 9 2

P P P

Since each function, that make up the above equation, only has

one coefficient in from of it, the coefficients can be solved by

the least square fit process. The process is performed by

substituting the above equation into a better format.

Let1 1

f :=Vol ;f :=1 ; f :- f :0 1 2 P 2 2

P

2T T T

f :=T ; f :=- f :-- ; ... f :=-3 4 P 5 2 9 2

P P

Now the equation is in a linear type equation, in addition n was

used instead of subscript 9, so as not to limit the process

capability.

f :=c f + c f + c f .. c f + c- f0 1 1 2 2 3 3 n-1 n-i n n

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NASA CR-179548RI/RD 86-214

The coefficients can than be solved by the matrix equation

I(f f) Z(f f) Z(f f) ... I(f f) C £(f f)1 1 1 2 1 3 1 n 1 0 1

,(f f) X(f f) L(f f) ... -(f f ) c I(f f)2 1 2 2 2 3 2 n 2 0 2

(ff(f f) f(f f Lf f) ... (f f) C (f f)n 1 n 2 n 3 n n n 0 n

Where the f functions are summation of all the data set given.

The values of the coefficients are obtained by solving the above

system of equations.

The coefficients obtained will be subject to run-off errors from

the accuracy of the computer solving the system of equation and

the software using the coefficients. For this reason, the

property curvefit equation may experience 1 to 2 percent errors.

Actually the coefficients are not off by much, but, because

performance calculations are usually deferrence of two points the

delta of big numbers are prone to higher inaccuracy.

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NASA CR-179548RI/RD 86-214

APPENDIX D Raw and Reduced Data, Tests 11-13

This appendix consists of three test data sets for the following three configurations:

s Design Configuration

1 1 Design

12 Half Design

13 Quarter Design

The measurements consisted of three temperatures, five pressures, nozzle position and

speed. These tests were essentially performed in order to recover the overall turbine

efficiency not attainable from Tests 8 through 10 due to the two faulty thermocouples.

Each data set consists of an input and an output section. The input section is the test

measurement values ana the output section is the calculated performance.

-140-

NASA CR-179548RI/RD 86-214

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-16

S Report Documentation Page

1. Report No. 2. Government Accession No. 3. Recipient's Catalog No.CR-179548

4. Title and Subtitle 5. Report Date

Design and Experimental Performance of a Two Stage Partial 14 December 1992Admission Turbine 6. Performing Organization Code

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

R.F.Sutton, J.L.Boynton, and R.A.Akian RI/RD92-214

10. Work Unit No.

9. Performing Organization Name and Address TASK B.1 / B.4ROCKWELL INTERNATIONAL 11. Contract or Grant No.Rocketdyne Division NAS 3-237736633 Canoga AvenueCanoga Park, California 91303 13. Type of Report and Period Covered

12. Sponsoring Agency Name and Address Final Report; Dec 84 to Jun 86National Aeronautic Space Administration-Lewis Research CenterSpace Vehicle Propulsion Branch 14. Sponsoring Agency Code21000 Brookpark RoadCleveland, Ohio 44135

15. Supplementary Notes

Project Manager, G. Paul Richter, NASA Lewis Research Center, Cleveland, Ohio

16. AbstractA three-inch mean diameter, two-stage turbine with partial admission in each stage was experimentallyinvestigated over a range of admissions and angular orientations of admission arcs. Three configurationswere tested in which first stage admission varied from 37.4 percent (10 of 29 passages open, 5 per side) to 6.9percent (2 open, 1 per side). Corresponding second stage admissions were 45.2 percent (14 of 31 passagesopen, 7 per side) and 12.9 percent (4 open, 2 per side). Angular positions of the second stage admission arcswith respect to the first stage varied over a range of 70 degrees.

Design and off-design efficiency and flow characteristics for the three configurations are presented.The results indicated that peak efficiency and the corresponding isentropic velocity ratio decreased as the arcsof admission were decreased. Both efficiency and flow characteristics were sensitive to the second stagenozzle orientation angles.

17. Key Words (Suggested by Author(s)) 18. Distribution StatementTurbine Performance Unclassified - UnlimitedTwo-Stage Partial Admission TurbineImpulse Turbine

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

Unclassified Unclassified 161

NASA FORM 1626 Oct 86

Documents

NASA CR-1 79548 RI/RD 8 -214 94-08574 - DTICAD-A277 095 NASA CR-1 79548 RI/RD 8 -214 94-08574 FINAL REPORT Design and Experimental Performance of a Two Stage Partial Admission Turbine