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  • 1
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    Analytical study of flow phenomena in SSME turnaround duct geometries (1985)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The SSME fuel turbopump hot gas manifold was identified as a source of loss and flow distortion which significantly affects the performance and durability of both the drive turbine and the LOX injector area of the main combustion chamber. Two current SSME geometries were studied, the full power level (FPL) and the first manned orbital flight (FMOF) configuration. The effects of turnaround duct geometry on flow losses and distortions, by varying wall curvature and flow area variation in the 180 deg turnaround region were examined. The effects of the duct inlet flow phenomena such as the radial distortion of the inlet flow and inlet swirl level on turnaround duct performance were also investigated. It is shown that of the two current geometries, the FMOF configuration had lower pressure losses and generated less flow distortion, but had a small flow separation bubble at the 180 deg turnaround exit. It is found that by optimizing wall curvature and flow diffusion in the turnaround, improved duct performance can be achieved.
    Keywords: FLUID MECHANICS AND HEAT TRANSFER
    Type: NASA. Marshall Space Flight Center Advan. High Pressure O2(H2 Technol.; p 579-598
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 2
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    Effect of endwall cooling on secondary flows in turbine stator vanes (1977)
    In:  CASI
    Details
    Publication Date: 2012-05-19
    Description: An experimental investigation was performed to determine the effect of endwall cooling on the secondary flow behavior and the aerodynamic performance of a coreturbine stator vane. The investigation was conducted in a cold-air, full-annular cascade, where three-dimensional effects could be obtained. Two endwall cooling configurations were tested. In the first configuration, the cooling holes were oriented so that the coolant was injected in line with the inviscid streamline direction. In the second configuration, the coolant was injected at an angle of 15 deg to the inviscid streamline direction and oriented toward the vane pressure surface. In both cases the stator vanes were solid and uncooled so that the effect of endwall cooling could be obtained directly. Total-pressure surveys were taken downstream of the stator vanes over a range of cooling flows at the design, mean-radius, critical velocity ratio of 0.778. Changes in the total-pressure contours downstream of the vanes were used to obtain the effect of endwall cooling on the secondary flows in the stator. Comparisons were made between the two cooled-endwall configurations and with the results obtained previously for solid endwalls.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: AGARD Secondary Flows in Turbomachines; 29 p
    Format: text
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    NASA TECHNICAL REPORTS
  • 3
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    Solar dynamic space power system heat rejection (1986)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A radiator system concept is described that meets the heat rejection requirements of the NASA Space Station solar dynamic power modules. The heat pipe radiator is a high-reliability, high-performance approach that is capable of erection in space and is maintainable on orbit. Results are present of trade studies that compare the radiator system area and weight estimates for candidate advanced high performance heat pipes. The results indicate the advantages of the dual-slot heat pipe radiator for high temperature applications as well as its weight-reduction potential over the range of temperatures to be encountered in the solar dynamic heat rejection systems.
    Keywords: SPACECRAFT PROPULSION AND POWER
    Format: text
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  • 4
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    Effects of interstage diffuser flow distortion on the performance of a 15.41-centimeter tip diameter axial power turbine (1983)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: The performance of a variable-area stator, axial flow power turbine was determined in a cold-air component research rig for two inlet duct configurations. The two ducts were an interstage diffuser duct and an accelerated-flow inlet duct which produced stator inlet boundary layer flow blockages of 11 percent and 3 percent, respectively. Turbine blade total efficiency at design point was measured to be 5.3 percent greater with the accelerated-flow inlet duct installed due to the reduction in inlet blockage. Blade component measurements show that of this performance improvement, 35 percent occurred in the stator and 65 percent occurred in the rotor. Analysis of inlet duct internal flow using an Axisymmetric Diffuser Duct Code (ADD Code) were in substantial agreement with the test data.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: AIAA PAPER 83-1179
    Format: text
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    NASA TECHNICAL REPORTS
  • 5
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    Improved automobile gas turbine engine (1976)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: Upgraded engine delivers 100 hp in 3500 lb vehicle. Improved fuel economy is due to combined effects of reduced weight, reduced power-to-weight ratio, increased turbine inlet pressure, and improved component efficiencies at part power.
