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Whitney, W. J.

Publications and source records attributed to Whitney, W. J..

At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1984

Performance of a high-work low aspect ratio turbine tested with a realistic inlet radial temperature profile

Experimental results are presented for a 0.767 scale model of the first stage of a two-stage turbine designed for a high by-pass ratio engine. The turbine was tested with both uniform inlet conditions and with an inlet radial temperature profile simulating engine conditions. The inlet temperature profile was essentially mixed-out in the rotor. There was also substantial underturning of the exit flow at the mean diameter. Both of these effects were attributed to strong secondary flows in the rotor blading. There were no significant differences in the stage performance with either inlet condition when differences in tip clearance were considered. Performance was very close to design intent in both cases. Previously announced in STAR as N84-24589

Stabe, R. G.↗
NASA NTRS1984

Performance of a high-work low aspect ration turbine tested with a realistic inlet radial temperature profile

Experimental results are presented for a 0.767 scale model of the first stage of a two-stage turbine designed for a high by-pass ratio engine. The turbine was tested with both uniform inlet conditions and with an inlet radial temperature profile simulating engine conditions. The inlet temperature profile was essentially mixed-out in the rotor. There was also substantial underturning of the exit flow at the mean diameter. Both of these effects were attributed to strong secondary flows in the rotor blading. There were no significant differences in the stage performance with either inlet condition when differences in tip clearance were considered. Performance was very close to design intent in both cases.

Stabe, R. G.↗
NASA NTRS1983

Analytical and experimental study of flow through an axial turbine stage with a nonuniform inlet radial temperature profile

Results are presented for a typical nonuniform inlet radial temperature profile through an advanced single-stage axial turbine and compared with the results obtained for a uniform profile. Gas temperature rises of 40 K to 95 K are predicted at the hub and tip corners at the trailing edges of the pressure surfaces in both the stator and rotor due to convection of hot fluid from the mean by the secondary flow. The inlet temperature profile is shown to be mixed out at the rotor exit survey plane (2.3 axial chords downstream of the rotor trailing edge) in both the analysis and the experiment. The experimental rotor exit angle profile for the nonuniform inlet temperature profile indicates underturning at the tip caused by increased clearance. Severe underturning also occurs at the mean, both with and without the nonuniform inlet temperature profile. The inviscid rotational flow code used in the analysis fails to predict the underturning at the mean, which may be caused by viscous effects. Previously announced in STAR as N83-27958

Schwab, J. R.↗
NASA NTRS1983

Aerodynamic effect of a honeycomb rotor tip shroud on a 50.8-centimeter-tip-diameter core turbine

A 50.8-cm-tip-diameter turbine equipped with a rotor tip shroud of hexagonal cell (or honeycomb) cross section has been tested in warm air (416 K) for a range of shroud coolant to primary flow rates. Test results were also obtained for the same turbine operated with a solid shroud for comparison. The results showed that the combined effect of the honeycomb shroud and the coolant flow was to cause a reduction of 2.8 points in efficiency at design speed, pressure ratio, and coolant flow rate. With the coolant system inactivated, the honeycomb shroud caused a decrease in efficiency of 2.3 points. These results and those obtained from a small reference turbine indicate that the dominant factor governing honeycomb tip shroud loss is the ratio of honeycomb depth to blade span. The loss results of the two shrouds could be correlated on this basis. The same honeycomb and coolant effects are expected to occur for the hot (2200 K) version of this turbine.

Moffitt, T. P.↗
NASA NTRS1983

Analytical and Experimental Study of Flow Through an Axial Turbine Stage with a Nonuniform Inlet Radial Temperature Profile

Results are presented for a typical nonuniform inlet radial temperature profile through an advanced single-stage axial turbine and compared with the results obtained for a uniform profile. Gas temperature rises of 40 K to 95 K are predicted at the hub and tip corners at the trailing edges of the pressure surfaces in both the stator and rotor due to convection of hot fluid from the mean by the secondary flow. The inlet temperature profile is shown to be mixed out at the rotor exit survey plane (2.3 axial chords downstream of the rotor trailing edge) in both the analysis and the experiment. The experimental rotor exit angle profile for the nonuniform inlet temperature profile indicates underturning at the tip caused by increased clearance. Severe underturning also occurs at the mean, both with and without the nonuniform inlet temperature profile. The inviscid rotational flow code used in the analysis fails to predict the underturning at the mean, which may be caused by viscous effects.

