Predicted turbine stage performance using quasi-three-dimensional and boundary-layer analyses
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Boyle, R. J.↗
Engineering topics
Haas, J. E.
Publications and source records attributed to Haas, J. E..
Fuel and oxidizer turbine loss analysis
The turbine losses for the fuel and oxidizer turbines at the FPL condition were assessed by a quasi-3D loss analysis method. This loss analysis method uses two flow codes - MERIDL and TSONIC - to calculate the flow velocities along the blade surfaces and endwalls. The velocities are then used as input to the boundary layer code - BLAYER - to calculate the friction losses due to incidence, secondary flow, and tip clearance. The loss analysis for the fuel turbine indicated an overall two-stage efficiency of about 90%. The largest loss was due to rotor tip clearance. The loss analysis for the oxidizer turbine is nearly completed. Results for the first stage of the two-stage design indicates an efficiency of about 80%, with high losses due to rotor incidence and blade and endwall friction.
Haas, J. E.↗
Analytical and experimental investigation of stator endwall countouring in a small axial-flow turbine
An experimental and analytical investigation was conducted to determine the effect of stator endwall contouring on turbine stage performance. In this investigation three stator configurations were evaluated using a common rotor. The three stator configurations were a cylindrical endwall design and two contoured endwall designs, one having a S-shaped outer wall profile and the other having a conical-shaped outer wall profile. Experimental data were obtained over a range of equivalent speeds, total pressure ratios, and rotor tip clearances for each stator-rotor combination. Detailed analytical loss assessments were conducted to aid in the determination of the contouring effect on turbine performance.
Haas, J. E.↗
Comparison between measured turbine stage performance and the predicted performance using quasi-3D flow and boundary layer analyses
A method for calculating turbine stage performance is described. The usefulness of the method is demonstrated by comparing measured and predicted efficiencies for nine different stages. Comparisons are made over a range of turbine pressure ratios and rotor speeds. A quasi-3D flow analysis is used to account for complex passage geometries. Boundary layer analyses are done to account for losses due to friction. Empirical loss models are used to account for incidence, secondary flow, disc windage, and clearance losses. Previously announced in STAR as N84-22564
Boyle, R. J.↗
Comparison between measured turbine stage performance and the predicted performance using quasi-3D flow and boundary layer analyses
A method for calculating turbine stage performance is described. The usefulness of the method is demonstrated by comparing measured and predicted efficiencies for nine different stages. Comparisons are made over a range of turbine pressure ratios and rotor speeds. A quasi-3D flow analysis is used to account for complex passage geometries. Boundary layer analyses are done to account for losses due to friction. Empirical loss models are used to account for incidence, secondary flow, disc windage, and clearance losses.
Boyle, R. J.↗
Cold-air performance of compressor-drive turbine of department of energy upgraded automobile gas turbine engine. 3: Performance of redesigned turbine
The aerodynamic performance of a redesigned compressor drive turbine of the gas turbine engine is determined in air at nominal inlet conditions of 325 K and 0.8 bar absolute. The turbine is designed with a lower flow factor, higher rotor reaction and a redesigned inlet volute compared to the first turbine. Comparisons between this turbine and the originally designed turbine show about 2.3 percentage points improvement in efficiency at the same rotor tip clearance. Two versions of the same rotor are tested: (1) an as cast rotor, and (2) the same rotor with reduced surface roughness. The effect of reducing surface roughness is about one half percentage point improvement in efficiency. Tests made to determine the effect of Reynolds number on the turbine performance show no effect for the range from 100,000 to 500,000.
Roelke, R. J.↗
Application of viscous and inviscid computation methods for rocket turbopump systems
This paper examines current research efforts underway at Lewis Research Center which have focused on improving the structural durability of the SSME turbodrive system through more fully examining the aerothermodynamics of the flow system. This objective was achieved by developing three-dimensional viscous and inviscid computer codes to predict the pressure and temperature distributions in the engine flow passages, particularly within the SSME turbines and turnaround ducts.
Povinelli, L. A.↗
Cold-air performance of compressor-drive turbine of Department of Energy upgraded automobile gas turbine engine. 2: Stage performance
The aerodynamic performance of the compressor-drive turbine of the DOE upgraded gas turbine engine was determined in low temperature air. The as-received cast rotor blading had a significantly thicker profile than design and a fairly rough surface finish. Because of these blading imperfections a series of stage tests with modified rotors were made. These included the as-cast rotor, a reduced-roughness rotor, and a rotor with blades thinned to near design. Significant performance changes were measured. Tests were also made to determine the effect of Reynolds number on the turbine performance. Comparisons are made between this turbine and the compressor-drive turbine of the DOE baseline gas turbine engine.
