Aerodynamic Performance and
Turbulence Measurements in a Turbine Vane Cascade
Boyle, Robert J. (NASA Glenn Research
Center, Cleveland, OH United States); Lucci, Barbara
L. (NASA Glenn Research Center, Cleveland, OH United States); Senyitko, Richard G. (QSS Group, Inc., Cleveland, OH United
States)
NASA Center for AeroSpace Information (CASI)
NASA/TM-2002-211709
Turbine vane aerodynamics were measured in a three vane linear cascade. Surface
pressures and blade row losses were obtained over a range of Reynolds and Mach
number for three levels of turbulence. Comparisons are made with predictions
using a quasi-3D Navier-Stokes analysis. Turbulence
intensity measurement were made upstream and downstream of the vane. The
purpose of the downstream measurements was to determine how the turbulence was
affected by the strong contraction through 75 deg turning.
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Effect of Favorable Pressure
Gradients on Turbine Blade Pressure Surface Heat Transfer
Boyle, Robert J. (NASA Glenn Research
Center, Cleveland, OH United States); Giel, P. W.
(DYNACS Engineering Co., Inc., Cleveland, OH United States)
NASA Center for AeroSpace Information (CASI)
2002
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.
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Prediction of Relaminarization Effects on Turbine Blade Heat Transfer
Boyle, R. J. (NASA Glenn Research Center, Cleveland, OH
United States); Giel, P. W. (QSS Group, Inc., Brook
Park, OH United States)
NASA Center for AeroSpace Information (CASI)
NASA/TM-2001-210978
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.
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Comparison of Predicted and
Measured Turbine Vane Rough Surface Heat Transfer
Boyle, R. J. (NASA Glenn Research Center, Cleveland, OH
United States); Spuckler, C. M. (NASA Glenn Research
Center, Cleveland, OH United States); Lucci, B. L.
(NASA Glenn Research Center, Cleveland, OH United States)
NASA Center for AeroSpace Information (CASI)
NASA/TM-2000-210219
The proposed paper compares predicted turbine vane heat transfer for a rough
surface over a wide range of test conditions with experimental data.
Predictions were made for the entire vane surface. However, measurements were
made only over the suction surface of the vane, and the leading edge region of
the pressure surface. Comparisons are shown for a wide range of test
conditions. Inlet pressures varied between 3 and 15 psia,
and exit Mach numbers ranged between 0.3 and 0.9. Thus, while a single
roughened vane was used for the tests, the effective rougness,(k(sup
+)), varied by more than a factor of ten. Results were obtained for freestream turbulence levels of 1 and 10%. Heat transfer
predictions were obtained using the Navier-Stokes
computer code RVCQ3D. Two turbulence models, suitable for rough surface
analysis, are incorporated in this code. The Cebeci-Chang
roughness model is part of the algebraic turbulence model. The k-omega
turbulence model accounts for the effect of roughness in the application of the
boundary condition. Roughness causes turbulent flow over the vane surface. Even
after accounting for transition, surface roughness significantly increased heat
transfer compared to a smooth surface. The k-omega results agreed better with
the data than the Cebeci-Chang model. However, the
low Reynolds number k-omega model did not accurately account for roughness when
the freestream turbulence level was low. The high
Reynolds number version of this model was more suitable when the freestream turbulence was low.
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Infrared Low-Temperature
Turbine Vane Rough Surface Heat Transfer Measurements
Boyle, R. J. (NASA Glenn Research Center, Cleveland, OH
United States); Spuckler, C. M. (NASA Glenn Research
Center, Cleveland, OH United States); Lucci, B. L.
(NASA Glenn Research Center, Cleveland, OH United States); Camperchioli,
W. P. (NASA Glenn Research Center, Cleveland, OH United States)
NASA Center for AeroSpace Information (CASI)
Paper-2000-GT-216 , 2001
Turbine vane heat transfer distributions obtained using an infrared camera
technique are described. Infrared thermography was
used because noncontact surface temperature
measurements were desired. Surface temperatures were 80 C or less. Tests were
conducted in a three-vane linear cascade, with inlet pressures between 0.14 and
1.02 atm, and exit Mach numbers of 0.3, 0.7, and 0.9,
for turbulence intensities of approximately 1 and 10 percent. Measurements were
taken on the vane suction side, and on the pressure side leading edge region.
