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
94 8 16 090
NASA CR-1 79548
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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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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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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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NASA CR-179548
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Figure 4
VELOCITY VECTOR DIAGRAMMK 49-F TURBINE
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(0. 327)• 876 (0.164)
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First SecondStage SecondNozzle NozzleStage Stage
Rotor Rotor
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Figure 5 MK49-F Turbine Cross Section
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NASA CR-i179548RI/RD 86-214
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NASA CR-179548RI/RD 86-214
Figure 7 First and Second Stage Nozzle Angular Orientation
MK 49-F TURBINE
DESIGN SECOND NOZZLE INLET CIRCUMFERENTIAL DISPLACEMENT- 40 DEGREES IN DIRECTION OF ROTATION
40 Degree Angle
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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-i179548RI/RD BE-214
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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
Rctating Assembly Model
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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
*~ ~ . .. .. , . . . . . .
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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
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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
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NASA CR-i179548RI/RD 86-214
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-32
NASA CR-179548RI/RD 86-214
0
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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~
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NASA CR-i179548RI/RD 86-214
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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-
NASA CR-i179548RI/RD 86-214
~~dc
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'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
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NA:SA CR-179548RI/RD 86-214
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-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-
NASA CR-i179548RI/RD 86-214
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-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
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NASA CR-i /9548RI/RD 86-214
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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
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-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
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-54
NASA CR-i179548RI/RD 86-214
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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
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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
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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
CIRCUMFERENTIAL ANGLE, DEGREES
-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
0.62SECOND STAGE NOZZLE OUTLET
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
0.80SECOND STAGE NOZZLE INLET
0.78
z 0.76 A
Cn 0.74
0.72
0.70
0.62SECOND STAGE NOZZLE OUTLET
0.60
Z 0.58
Cn 0.56
0.54 -
0.52 ' .. I ..... I.. .. , ..0 45 90 135 180 225 270 315 360
CIRCUMFERENTIAL ANGLE, DEGREES
-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
0.80SECOND STAGE NOZZLE INLET
0.78
Z 0.76
W 0.74 A
0.72
0.70
0.62SECOND STAGE NOZZLE OUTLET
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-
NASA CR-i179548RI/RD 86-214
C-)
0U-z IVtoc
z . T , uiI
Cu 0
C6 l
cocoCV) 0 0 Z wZ-
10 - Cr
C-)
4) n i - ''d(
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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-
NASA CR-i179548RI/RD 86-214
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)
DEGREES TEST 12 RPM LB/SEC
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-
NASA CR-i179548RI/RD 86-214
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 -
NASA CR-i179548RI/RD 86-214
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
COMPARISON OF TEST (TEST 9, LINE 10) AND PREDICTION
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
N-1 :FIRST STAGE NOZZLE
N-2: SECOND STAGE NOZZLE
(1) MEAN GAS PATH PRESSURE
-77-
NASA CR-i179548RI/RD 86-214
-j
0
zLU C4~
9- IM 9 0
CI U
0C jO uz o$, M0
(P 0s
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NASA CR-i179548RI/RD 86-214
w VIý
co 0!C)cI-
C4 T- L)
q~4.C*j wJY 0ww
L.j 0>
0 0
U,
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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)
DEGREES TEST 13 RPM LB/SEC
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
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339/21 '3LVgMO-U i-N3-hAIflO3
-85-
NASA CR-179548RI/RD 86-214
Lfl
LLjJ
co
Cl.R
000-4
I....
g~i=
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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-
NASA CR-i179548RI/RD 86-214
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
COMPARISON OF TEST (TEST 10, LINE 14) AND PREDICTION
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
N-1 :FIRST STAGE NOZZLE
N-2: SECOND STAGE NOZZLE
(1) MEAN GAS PATH PRESSURE
-89-
NASA CR-i179548RI/RD 86-214
-jui
U.cooL
vi-
1*T)
.4
CYc
0 an o t
0 )o 1'(Oc i i 4
in'I t
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NASA CR-179548RI/RD 86-214
0
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U z
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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.
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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-
NASA CR-i179548RI/RD 86-214
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NASA CR-179548RI/RD 86-214
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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