Roelke, R. J.↗
Analytical and experimental investigation of stator endwall contouring in a small axial-flow turbine. 1: Stator performance
Three stator configurations were studied to determine the effect of stator outer endwall contouring on stator performance. One configuration was a cylindrical stator design. One contoured stator configuration had an S-shaped outer endwall, the other had a conical-convergent outer endwall. The experimental investigation consisted of annular surveys of stator exit total pressure and flow angle for each stator configuration over a range of stator pressure ratio. Radial variations in stator loss and aftermixed flow conditions were obtained when these data were compared with the analytical results to assess the validity of the analysis, good agreement was found.
Haas, J. E.↗
Experimental evaluation of inlet turbulence, wall boundary layer, surface finish, and fillet radius on small axial turbine state performance
The results of an experimental investigation are presented which establish the effects of surface finish, fillet radius, inlet boundary layer thickness, and free-stream inlet turbulence level on the aerodynamic performance of a small axial flow turbine stator. The principal objective was to help understand why large turbine efficiency is not maintained when a large turbine is scaled to a smaller size and to provide the turbine designer with the performance compromises expected for a small scale design. A comprehensive test matrix was used to gain an understanding of the effects of each variable over the full range of all the other variables.
Kozak, A. A.↗
Comparison of experimental and analytic performance for contoured endwall stators
(Previously announced in STAR as N82-26299)
Boyle, R. J.↗
The effect of rotor blade thickness and surface finish on the performance of a small axial flow turbine
(Previously announced in STAR as N82-13114)
Roelke, R. J.↗
The effect of rotor blade thickness and surface finish on the performance of a small axial flow turbine
An experimental investigation was conducted to determine the effect of blade profile inaccuracies and surface finish on the aerodynamic performance of a 11.13 cm tip diameter turbine. The as-received cast rotor blades had a significantly thicker profile than the design intent and a fairly rough surface finish. Stage test results showed an increase of one point in efficieny by smoothing the surface finish and another three points by thinning the blade profiles to near the design profile. Most of the performance gain between the as-cast thick and the thinned rotor blades both with the same surface finish, was attributed to reduced trailing edge losses of the recontoured blades.
Roelke, R. J.↗
Comparison of Experimental and Analytical Performance for Contoured Endwall Stators
Comparisons between predicted and experimental stator losses showed that the analysis was able to predict the change in stator loss when contoured endwalls with highly three dimensional passage geometry were used. The level of loss was predicted to within 75 percent of that measured. The predicted loss was due only to profile loss and boundary layer growth on the endwalls. The 25 percent difference was approximately 0.015 at design pressure ratio. The analysis was shown to predict the trend in stator flow angle, even for small stator geometries.
Boyle, R. J.↗
Cold-air performance of compressor-drive turbine of Department of Energy upgraded automobile gas turbine engine. 1: Volute-manifold and stator performance
The aerodynamic performance of the inlet manifold and stator assembly of the compressor drive turbine was experimentally determined with cold air as the working fluid. The investigation included measurements of mass flow and stator-exit fluid torque as well as radial surveys of total pressure and flow angle at the stator inlet and annulus surveys of total pressure and flow angle at the stator exit. The stator-exit aftermixed flow conditions and overall stator efficiency were obtained and compared with their design values and the experimental results from three other stators. In addition, an analysis was made to determine the constituent aerodynamic losses that made up the stator kinetic energy loss.
Roelke, R. J.↗
Reasons for low aerodynamic performance of 13.5-centimeter-tip-diameter aircraft engine starter turbine
The reasons for the low aerodynamic performance of a 13.5 cm tip diameter aircraft engine starter turbine were investigated. Both the stator and the stage were evaluated. Approximately 10 percent improvement in turbine efficiency was obtained when the honeycomb shroud over the rotor blade tips was filled to obtain a solid shroud surface. Efficiency improvements were obtained for three rotor configurations when the shroud was filled. It is suggested that the large loss associated with the open honeycomb shroud is due primarily to energy loss associated with gas transportation as a result of the blade to blade pressure differential at the tip section.
Haas, J. E.↗
Experimental performance and analysis of 15.04-centimeter-tip-diameter, radial-inflow turbine with work factor of 1.126 and thick blading
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.
Mclallin, K. L.↗
An experimental evaluation of the performance deficit of an aircraft engine starter turbine
An experimental investigation is presented to determine the aerodynamic performance deficit of a 13.5 - centimeter-tip-diameter aircraft engine starter turbine. The two-phased evaluation comprised both the stator and the stage performance, and the experimental design is described in detail. Data obtained from the investigation of three honeycomb shrouds clearly showed that the filled honeycomb reached a total efficiency of 0.868, 8.2 points higher than the open honeycomb shroud, at design equivalent conditions of speed and blade-jet speed ratio. It was concluded that the use of an open honeycomb shroud caused the large performance deficit for the starter turbine. Further research is suggested to ascertain stator inlet boundary layer measurements.
Haas, J. E.↗