Low-Pressure Turbine
Separation Control: Comparison With Experimental Data
Garg, Vijay K. (Toledo Univ.,
OH United States)
NASA Center for AeroSpace Information (CASI)
NASA/CR-2002-211689
The present work details a computational study, using the Glenn HT code, that
analyzes the use of vortex generator jets (VGJs) to
control separation on a low-pressure turbine (LPT) blade at low Reynolds
numbers. The computational results are also compared with the experimental data
for steady VGJs. It is found that the code determines
the proper location of the separation point on the suction surface of the
baseline blade (without any VGJ) for Reynolds numbers of 50,000 or less. Also,
the code finds that the separated region on the suction surface of the blade
vanishes with the use of VGJs. However, the separated
region and the wake characteristics are not well predicted. The wake width is
generally over-predicted while the wake depth is under-predicted.
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Two-Equation Turbulence
Models for Prediction of Heat Transfer on a Transonic Turbine Blade
Garg, Vijay K. (AYT Corp.,
Brook Park, OH United States); Ameri, Ali A. (AYT
Corp., Brook Park, OH United States)
NASA Center for AeroSpace Information (CASI)
NASA/CR-2000-210810 , 2001
Two versions of the two-equation k-omega model and a shear stress transport
(SST) model are used in a three-dimensional, multi-block, Navier-Stokes
code to compare the detailed heat transfer measurements on a transonic turbine
blade. It is found that the SST model resolves the passage vortex better on the
suction side of the blade, thus yielding a better comparison with the
experimental data than either of the k-w models. However, the comparison is
still deficient on the suction side of the blade. Use of the SST model does
require the computation of distance from a wall, which for a multiblock grid, such as in the present case, can be
complicated. However, a relatively easy fix for this problem was devised. Also
addressed are issues such as (1) computation of the production term in the
turbulence equations for aerodynamic applications, and (2) the relation between
the computational and experimental values for the turbulence length scale, and
its influence on the passage vortex on the suction side of the turbine blade.
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Heat Transfer in Gas Turbines
Garg, Vijay K. (AYT Corp.,
NASA/CR-2001-210942
The turbine gas path is a very complex flow field. This is due to a variety of
flow and heat transfer phenomena encountered in turbine passages. This
manuscript provides an overview of the current work in this field at the
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Modeling Film-Coolant Flow
Characteristics at the Exit of Shower-Head Holes
Garg, Vijay K. (AYT Corp.,
Brook Park, OH United States)
NASA Center for AeroSpace Information (CASI)
NASA/CR-2000-210510
The coolant flow characteristics at the hole exits of a film-cooled blade are
derived from an earlier analysis where the hole pipes and coolant plenum were
also discretized. The blade chosen is the VKI rotor
with three staggered rows of shower-head holes. The present analysis applies
these flow characteristics at the shower-head hole
exits. A multi-block three-dimensional Navier-Stokes
code with Wilcox's k-omega model is used to compute the heat transfer
coefficient on the film-cooled turbine blade. A reasonably good comparison with
the experimental data as well as with the more complete earlier analysis where
the hole pipes and coolant plenum were also gridded is obtained. If the 1/7th power law is assumed for
the coolant flow characteristics at the hole exits, considerable differences in
the heat transfer coefficient on the blade surface, specially in the
leading-edge region, are observed even though the span-averaged values of h
(heat transfer coefficient based on T(sub o)-T(sub w)) match well with the
experimental data. This calls for span-resolved experimental data near
film-cooling holes on a blade for better validation of the code.
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Heat Transfer on a
Film-Cooled Rotating Blade Using Different Turbulence Models
Garg, Vijay K. (AYT Corp.,
1999
A three-dimensional Navier Stokes code has been used
to compute the heat transfer coefficient on a film-cooled, rotating turbine
blade. The blade chosen is the ACE rotor with five rows containing 93 film
cooling holes covering the entire span. This is the only film-cooled rotating
blade over which experimental data is available for comparison. Over 2.278
million grid points are used to compute the flow over the blade including the
tip clearance region. using Wilcox's k-omega model, Coakley's q-omega model, and the zero-equation Baldwin-Lomax (B-L) model. A reasonably good comparison with the
experimental data is obtained on the suction surface for all the turbulence
models. At the leading edge, the B-L model yields a better comparison than tile
two-equation models. On the pressure surface however the comparison between the
experimental data and the prediction from the k-omega model is much better than
from the other two models. Overall, the k-omega model provides the best
comparison with the experimental data. However, the two-equation models require
at least 40% more computational resources than the B-L model.