    Keywords: MACHINERY
    Type: LEW-12521 , NASA Tech Briefs (ISSN 0145-319X); 1; 1; P. 116
    Format: text
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  • 6
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    Experimental performance and analysis of 15.04-centimeter-tip-diameter, radial-inflow turbine with work factor of 1.126 and thick blading (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The aerodynamic design, the performance, and an internal loss breakdown were examined for a 15.04 cm tip diameter, radial-inflow turbine. The design application was to drive a two stage, 10 to 1 pressure ratio compressor with a mass flow of 0.952 kg/sec and a rotative speed of 70,000 rmp. The turbine inlet temperature was 1478 K, and the turbine was designed with blades thick enough for internal cooling passages. The rotor tip diameter was limited to 86 percent of optimum in order to obtain a reduced tip speed design. The turbine was fabricated with solid, uncooled blading and tested in air at nominal inlet pressure and temperature of 1.379 x 10000 N/sq m and 322.2 K, respectively. Results indicated the turbine total efficiency to be 5.3 points less than design. Analysis of these results has indicated the deficit in performance to be due to stator secondary flow losses, vaneless space surface friction losses, and trailing edge wake mixing losses.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: NASA-TP-1730 , E-391 , AVRADCOM-TR-80-09
    Format: application/pdf
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  • 7
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    Cold-air performance of a 15.41-cm-tip-diameter axial-flow power turbine with variable-area stator designed for a 75-kW automotive gas turbine engine (1982)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: An experimental evaluation of the aerodynamic performance of the axial flow, variable area stator power turbine stage for the Department of Energy upgraded automotive gas turbine engine was conducted in cold air. The interstage transition duct, the variable area stator, the rotor, and the exit diffuser were included in the evaluation of the turbine stage. The measured total blading efficiency was 0.096 less than the design value of 0.85. Large radial gradients in flow conditions were found at the exit of the interstage duct that adversely affected power turbine performance. Although power turbine efficiency was less than design, the turbine operating line corresponding to the steady state road load power curve was within 0.02 of the maximum available stage efficiency at any given speed.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: NASA-TM-82644 , E-899 , NAS 1.15:82644 , DOE/NASA/51040-30
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  • 8
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    The aerodynamic design of a compressor-drive turbine for use in a 75 kw automotive engine (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The design of a single stage axial-flow turbine with a tip diameter of 11.15 cm is presented. The design specifications are given, and the aerodynamic design procedure is described. The aerodynamic information includes the results of flow path, velocity diagram, and blade profile studies. Predicted off-design performance characteristics are also presented.
    Keywords: MECHANICAL ENGINEERING
    Type: NASA-TM-X-71717
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  • 9
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    Cold-air annular-cascade investigation of aerodynamic performance of core-engine-cooled turbine vanes. 2: Pressure surface trailing edge ejection and split trailing edge ejection (1976)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The aerodynamic performance of two trailing edge ejection cooling configurations of a core-engine stator vane were experimentally determined in an ambient inlet-air full-annular cascade where three-dimensional effects could be obtained. The tests were conducted at the design mean-radius ideal aftermixed critical velocity ratio of 0.778. Overall vane aftermixed thermodynamic and primary efficiencies were obtained over a range of coolant flows to about 10 percent of the primary flow at a primary to coolant total temperature ratio of 1.0. The radial variation in efficiency and the circumferential and radial variations in vane-exit total pressure were determined. Comparisons are made with the solid (uncooled) vane.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: NASA-TM-X-3369 , E-8566
    Format: application/pdf
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  • 10
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    Cold-air performance of the compressor-drive turbine of the Department of Energy baseline automobile gas-turbine engine (1978)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The aerodynamic performance of the compressor-drive turbine of the DOE baseline gas-turbine engine was determined over a range of pressure ratios and speeds. In addition, static pressures were measured in the diffusing transition duct located immediately downstream of the turbine. Results are presented in terms of mass flow, torque, specific work, and efficiency for the turbine and in terms of pressure recovery and effectiveness for the transition duct.
    Keywords: AERODYNAMICS
    Type: NASA-TM-78894 , E-9480 , DOE/NASA/1011-78/25
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