Perry L. Johnson

University of Central Florida, Orlando, Florida

Vikram Shyam

Glenn Research Center, Cleveland, Ohio

Large Eddy Simulation of a Film Cooling Flow

Injected From an Inclined Discrete Cylindrical

Hole Into a Crossfl ow With Zero-Pressure

Gradient Turbulent Boundary Layer

NASA/TM—2012-217695

August 2012
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Perry L. Johnson

University of Central Florida, Orlando, Florida

Vikram Shyam

Glenn Research Center, Cleveland, Ohio

Large Eddy Simulation of a Film Cooling Flow

Injected From an Inclined Discrete Cylindrical

Hole Into a Crossfl ow With Zero-Pressure

Gradient Turbulent Boundary Layer

NASA/TM—2012-217695

August 2012

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared under Contract NNC097A01C
Acknowledgments

The author is grateful for the mentoring support of Dr. Vikram Shyam through the Lewis Educational Research Collaborative 

Internship Program (LERCIP) with the Ohio Aerospace Institute (OAI). The author benefi ted from helpful discussions with

Drs. Shyam, Lamyaa El-Gabry, and Ali Ameri concerning the experimental data and numerical techniques.

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This work was sponsored by the Fundamental Aeronautics Program 

at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2012-217695 1 

Large Eddy Simulation of a Film Cooling Flow Injected From an 

Inclined Discrete Cylindrical Hole Into a Crossflow With ZeroPressure Gradient Turbulent Boundary Layer 

Perry L. Johnson 

University of Central Florida 

Orlando, Florida 32816 

Vikram Shyam 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

A Large Eddy Simulation (LES) is performed of a high blowing ratio (M = 1.7) film cooling flow 

with density ratio of unity. Mean results are compared with experimental data to show the degree of 

fidelity achieved in the simulation. While the trends in the LES prediction are a noticeable improvement 

over Reynolds-Averaged Navier-Stokes (RANS) predictions, there is still a lack a spreading on the 

underside of the lifted jet. This is likely due to the inability of the LES to capture the full range of 

influential eddies on the underside of the jet due to their smaller structure. The unsteady structures in the 

turbulent coolant jet are also explored and related to turbulent mixing characteristics. 

Introduction 

It is well known that film cooling is one of the useful techniques that help the metal components survive 

in the extremely harsh environment immediately downstream of the combustion chamber of a gas turbine. 

With improvement of film cooling design, this combustion temperature continues to increase over time. 

Higher temperatures of the combustion products yield higher efficiencies for the turbine. For film cooling, 

compressed air is drawn from the compressor stage and injected through film holes, which are machined in 

the metal surface of the blades and endwalls. These coolant jets then create a blanket-like effect to protect 

the solid surface from the hot main flow. Although the film cooling technique has been successfully 

implemented in the gas turbine industry, it is still in the process of both improvement and optimization. The 

designer always wants to achieve the best capability to protect the surface from the high mainstream 

temperature using least amount of coolant in order to achieve higher gas turbine cycle efficiency. 

The simplest discrete hole film cooling geometry is the cylindrical hole (Goldstein et al. 1968, 

Goldstein et al. 1970, Eriksen et al. 1974), which is typically inclined at an angle around 30° to 35° to the 

crossflow. The dominance of the detrimental counter-rotating vortex pair leads to diminished 

effectiveness on the downstream wall (Haven et al. 1997). While the streamwise cylindrical hole is the 

simplest film hole configuration, many alternate geometries have been shown to provide better 

effectiveness downstream of the hole. The most notable of these variations include: compound angle 

holes (Ligrani et al. 1994a-b, Sen et al. 1996, Schmidt et al. 1996), diffusion-shaped holes(Bunker 2005), 

and trenched holes (Bunker 2001, Bunker 2002). 

The focus of the present paper is to study the unsteady interactions of a compound angle film cooling 

hole using LES. Film cooling scenarios have typically been studied using the Reynolds-Averaged NavierStokes equations (RANS) with two-equation turbulence models (Walters and Leylek 2000, McGovern and 

Leylek 2000, Hyams and Leylek 2000, Brittingham and Leylek 2000). In particular, latter two study the 

time-averaged flow features of compound angle holes using the RANS numerical approach. RANS analysis 

is a steady flow approach, which yields time-averaged results by using various viscous models to account 

for the effects of turbulence. LES, however, directly resolves the more influential turbulent scales by solving 
NASA/TM—2012-217695 2 

the Navier-Stokes equations in both the spatial and temporal domains. In this way, LES captures the effects 

of the problem-dependent, high energy large scale velocity fluctuations in the studied flow field. Acharya et 

al. (2001) studied LES predictions of square hole geometries and concluded that the two equation turbulence 

models, used with the RANS equations, overpredict the vertical penetration and underpredict the lateral 

spreading of the coolant jet. They showed that LES solutions predicted mean velocities and turbulent 

stresses better than RANS models. Furthermore, Tyagi and Acharya (2003) used LES to study cylindrical 

inclined coolant holes more closely. Not only did LES enable them to identify the known vortical structures 

of a jet in crossflow (Fric and Roshko 1994, Kelso et al. 1996, Haven and Kurosaka 1997), but also led them 

to propose a unifying ‘hairpin’ vortex structure. They concluded that their LES results matched previous 

experimental results of Lavrich and Chiappetta (1990) and Sinha et al. (1991) with few discrepancies. 