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Heat Transfer on a
Film-Cooled Rotating Blade
Garg, Vijay K. (AYT Corp.,
NASA/CR-1999-209301
A multi-block, three-dimensional Navier-Stokes code
has been used to compute heat transfer coefficient on the blade, hub and shroud
for a rotating high-pressure turbine blade with 172 film-cooling holes in eight
rows. Film cooling effectiveness is also computed on the adiabatic blade.
Wilcox's k-omega model is used for modeling the turbulence. Of the eight rows
of holes, three are staggered on the shower-head with compound-angled holes.
With so many holes on the blade it was somewhat of a challenge to get a good
quality grid on and around the blade and in the tip clearance region. The final
multi-block grid consists of 4784 elementary blocks which were merged into 276
super blocks. The viscous grid has over 2.2 million cells. Each hole exit, in its true oval shape, has 80 cells within it so
that coolant velocity, temperature, k and omega distributions can be specified
at these hole exits. It is found that for the given parameters, heat transfer
coefficient on the cooled, isothermal blade is highest in the leading edge
region and in the tip region. Also, the effectiveness over the cooled,
adiabatic blade is the lowest in these regions. Results for an uncooled blade are also shown, providing a direct
comparison with those for the cooled blade. Also, the heat transfer coefficient
is much higher on the shroud as compared to that on the hub for both the cooled
and the uncooled cases.
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Heat Transfer on a
Film-Cooled Blade: Effect of Hole Physics
Garg, Vijay K. (AYT Corp., Brook
Park, OH United States); Rigby, David L. (DYNACS Engineering Co., Inc., Brook
Park, OH United States)
NASA Center for AeroSpace Information (CASI)
ASME Paper 98-GT-404 , 1999
A multiblock, three-dimensional Navier
Stokes code has been used to study the within-hole and near-hole physics in
relation to heat transfer on a film-cooled blade. The flow domain consists of
the coolant flow through the plenum and hole-pipes for the three staggered rows
of shower-head holes on the VKI rotor, and the main flow over the blade. A multiblock grid is generated that is nearly orthogonal to
the various surfaces. It may be noted that for the VKI rotor the shower-head
holes are inclined at 30 deg to the spanwise
direction, and are normal to the streamwise direction
on the blade. Wilcox's k-omega turbulence model is used. The present study
provides a much better comparison for the span-averaged heat transfer
coefficient on the blade surface with the experimental data than an earlier
analysis wherein coolant velocity and temperature distributions were specified
at the hole exits rather than extending the computational domain into the
hole-pipe and plenum. Details of the distributions of coolant velocity,
temperature, k and omega at the hole exits are also
presented.
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Heat Transfer on a
Film-Cooled Rotating Blade Using a Two Equation Turbulence Model
Garg, Vijay K. (AYT Corp.,
1998
A three-dimensional Navier-Stokes code has been used
to compare the heat transfer coefficient on a film-cooled, rotating turbine
blade. The blade chosen is the ACE rotor with five rows containing 93 film
cooling holes covering the entire span. This is the only film-cooled rotating
blade over which experimental data is available for comparison. Over 2.278
million grid points are used to compute the flow over the blade including the
tip clearance region, using Coakley's q-omega
turbulence model. Results are also compared with those obtained by Garg and Abhari
(1997) using the zero-equation Baldwin-Lomax (B-L)
model. A reasonably good comparison with the experimental data is obtained on
the suction surface for both the turbulence models. At the leading edge, the
B-L model yields a better comparison than the q-omega model. On the pressure
surface, however, the comparison between the experimental data and the
prediction from either turbulence model is poor. A potential reason for the
discrepancy on the pressure surface could be the presence of unsteady effects
due to stator-rotor interaction in the experiments which are not modeled in the
present computations. Prediction using the two-equation model is in general
poorer than that using the zero-equation model, while the former requires at
least 40% more computational resources.
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Heat Transfer on a
Film-Cooled Blade - Effect of Hole Physics
Garg, Vijay K. (AYT Corp.,
Brook Park, OH United States); Rigby, David L. (NYMA, Inc., Brook Park, OH
United States)
NASA Center for AeroSpace Information (CASI)
NASA/CR-1998-206609
A multi-block, three-dimensional Navier-Stokes code
has been used to study the within-hole and near-hole physics in relation to
heat transfer on a film-cooled blade. The flow domain consists of the coolant
flow through the plenum and hole-pipes for the three staggered rows of
shower-head holes on the VK1 rotor, and the main flow over the blade. A
multi-block grid is generated that is nearly orthogonal to the various
surfaces. It may be noted that for the VK1 rotor the shower-head holes are
inclined at 30 deg. to the spanwise direction, and
are normal to the streamwise direction on the blade.
