Engineering topics

Giel, P. W.

Publications and source records attributed to Giel, P. W..

10 records

NASA NTRS2002

Effect of Favorable Pressure Gradients on Turbine Blade Pressure Surface Heat Transfer

Recent measurements on a turbine rotor showed significant relaminarization effects. These effects were evident on the pressure surface heat transfer measurements. The character of the heat transfer varied with Reynolds number. Data were obtained for exit Reynolds numbers between 500,000 and 880,000. Tests were done with a high level of inlet turbulence, 7.5%. At lower Reynolds numbers the heat transfer was similar to that for laminar flow, but at a level higher than for laminar flow. At higher Reynolds numbers the heat transfer was similar to turbulent flow, when the acceleration parameter, K, was sufficiently small. The proposed paper discusses the experimental results, and also discusses approaches to calculating the surface heat transfer for the blade surface. Calculations were done using a three-dimensional Navier-Stokes CFD analysis. The results of these tests, when compared with previous blade tests in the same facility, illustrate modeling difficulties that were encountered in CFD predictions. The two blades were in many ways similar. However, the degree of agreement between the same analysis and the experimental data was significantly different. These differences are highlighted to illustrate where improvements in modeling approaches are needed for transitional flows.

Boyle, Robert J.↗

NASA NTRS2001

Prediction of Relaminarization Effects on Turbine Blade Heat Transfer

An approach to predicting turbine blade heat transfer when turbulent flow relaminarizes due to strong favorable pressure gradients is described. Relaminarization is more likely to occur on the pressure side of a rotor blade. While stators also have strong favorable pressure gradients, the pressure surface is less likely to become turbulent at low to moderate Reynolds numbers. Accounting for the effects of relaminarization for blade heat transfer can substantially reduce the predicted rotor surface heat transfer. This in turn can lead to reduced rotor cooling requirements. Two-dimensional midspan Navier-Stokes analyses were done for each of eighteen test cases using eleven different turbulence models. Results showed that including relaminarization effects generally improved the agreement with experimental data. The results of this work indicate that relatively small changes in rotor shape can be utilized to extend the likelihood of relaminarization to high Reynolds numbers. Predictions showing how rotor blade heat transfer at a high Reynolds number can be reduced through relaminarization are given.

Boyle, R. J.↗

NASA NTRS1999

Blade Heat Transfer Measurements and Prediction in a Transonic Turbine Cascade

Detailed heat transfer measurements and predictions are given for a turbine rotor with 136 deg of turning and an axial chord of 12.7 cm. Data were obtained for inlet Reynolds numbers of 0.5 and 1.0 x 10(exp 6), for isentropic exit Mach numbers of 1.0 and 1.3, and for inlet turbulence intensities of 0.25% and 7.0%. Measurements were made in a linear cascade having a highly three-dimensional flow field resulting from thick inlet boundary layers. The purpose of the work is to provide benchmark quality data for three-dimensional CFD code and model verification. Data were obtained by a steady-state technique using a heated, isothermal blade. Heat fluxes were determined from a calibrated resistance layer in conjunction with a surface temperature measured by calibrated liquid crystals. The results show the effects of strong secondary vortical flows, laminar-to-turbulent transition, shock impingement, and increased inlet turbulence on the surface heat transfer.

Giel, P. W.↗

NASA NTRS1996

Endwall Heat Transfer Measurements in a Transonic Turbine Cascade

Turbine blade endwall heat transfer measurements are given for a range of Reynolds and Mach numbers. Data were obtained for Reynolds numbers based on inlet conditions of 0.5 and 1.0 x 106, for isentropic exit Mach numbers of 1.0 and 1.3, and for freestream turbulence intensities of 0.25% and 7.0%. Tests were conducted in a linear cascade at the NASA Lewis Transonic Turbine Blade Cascade Facility. The test article was a turbine rotor with 136' of turning and an axial chord of 12.7 cm. The large scale allowed for very detailed measurements of both flow field and surface phenomena. The intent of the work is to provide benchmark quality data for computational fluid dynamics (CFD) code and model verification. The flow field in the cascade is highly three-dimensional as a result of thick boundary layers at the test section inlet. Endwall heat transfer data were obtained using a steady-state liquid crystal technique.

Giel, P. W.↗

NASA NTRS1996

Three-Dimensional Flow Field Measurements in a Transonic Turbine Cascade

Three-dimensional flow field measurements are presented for a large scale transonic turbine blade cascade. Flow field total pressures and pitch and yaw flow angles were measured at an inlet Reynolds number of 1.0 x 10(exp 6) and at an isentropic exit Mach number of 1.3 in a low turbulence environment. Flow field data was obtained on five pitchwise/spanwise measurement planes, two upstream and three downstream of the cascade, each covering three blade pitches. Three-hole boundary layer probes and five-hole pitch/yaw probes were used to obtain data at over 1200 locations in each of the measurement planes. Blade and endwall static pressures were also measured at an inlet Reynolds number of 0.5 x 10(exp 6) and at an isentropic exit Mach number of 1.0. Tests were conducted in a linear cascade at the NASA Lewis Transonic Turbine Blade Cascade Facility. The test article was a turbine rotor with 136 deg of turning and an axial chord of 12.7 cm. The flow field in the cascade is highly three-dimensional as a result of thick boundary layers at the test section inlet and because of the high degree of flow turning. The large scale allowed for very detailed measurements of both flow field and surface phenomena. The intent of the work is to provide benchmark quality data for CFD code and model verification.