Schwab, J. R.↗
NASA NTRS1981

Cold-air investigation of first stage of 4-1/2-stage, fan drive turbine with average stage-loading factor of 4.66

The design procedure and the development of the blading geometry for the 4 and 1/2 stage turbine are discussed. Results obtained with the first stage, operated as a single stage turbine, are presented. A free vortex design meets the design requirements without incurring problems such as excessive turning, negative reaction, or high Mach number. Cold air tests of the single stage turbine showed that the turbine developed design work (stage loading factor of 5.26) at an efficiency of 0.86, which was the efficiency predicted by a reference method. The mass flow at this condition was 0.88, which occurred at design speed and a pressure ratio of 1.407, corresponding to a stage loading factor of 4.35. The efficiency at this condition was 0.003 higher than that predicted by the reference method.

Whitney, W. J.↗
NASA NTRS1980

Description of the warm core turbine facility and the warm annular cascade facility recently installed at NASA Lewis Research Center

The two new facilities have been installed and operated at their design or rated conditions. The important feature of both of these facilities is that the ratio of turbine inlet temperature to coolant temperature encountered in high temperature engines can be duplicated at moderate turbine inlet temperature. Included in the discussion are the limits of the facilities with regard to maximum temperature, maximum pressure, maximum mass flow rate, turbine size, and dynamometer torque-speed characteristics.

Whitney, W. J.↗
NASA NTRS1980

Design and cold-air test of single-stage uncooled turbine with high work output

A solid version of a 50.8 cm single stage core turbine designed for high temperature was tested in cold air over a range of speed and pressure ratio. Design equivalent specific work was 76.84 J/g at an engine turbine tip speed of 579.1 m/sec. At design speed and pressure ratio, the total efficiency of the turbine was 88.6 percent, which is 0.6 point lower than the design value of 89.2 percent. The corresponding mass flow was 4.0 percent greater than design.

Moffitt, T. P.↗
NASA NTRS1980

Cold-air investigation of a 4 1/2 stage turbine with stage-loading factor of 4.66 and high specific work output. 2: Stage group performance

The stage group performance of a 4 1/2 stage turbine with an average stage loading factor of 4.66 and high specific work output was determined in cold air at design equivalent speed. The four stage turbine configuration produced design equivalent work output with an efficiency of 0.856; a barely discernible difference from the 0.855 obtained for the complete 4 1/2 stage turbine in a previous investigation. The turbine was designed and the procedure embodied the following design features: (1) controlled vortex flow, (2) tailored radial work distribution, and (3) control of the location of the boundary-layer transition point on the airfoil suction surface. The efficiency forecast for the 4 1/2 stage turbine was 0.886, and the value predicted using a reference method was 0.862. The stage group performance results were used to determine the individual stage efficiencies for the condition at which design 4 1/2 stage work output was obtained. The efficiencies of stages one and four were about 0.020 lower than the predicted value, that of stage two was 0.014 lower, and that of stage three was about equal to the predicted value. Thus all the stages operated reasonably close to their expected performance levels, and the overall (4 1/2 stage) performance was not degraded by any particularly inefficient component.

Whitney, W. J.↗
NASA NTRS1980

Description of the warm core turbine facility recently installed at NASA Lewis Research Center

The two net facilities were installed and operated at their design, or rated conditions. The important feature of both of these facilities is that the ratio of turbine inlet temperature to coolant temperature encountered in high temperature engines can be duplicated at moderate turbine inlet temperature. The limits of the facilities with regard to maximum temperature, maximum pressure, maximum mass flow rate, turbine size, and dynamometer torque-speed characteristics are discussed.

Whitney, W. J.↗
NASA NTRS1977

Cold-air investigation of a 31/2-stage fan-drive turbine with a stage loading factor of 4 designed for an integral lift engine. 2: Performance of 2-, 3- and 3 1/2-stage configurations

The stage work distribution among the three stages was very close to the design value. The specific work output-mass flow characteristics of the three stages were closely matched. The efficiency of the 3 1/2 stage turbine at design specific work output and design speed was within 0.008 of the estimated value, and this agreement was felt to demonstrate the adequacy of the prediction method in the high stage loading factor regime.