Wilcox's k-omega turbulence model is used. The present study provides a much
better comparison for the heat transfer coefficient at the blade mid-span with
the experimental data than an earlier analysis wherein coolant velocity and
temperature distributions were specified at the hole exits rather than
extending the computational domain into the hole-pipe and plenum. Details of
the distributions of coolant velocity, temperature, k and omega at the hole exits are also presented.
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Effect of Coolant Temperature
and Mass Flow on Film Cooling of Turbine Blades
Garg, Vijay K. (AYT Corp.,
Cleveland, OH United States); Gaugler, Raymond E.
(NASA Lewis Research Center, Cleveland, OH United States)
NASA Center for AeroSpace Information (CASI)
NASA-TM-112760 , 1997
A three-dimensional Navier Stokes code has been used
to study the effect of coolant temperature, and coolant to mainstream mass flow
ratio on the adiabatic effectiveness of a film-cooled turbine blade. The blade
chosen is the VKI rotor with six rows of cooling holes including three rows on
the shower head. The mainstream is akin to that under real engine conditions
with stagnation temperature = 1900 K and stagnation pressure = 3 MPa. Generally, the adiabatic effectiveness is lower for a
higher coolant temperature due to nonlinear effects via the compressibility of
air. However, over the suction side of shower-head holes, the effectiveness is
higher for a higher coolant temperature than that for a lower coolant
temperature when the coolant to mainstream mass flow ratio is 5% or more. For a
fixed coolant temperature, the effectiveness passes through a
minima on the suction side of shower-head holes as the coolant to
mainstream mass flow, ratio increases, while on the pressure side of
shower-head holes, the effectiveness decreases with increase in coolant mass
flow due to coolant jet lift-off. In all cases, the adiabatic effectiveness is
highly three-dimensional.
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Adiabatic Effectiveness and
Heat Transfer Coefficient on a Film-Cooled Rotating Blade
Garg, Vijay K. (AYT Corp.,
Cleveland, OH United States)
NASA Center for AeroSpace Information (CASI)
ASME Paper 96-GT-221 , 1997
three-dimensional Navier-Stokes code has been used to
compute the adiabatic effectiveness and heat transfer coefficient on a rotating
film-cooled turbine blade. The blade chosen is the United Technologies Research
Center(UTRC) rotor with five film-cooling rows
containing 83 holes, including three rows on the shower head with 49 holes,
covering about 86% of the blade span. The mainstream is akin to that under real
engine conditions with stagnation temperature 1900 K and stagnation pressure 3 MPa. The blade speed is taken to be 5200 rpm. The adiabatic
effectiveness is higher for a rotating blade as compared to that for a
stationary blade. Also, the direction of coolant injection from the shower-head
holes considerably affects the effectiveness and heat transfer coefficient
values on both the pressure and suction surfaces. In all cases the heat
transfer coefficient and adiabatic effectiveness are highly three-dimensional
in the vicinity of holes but tend to become two-dimensional far downstream.
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Effect of Velocity and
Temperature Distribution at the Hole Exit on Film Cooling of Turbine Blades
Garg, V. K. (AYT Corp.,
Cleveland, OH United States); Gaugler, R. E. (NASA Lewis
Research Center, Cleveland, OH United States)
NASA Center for AeroSpace Information (CASI)
NASA/TM-97-208146 , 1997
An existing three-dimensional Navier-Stokes code (Arnone et al, 1991), modified Turbine Branch, to include
film cooling considerations (Garg
and Gaugler, 1994), has been used to study the effect
of coolant velocity and temperature distribution at the hole exit on the heat
transfer coefficient on three film-cooled turbine blades, namely, the C3X vane,
the VKI rotor, and the ACE rotor. Results are also compared with the
experimental data for all the blades. Moreover, Mayle's
transition criterion (1991),
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Comparison of Two-Equation
Turbulence Models for Prediction of Heat Transfer on Film-Cooled Turbine Blades
Garg, Vijay K. (AYT Corp.,
Cleveland, OH United States); Ameri, Ali A. (AYT
Corp., Cleveland, OH United States)
NASA Center for AeroSpace Information (CASI)
ASME Paper 97-GT-024 , 1997
A three-dimensional Navier-Stokes code has been used
to compute the heat transfer coefficient on two film-cooled turbine blades,
namely, the VKI rotor with six rows of cooling holes, including three rows on
the shower head and the C3X vane with nine rows of holes, including five rows
on the shower head. Predictions of heat transfer coefficient at the blade
surface using three two-equation turbulence model specifically, Coakley's q-omega model, Chien's
k-epsilon model and Wilcox's k-omega model with Menter's
modifications, have been compared with the experimental data of Camci and Arts for the VKI rotor, and of Hylton et al. for the C3X vane along with predictions using
the Baldwin-Lomar (B-L) model taken from Garg and Gaugler.