The designs for both the vane and test facility are discussed. The approach
used to account for conduction within the vane is described. Midspan heat transfer distributions are given for the range
of test conditions.
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Mach Number Effects on
Turbine Blade Transition Length Prediction
Boyle, R. J. (NASA Lewis Research Center, Cleveland, OH
United States); Simon, F. F. (NASA Lewis Research Center, Cleveland, OH United
States)
NASA Center for AeroSpace Information (CASI)
NASA/TM-1998-208404
The effect of a Mach number correction on a model for predicting the length of
transition was investigated. The transition length decreases as the turbulent
spot production rate increases. Much of the data for predicting the spot
production rate comes from low speed flow experiments. Recent data and analysis
showed that the spot production rate is affected by Mach number. The degree of
agreement between analysis and data for turbine blade heat transfer without
film cooling is strongly dependent of accurately predicting the length of
transition. Consequently, turbine blade heat transfer data sets were used to
validate a transition length turbulence model. A method for modifying models
for the length of transition to account for Mach number effects is presented.
The modification was made to two transition length models. The modified models
were incorporated into the two-dimensional Navier-Stokes
code, RVCQ3D. Comparisons were made between predicted and measured midspan surface heat transfer for stator and rotor turbine
blades. The results showed that accounting for Mach number effects
significantly improved the agreement with the experimental data.
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Aerodynamics of a
Transitioning Turbine Stator Over a Range of Reynolds Numbers
Boyle, R. J. (NASA Lewis Research Center, Cleveland, OH
United States); Lucci, B. L. (NASA Lewis Research
Center, Cleveland, OH United States); Verhoff, V. G.
(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-1998-208408
Midspan aerodynamic measurements for a three
vane-four passage linear turbine vane cascade are given. The vane axial chord
was 4.45 cm. Surface pressures and loss coefficients were measured at exit Mach
numbers of 0.3, 0.7, and 0.9. Reynolds number was varied by a factor of six at
the two highest Mach numbers, and by a factor of ten at the lowest Mach number.
Measurements were made with and without a turbulence grid. Inlet turbulence
intensities were less than I% and greater than IO%. Length scales were also
measured. Pressurized air fed the test section, and exited to a low pressure
exhaust system. Maximum inlet pressure was two atmospheres. The minimum inlet
pressure for an exit Mach number of 0.9 was one-third of an atmosphere, and at
a Mach number of 0.3, the minimum pressure was half this value. The purpose of
the test was to provide data for verification of turbine vane aerodynamic
analyses, especially at low Reynolds numbers. Predictions obtained using a Navier-Stokes analysis with an algebraic turbulence model
are also given.
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Prediction of Nonuniform
Inlet Temperature Effects on Vane and Rotor Heat Transfer
Boyle, Robert J. (NASA Lewis Research Center,
Cleveland, OH United States); Giel, Paul W. (NYMA,
Inc., Brook Park, OH United States)
NASA Center for AeroSpace Information (CASI)
NASA-TM-107539 , 1997
The effects of nonuniform
combustor exit temperature profiles on vane and rotor heat
transfer were determined using a steady-state three-dimensional Navier-Stokes analysis. Both radial and tangential nonuniform temperature profiles
were individually considered. Comparisons are made with experimental data for
the effects of a radial temperature nonuniformity
on rotor heat transfer. There was a decrease in stator heat load, and an
increase in rotor heat load for a radial temperature distribution typically
seen at the combustor exit. Tangential variations in stator inlet temperature
produced significant variations in stator heat load, and resulted in average rotor
heat load greater than for the uniform inlet temperature case. Rotor heat load
was also calculated for different stator wake locations. Accounting for the
stator wake position at the rotor inlet gave a greater average rotor heat load
than that obtained by averaging the stator exit flow field in the tangential
direction. The increase was most notable on the rotor pressure surface.
