Heat Transfer Measurements
and Predictions on a Power Generation Gas Turbine Blade
Giel, Paul W. (DYNACS
Engineering Co., Inc., Brook Park, OH United States); Bunker, Ronald S.
(General Electric Co., Corporate Research and Development Schenectady, NY
United States); VanFossen, G. James (NASA Glenn
Research Center, Cleveland, OH United States); Boyle, Robert J. (NASA Glenn
Research Center, Cleveland, OH United States)
NASA Center for AeroSpace Information (CASI)
NASA/TM-2000-210021
Detailed heat transfer measurements and predictions are given for a power
generation turbine rotor with 129 deg of nominal turning and an axial chord of
137 mm. Data were obtained for a set of four exit Reynolds numbers comprised of
the design point of 628,000, -20%, +20%, and +40%. Three ideal exit pressure
ratios were examined including the design point of 1.378, -10%, and +10%. Inlet
incidence angles of 0 deg and +/-2 deg were also examined. Measurements were
made in a linear cascade with highly three-dimensional blade passage flows that
resulted from the high flow turning and thick inlet boundary layers. Inlet
turbulence was generated with a blown square bar grid. The purpose of the work
is the extension of three-dimensional predictive modeling capability for
airfoil external heat transfer to engine specific conditions including blade
shape, Reynolds numbers, and Mach numbers. Data were obtained by a steady-state
technique using a thin-foil heater wrapped around a low thermal conductivity
blade. Surface temperatures were measured using calibrated liquid crystals. The
results show the effects of strong secondary vortical
flows, laminar-to-turbulent transition, and also show good detail in the
stagnation region.
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Blade Heat Transfer
Measurements and Prediction in a Transonic Turbine Cascade
Giel, P. W. (DYNACS
Engineering Co., Inc., Brook Park, OH United States); VanFossen,
G. J. (NASA Glenn Research Center, Cleveland, OH United States); Boyle, R. J.
(NASA Glenn Research Center, Cleveland, OH United States); Thurman, D. R. (Army
Research Lab., Cleveland, OH United States); Civinskas,
K. C. (Army Research Lab., Cleveland, OH United States)
NASA Center for AeroSpace Information (CASI)
NASA/TM-1999-209296
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.
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Three-Dimensional Flow Field
Measurements in a Transonic Turbine Cascade
Giel, P. W. (NYMA, Inc.,
Brook Park, OH United States); Thurman, D. R. (Army Research Lab., Cleveland,
OH United States); Lopez, I. (Army Research Lab., Cleveland, OH United States);
Boyle, R. J. (NASA Lewis Research Center, Cleveland, OH United States); VanFossen, G. J. (NASA Lewis Research Center, Cleveland, OH
United States); Jett, T. A. (NASA Lewis Research Center, Cleveland, OH United
States); Camperchioli, W. P. (NASA Lewis Research
Center, Cleveland, OH United States); La, H. (NASA Lewis Research Center,
Cleveland, OH United States)
NASA Center for AeroSpace Information (CASI)
NASA-TM-107388 , 1996
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.
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Endwall Heat Transfer Measurements in a Transonic Turbine
Cascade
Giel, P. W. (NYMA, Inc.,
Brook Park, OH United States); Thurman, D. R. (Army Research Lab., Cleveland,
OH United States); VanFossen, G. J. (NASA Lewis
Research Center, Cleveland, OH United States); Hippensteele,
S. A. (NASA Lewis Research Center, Cleveland, OH United States); Boyle, R. J.
(NASA Lewis Research Center, Cleveland, OH United States)
NASA Center for AeroSpace Information (CASI)
NASA-TM-107387 , 1996
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.
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Three-dimensional Navier-Stokes analysis and redesign of an imbedded bellmouth nozzle in a turbine cascade inlet section
Giel, P. W. (Sverdrup Technology, Inc., Brook Park, OH, United States); Sirbaugh, J. R. (Sverdrup
Technology, Inc., Brook Park, OH, United States)
NASA Center for AeroSpace Information (CASI)
1993
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.
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