It is found that for the cases considered here the two equation models predict
the blade heat transfer somewhat better than the B-L model except immediately
downstream of the film-cooled holes on the suction surface of the VKI rotor, and
over most of the suction surface of the C3X vane. However, all two-equation
models require 40% more computer core than the B-L model for solution, and
while the q-omega and k-epsilon models need 40% more computer time than the B-L
model the k-omega model requires at least 65% more time because of the slower
rate of convergence. It is found that the heat transfer coefficient exhibit a
strong spanwise as well as streamwise
variation for both blades and all turbulence models.
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Leading edge film cooling
effects on turbine blade heat transfer
Garg, Vijay K. (AYT Corp.,
Brook Park, OH., United States); Gaugler, Raymond E.
(NASA Lewis Research Center, Cleveland, OH, United States)
NASA Center for AeroSpace Information (CASI)
NASA-TM-106955 , 1995
An existing three dimensional Navier-Stokes code,
modified to include film cooling considerations, has been used to study the
effect of spanwise pitch of shower-head holes and
coolant to mainstream mass flow ratio on the adiabatic effectiveness and heat
transfer coefficient on a film-cooled turbine vane. The mainstream is akin to
that under real engine conditions with stagnation temperature = 1900 K and
stagnation pressure = 3 MPa. It is found that with
the coolant to mainstream mass flow ratio fixed, reducing P, the spanwise pitch for shower-head holes, from 7.5 d to 3.0 d,
where d is the hole diameter, increases the average
effectiveness considerably over the blade surface. However, when P/d= 7.5,
increasing the coolant mass flow increases the effectiveness on the pressure
surface but reduces it on the suction surface due to coolant jet lift-off. For
P/d = 4.5 or 3.0, such an anomaly does not occur within the range of coolant to
mainstream mass flow ratios analyzed. In all cases, adiabatic effectiveness and
heat transfer coefficient are highly three-dimensional.
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Effect of velocity and
temperature distribution at the hole exit on film cooling of turbine blades
Garg, Vijay K. (AYT Corp.,
Brook Park, OH., United States); Gaugler, Raymond E.
(NASA Lewis Research Center, Cleveland, OH, United States)
NASA Center for AeroSpace Information (CASI)
NASA-TM-106954 , 1995
An existing three-dimensional Navier-Stokes code,
modified to include film cooling considerations, has been used to study the
effect of coolant velocity and temperature distribution at the hole exit on the
heat transfer coefficient on three-film-cooled turbine blades, namely, the C3X
vane, the VKI rotor, and the ACE rotor. Results are also compared with the
experimental data for all the blades. Moreover, Mayle's
transition criterion,
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Prediction of film cooling on
gas turbine airfoils
Garg, Vijay K. (NASA Lewis
Research Center, Cleveland, OH, United States); Gaugler,
Raymond E. (NASA Lewis Research Center, Cleveland, OH, United States)
NASA Center for AeroSpace Information (CASI)
NASA-TM-106653 , 1994
A three-dimensional Navier-Stokes analysis tool has
been developed in order to study the effect of film cooling on the flow and
heat transfer characteristics of actual turbine airfoils. An existing code (Arnone et al., 1991) has been modified for the purpose. The
code is an explicit, multigrid, cell-centered, finite
volume code with an algebraic turbulence model. Eigenvalue
scaled artificial dissipation and variable-coefficient implicit residual
smoothing are used with a full-multigrid
technique. Moreover, Mayle's transition criterion (Mayle, 1991) is used. The effects of film cooling have been
incorporated into the code in the form of appropriate boundary conditions at
the hole locations on the airfoil surface. Each hole exit is represented by several control volumes, thus
providing an ability to study the effect of hole shape on the film-cooling
characteristics. Comparison is fair with near mid-span experimental data for
four and nine rows of cooling holes, five on the shower head, and two rows each
on the pressure and suction surfaces. The computations, however, show a strong spanwise variation of the heat transfer coefficient on the
airfoil surface, specially with shower-head cooling.
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