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Predicted Turbine Heat
Transfer for a Range of Test Conditions
Boyle, R. J. (NASA Lewis Research Center, Cleveland, OH
United States); Lucci, B. L. (NASA Lewis Research
Center, Cleveland, OH United States)
NASA Center for AeroSpace Information (CASI)
NASA-TM-107374 , 1996
Comparisons are shown between predictions and experimental data for blade and endwall heat transfer. The comparisons of computational
domain parisons are given for both vane and rotor
geometries over an extensive range of Reynolds and Mach numbers. Comparisons
are made with experimental data from a variety of sources. A number of
turbulence models are available for predicting blade surface heat transfer, as
well as aerodynamic performance. The results of an investigation to determine the
turbulence model which gives the best agreement with experimental data over a
wide range of test conditions are presented.
No
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Heat transfer predictions for
two turbine nozzle geometries at high Reynolds and Mach numbers
Boyle, R. J. (NASA Lewis Research Center, Cleveland,
OH, United States); Jackson, R. (Defence Research
Agency, Farnborough, Hampshire, United Kingdom)
NASA Center for AeroSpace Information (CASI)
NASA-TM-106956 , 1995
Predictions of turbine vane and endwall heat transfer
and pressure distributions are compared with experimental measurements for two
vane geometries. The differences in geometries were due to differences in the
hub profile, and both geometries were derived from the design of a high rim
speed turbine (HRST). The experiments were conducted in the Isentropic Light
Piston Facility (ILPF) at Pyestock at a Reynolds
number of 5.3 x 10(exp 6), a Mach number of 1.2, and a wall-to-gas temperature
ratio of 0.66. Predictions are given for two different steady-state
three-dimensional Navier-Stokes computational
analyses. C-type meshes were used, and algebraic models were employed to
calculate the turbulent eddy viscosity. The effects of different turbulence
modeling assumptions on the predicted results are examined. Comparisons are
also given between predicted and measured total pressure distributions behind
the vane. The combination of realistic engine geometries and flow conditions
proved to be quite demanding in terms of the convergence of the CFD solutions.
An appropriate method of grid generation, which resulted in consistently
converged CFD solutions, was identified.
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Grid orthogonality
effects on predicted turbine midspan heat transfer
and performance
Boyle, R. J. (NASA Lewis Research Center, Cleveland, OH,
United States); Ameri, A. A. (Kansas Univ., Lawrence,
KS., United States)
NASA Center for AeroSpace Information (CASI)
NASA-TM-106931 , 1995
The effect of five different C type grid geometries on the predicted heat
transfer and aerodynamic performance of a turbine stator is examined.
Predictions were obtained using two flow analysis codes. One was a finite difference
analysis, and the other was a finite volume analysis. Differences among the
grids in terms of heat transfer and overall performance were small. The most
significant difference among the five grids occurred in the prediction of pitchwise variation in total pressure. There was
consistency between results obtained with each of the flow analysis codes when
the same grid was used. A grid generating procedure in which the viscous grid
is embedded within an inviscid type grid resulted in
the best overall performance.
No
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Prediction of incidence and
surface roughness effects on turbine performance
Boyle, R. J. (NASA Lewis Research Center, Cleveland,
OH, United States)
NASA Center for AeroSpace Information (CASI)
1993
The results of a Navier-Stokes analysis for predicting
the change in turbine efficiency due to a change in either incidence or surface
roughness is discussed. It was experimentally determined by Boynton, Tabibzadeh, and Hudson that polishing the SSME high
pressure fuel turbine blades improved turbine efficiency by about 2 points over
a wide range of operating conditions. These conditions encompassed the range of
incidence seen by the turbine blading during flight.
It is also necessary to be able to predict turbine performance at various
operating points for future rocket turbopump
applications. The code RVCQ3D, developed by Rod Chima,
was used to determine the effects of changes in incidence angle on turbine
blade row efficiency. The midspan Navier-Stokes
results were used in conjunction with an inviscid flow
analysis code to predict the efficiency of the two stage SSME over a wide range
of operating conditions for smooth and rough turbine blades. The use of the Navier-Stokes analysis to predict changes in turbine
efficiency due to variation in incidence angles was found to be superior to
other incidence loss correlations available in the literature. The sensitivity
of the Navier-Stokes results to grid parameters is
discussed. The effects of the surface roughness were accounted for using the Cebeci-Chang rough wall turbulence model. This model was
implemented in the code RVCQ3D. The implementation of this model for predicting
the change in efficiency is also discussed.