Giel, P. W.↗

NASA NTRS1996

Three-dimensional Navier-Stokes analysis and redesign of an imbedded bellmouth nozzle in a turbine cascade inlet section

Experimental measurements in the inlet of a transonic turbine blade cascade showed unacceptable pitchwise flow non-uniformity. A three-dimensional, Navier-Stokes computational fluid dynamics (CFD) analysis of the imbedded bellmouth inlet in the facility was performed to identify and eliminate the source of the flow non-uniformity. The blockage and acceleration effects of the blades were accounted for by specifying a periodic static pressure exit condition interpolated from a separate three-dimensional Navier-Stokes CFD solution of flow around a single blade in an infinite cascade. Calculations of the original inlet geometry showed total pressure loss regions consistent in strength and location to experimental measurements. The results indicate that the distortions were caused by a pair of streamwise vortices that originated as a result of the interaction of the flow with the imbedded bellmouth. Computations were performed for an inlet geometry which eliminated the imbedded bellmouth by bridging the region between it and the upstream wall. This analysis indicated that eliminating the imbedded bellmouth nozzle also eliminates the pair of vortices, resulting in a flow with much greater pitchwise uniformity. Measurements taken with an installed redesigned inlet verify that the flow non-uniformity has indeed been eliminated.

VanFossen, G. J.↗

NASA NTRS1993

Three-dimensional Navier-Stokes analysis and redesign of an imbedded bellmouth nozzle in a turbine cascade inlet section

Verification of proposed turbopump blading performance will involve evaluation of candidate blades in cascade test facilities. It is necessary to be able to predict the flow fields within these cascades for the results to be applicable to actual engine environments. This work presents the results of a study to predict the flow field for the NASA Lewis Transonic Turbine Blade Cascade Facility, which is similar to those used to evaluate rocket propulsion turbines. A pitchwise nonuniform total pressure distribution was observed at the blade row leading edge plane. A CFD analysis was used to show that the cause of the flow nonuniformity was a pair of vortices that originated in an embedded bellmouth inlet. Further CFD analysis was used to verify that a redesigned inlet section resulted in a flow with acceptable uniformity. A computational analysis was chosen because physical accessibility to the inlet section was limited, and because a computational approach also allows one to examine design changes cheaper and more quickly than an experimental approach would. The PARC code, a general purpose, three-dimensional, Navier-Stokes code with multiblock solution capability, was chosen for the present study. Results are presented detailing the computational requirements needed to accurately predict flows of this nature. Calculations of the original geometry showed total pressure loss regions consistent in strength and in location to experimental measurements. An examination of the results shows that the distortions are caused by a pair of vortices that originate as a result of the interaction of the flow with the imbedded bellmouth. Computations were performed for an inlet geometry which eliminated the imbedded bellmouth by bridging the region between it and the upstream wall. This analysis indicated that eliminating the imbedded bellmouth eliminates the troublesome pair of vortices, resulting in a flow with much greater pitchwise uniformity.

Giel, P. W.↗

NASA NTRS1993

An algebraic turbulence model for three-dimensional viscous flows

An algebraic turbulence model is proposed for use with three-dimensional Navier-Stokes analyses. It incorporates features of both the Baldwin-Lomax and Cebeci-Smith models. The Baldwin-Lomax model uses the maximum of a function f(y) to determine length and velocity scales. An analysis of the Baldwin-Lomax model shows that f(y) can have a spurious maximum close to the wall, causing numerical problems and non-physical results. The proposed model uses integral relations to determine delta(*) u(sub e) and delta used in the Cebeci-Smith mode. It eliminates a constant in the Baldwin-Lomax model and determines the two remaining constants by comparison to the Cebeci-Smith formulation. Pressure gradient effects, a new wake model, and the implementation of these features in a three-dimensional Navier-Stokes code are also described. Results are shown for a flat plate boundary layer, an annular turbine cascade, and endwall heat transfer in a linear turbine cascade. The heat transfer results agree well with experimental data which shows large variations in endwall Stanton number contours with Reynolds number.

Chima, R. V.↗

NASA NTRS1993

An algebraic turbulence model for three-dimensional viscous flows

An algebraic turbulence model is proposed for use with three-dimensional Navier-Stokes analyses. It incorporates features of both the Baldwin-Lomax and Cebeci-Smith models. The Baldwin-Lomax model uses the maximum of a function f(v) to determine length and velocity scales. An analysis of the Baldwin-Lomax model shows that f(v) can have a spurious maximum close to the wall, causing numerical problems and non-physical results. The proposed model uses integral relations to determine delta(*) u(sub e) and delta used in the Cebeci-Smith mode. It eliminates a constant in the Baldwin-Lomax model and determines the two remaining constants by comparison to the Cebeci-Smith formulation. Pressure gradient effects, a new wake model, and the implementation of these features in a three-dimensional Navier-Stokes code are also described. Results are shown for a flat plate boundary layer, an annular turbine cascade, and endwall heat transfer in a linear turbine cascade. The heat transfer results agree well with experimental data which shows large variations in endwall Stanton number contours with Reynolds number.

Chima, R. V.↗

NASA NTRS1992

Three-dimensional Navier-Stokes heat transfer predictions for turbine blade rows

Results are shown for a three-dimensional Navier-Stokes analysis of both the flow and the surface heat transfer for turbine applications. Heat transfer comparisons are made with the experimental shock-tunnel data of Dunn and Kim, and with the data of Blair for the rotor of the large scale rotating turbine. The analysis was done using the steady-state, three-dimensional, thin-layer Navier-Stokes code developed by Chima, which uses a multistage Runge-Kutta scheme with implicit residual smoothing. An algebraic mixing length turbulence model is used to calculate turbulent eddy viscosity. The variation in heat transfer due to variations in grid parameters is examined. The effects of rotation, tip clearance, and inlet boundary layer thickness variation on the predicted blade and endwall heat transfer are examined.

Boyle, R. J.↗