Whitney, W. J.↗
NASA NTRS1977

Cold air investigation of 4 1/2-stage turbine with stage loading factor of 4.66 and high specific work output. 1: Overall performance

The turbine developed design specific work output at design speed at a total pressure ratio of 6.745 with a corresponding efficiency of 0.855. The efficiency (0.855)was 3.1 points lower than the estimated efficiency quoted by the contractor in the design report and 0.7 of a point lower than that determined by a reference prediction method. The performance of the turbine, which was a forced vortex design, agreed with the performance determined by the prediction method to about the same extent as did the performance of three reference high stage loading factor turbines, which were free vortex designs.

Whitney, W. J.↗
NASA NTRS1975

Cold-air investigation of a 3 1/2-stage fan-drive turbine with a stage loading factor of 4 designed for an integral lift engine. 1: Turbine design and performance of first stage

The design of the 3 1/2-stage turbine is described, and the cold-air performance of the first stage, modified for axial inlet conditions, is presented. The performance of the modified single-stage turbine and of two comtemporary high-stage-loading-factor turbines is compared with that estimated with a reference prediction method.

Whitney, W. J.↗
NASA NTRS1972

Cold-air investigation of a turbine for high-temperature-engine application. 4: Two-stage turbine performance

The efficiency of a two-stage turbine is discussed. The turbine efficiency was 0.932 for equivalent design operating conditions (speed) and specific work, which compares closely to the value of 0.929 that would be estimated using the first-stage efficiency. The mass flow obtained with the two-stage configuration indicated that the mass flow characteristics of the two stages were closely matched at design operating conditions. The stage work split at these conditions was 0.505-0.495, which was close to the design work split of 0.515-0.485.

Whitney, W. J.↗
NASA NTRS1972

Drive turbine system for 20-inch turbofan simulators. 2: Core turbine design

Drive-turbines for a given set of 20-inch turbo-fan simulators are described. The simulators had both single-stage and two-stage fans that had design pressure ratios as low as 1.25 and as high as 3.0. The desired objective of the study was to be able to drive all of the single-stage fans with one core turbine and to drive all of the two-stage fans with this same core turbine in combination with a duct turbine. The core turbine is described. Included are the design operating conditions, design velocity diagram and a power-speed envelope determined by an off-design performance procedure. Also discussed is the adaption and scaling of an existing turbine design to this particular application.

Whitney, W. J.↗
NASA NTRS1972

Drive turbine systems for 20-inch turbofan simulators. 1: Duct turbine design

A study was made to evolve the turbine drive systems for 20-inch turbofan engine simulators. The fan designs used in the simulators included single-stage and two-stage configurations that covered a wide range of rotative speed and power requirement. The objective assumed for the study was to evolve one core turbine design that could drive all of the single-stage fans and, when operated in combination with one duct turbine design, drive all of the two-stage fans. The duct turbine power output is then needed to determine the make-up power required of the core turbine over the range of two-stage fan operating conditions. The duct turbine design analysis is reported and includes the selection of the duct turbine velocity diagram, a description of the blade design, and a determination of its off-design performance. Adjustable stators were found to be quite advantageous to the duct turbine off-design operation. The use of adjustable stators enabled the duct turbine to accommodate fan mass flow at all operating points and caused the duct turbine power output to increase as the total power requirement increased. This in turn resulted in a core turbine make-up power requirement that was not significantly greater than that required for driving the single-stage fans.

Whitney, W. J.↗
NASA NTRS1972

Velocity diagrams

The selection and design of velocity diagrams for axial flow turbines are considered. Application is treated in two parts which includes: (1) mean-section diagrams, and (2) radial variation of diagrams. In the first part, the velocity diagrams occurring at the mean section are assumed to represent the average conditions encountered by the turbine. The different types of diagrams, their relation to stage efficiency, and their selection when staging is required are discussed. In the second part, it is shown that in certain cases the mean-section diagrams may or may not represent the average flow conditions for the entire blade span. In the case of relatively low hub- to tip-radius ratios, substantial variations in the velocity diagrams are encountered. The radial variations in flow conditions and their effect on the velocity diagrams are considered.

Whitney, W. J.↗