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Three-dimensional Navier-Stokes heat transfer predictions for turbine blade
rows
Boyle, Robert J. (NASA Lewis Research
Center, Cleveland, OH, United States); Giel, Paul W.
(Sverdrup Technology, Inc., Brook Park, OH., United
States)
NASA Center for AeroSpace Information (CASI)
NASA-TM-105800 , 1992
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.
No
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Two-dimensional Navier-Stokes heat transfer analysis for rough turbine
blades
Boyle, R. J. (NASA Lewis Research Center, Cleveland,
OH, United States); Civinskas, K. C. (NASA Lewis
Research Center, Cleveland, OH, United States)
NASA Center for AeroSpace Information (CASI)
NASA-TM-106008 , 1991
A quasi-three-dimensional thin-layer Navier-Stokes
analysis was used to predict heat transfer to rough surfaces. Comparisons are
made between predicted and experimental heat transfer for turbine blades and
flat plates of known roughness. The effect of surface roughness on heat
transfer was modeled using a mixing length approach. The effect of near-wall
grid spacing and convergence criteria on the accuracy of the heat transfer
predictions are examined. An eddy viscosity mixing length model having an inner
and outer layer was used. A discussion of the appropriate model for the
crossover between the inner and outer layers is included. The analytic results
are compared with experimental data for both flat plates and turbine blade
geometries. Comparisons between predicted and experimental heat transfer showed
that a modeling roughness effects using a modified mixing length approach
results in good predictions of the trends in heat transfer due to roughness.
No
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Navier-Stokes analysis of turbine blade heat transfer
Boyle, R. J. (NASA Lewis Research Center, Cleveland,
OH, United States)
NASA Center for AeroSpace Information (CASI)
NASA-TM-102496 , 1990
Comparisons with experimental heat transfer and surface pressures were made for
seven turbine vane and blade geometries using a quasi-three-dimensional
thin-layer Navier-Stokes analysis. Comparisons are
made for cases with both separated and unseparated flow
over a range of Reynolds numbers and freestream
turbulence intensities. The analysis used a modified Baldwin-Lomax turbulent eddy viscosity mode. Modifications were
made to account for the effects of: (1) freestream
turbulence on both transition and leading edge heat transfer; (2) strong
favorable pressure gradients on relaminarization; and
(3) variable turbulent Prandtl number heat transfer.
In addition, the effect of heat transfer on the near wall model of Deissler is compared with the Van Driest model.
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Experimental determination of
stator endwall heat transfer
Boyle, Robert J. (NASA
Lewis Research Center, Cleveland, OH, United States); Russell, Louis M. (NASA
Lewis Research Center, Cleveland, OH, United States)
NASA Center for AeroSpace Information (CASI)
NASA-TM-101419 , 1989
Local Stanton numbers were experimentally determined for the endwall surface of a turbine vane passage. A six vane
linear cascade having vanes with an axial chord of 13.81 cm was used. Results
were obtained for Reynolds numbers based on inlet velocity and axial chord
between 73,000 and 495,000. The test section was connected to a low pressure
exhaust system. Ambient air was drawn into the test section, inlet velocity was
controlled up to a maximum of 59.4 m/sec. The effect of the inlet boundary
layer thickness on the endwall heat transfer was
determined for a range of test section flow rates. The liquid crystal
measurement technique was used to measure heat transfer. Endwall
heat transfer was determined by applying electrical power to a foil heater
attached to the cascade endwall. The temperature at
which the liquid crystal exhibited a specific color was known from a
calibration test. Lines showing this specific color were isotherms, and because
of uniform heat generation they were also lines of nearly constant heat
transfer. Endwall static pressures were measured,
along with surveys of total pressure and flow angles at the inlet and exit of
the cascade.
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