Ultimately, they suggested that their observations of the LES solution showed a high significance of the 

shortcomings of the RANS models. 

More recent film cooling LES studies were performed by: Iourokina and Lele (2006a-b), Peet and 

Lele (2008), Leedom and Acharya (2008), Guo et al. (2006), and Renze et al. (2008a-c, 2009). The 

present study simulates a cylindrical film cooling hole with high blowing ratio than typically simulated, 

specifically, M = 1.7. This high of a blowing ratio, while not ideal, can still play a role in engine 

performance during aircraft cruise, since film cooling is typically designed to survive the tougher 

conditions at take-off. 

Numerical Methodology 

The open source software OpenFOAM was used for solving the filtered governing equations for 

mass, momentum, and energy conservation. OpenFOAM uses the finite volume method (FVM) of 

discretization on unstructured grids. 

Experimental Data for Validation 

Data from the experiments of El-Gabry (2011) for film cooling flow fields is used for validation of 

mean and second-order statistics. The numerical domain and boundary conditions are constructed so as to 

best approximate the experimental conditions. 

Main Simulation 

Domain and Boundary Conditions 

The numerical domain (Fig. 1) included one full pitch with one cooling hole in the center and 

periodic planes on either side. The surface boundary was an adiabatic wall satisfying the no-slip 

conditions. The grid was clustered near the wall such that the viscous sublayer was satisfactorily resolved 

(y+ 1). The height of the solution domain was 5 diameters, as suggested by Johnson et al. (2010). At the 

free-stream boundary, the streamwise velocity was specified and the other two velocity components were 

given zero gradients. All other variables were specified with zero gradients. The domain for the main 

simulation was chosen to include enough space upstream of the hole for the inclusion of the recycling 

technique to be merged with the main solution, which provided the velocity inlet conditions. Specifically, 

the domain extended 12 diameters upstream of the injection, with the recycling plane 

~5.5 diameters upstream of injection and ~6.5 diameters (~70 displacement thicknesses) downstream of 

the inlet plane. In this way, no auxiliary simulation was needed in parallel with the main simulation (only 

for initialization of the main solution). A constant temperature was specified at the cross-flow inlet. 

Downstream, the domain was extended 12 diameters downstream of the injection, far enough for 

comparison with relevant experimental data but not too far as to reach a point where the merging of the 

jets completely invalidates the physics of the periodic boundary conditions. The advective boundary 

condition using local fluid velocity was specified on the outflow plane. 
NASA/TM—2012-217695 3 

5D

Figure 1.—Extent of numerical domain. 

Recycling 

Plane

Inlet

Outlet

Free-stream

Wall

Wall

Inlet

Periodic

Wall

Figure 2.—Boundaries for full numerical domain. 

No coolant plenum was modeled and the coolant hole was supplied with “fully-developed” turbulent 

pipe flow, so as to best match experimental conditions in a time-resolved manner. The cylindrical coolant 

hole was an adiabatic, no-slip wall. The different boundaries are shown in Figure 2. The specifics of the 

boundary conditions are summarized in Table 1. 
NASA/TM—2012-217695 4 

TABLE 1.—BOUNDARY CONDITIONS FOR FULL NUMERICAL DOMAIN 

Boundary Velocity Pressure Temperature Other 

Inlet recycled(crossflow rescaled) Zero-gradient 297 K Zero-gradient 

Wall Ux=Uy=Uz=0 m/s Zero-gradient Adiabatic Zero 

Free-stream Ux=U∞, Uy,Uz: zero-gradient p = p∞ Zero-gradient Zero-gradient 

Outlet Advective Zero-gradient Zero-gradient Zero-gradient 

Figure 3.—Computational grid.

Grid and Spatial Discretization 

The domain was discretized using the finite volume method in OpenFOAM. The numerical grids 

were generated using the GridPro software available commercially from Product Development Company 

(PDC). The grids were generated in a block-structured manner but converted to unstructured format for 

use with OpenFOAM. The final grid consisted of about 9 million hexahedral cells. Near the no-slip walls, 

the grid was refined such that the first cell fell within y+  1. For the boundary layer simulation, the 

streamwise spacing was x+ ~ 35, and the spanwise spacing was z+ ~ 16. These spacings were continued 

throughout the domain with the ratio of the boundary layer thickness to coolant hole diameter being 

~0.75. The main simulation grid is shown in Figure 3. 

All diffusive Laplacian terms were discretized using a second-order centered scheme, consistent with 

the nature of diffusion. The convective terms were likewise discretized with second-order central 

schemes, but limiters were introduced to maintain solution stability and boundedness. 
NASA/TM—2012-217695 5 

Solution Technique and Temporal Discretization 

The incompressible Navier-Stokes equations were solved along with the continuity and energy 

equations. A low-pass filtering operation in three-dimensional space was implicitly performed by the grid. 

Residual stresses not resolved by the grid were modeled using the dynamic procedure of Germano (1991), 

specifically with the modification of Lilly (1992). The solution was progressed in time such that the 

Courant number was kept well below 0.5 in regions of high interest. This allowed for the time-accurate 

resolution of the non-filtered scales. The second-order fully-implicit backwards-differencing method was 

used to discretize all temporal derivatives. A PISO-SIMPLE scheme was used for time advancement. The 

SIMPLE iteration technique was used to converge the solution at each time step, with multiple pressure 

corrections in each SIMPLE iteration (resembling the PISO method). Typical time steps employed three 

SIMPLE (“outer”) iterations and two pressure corrections (“inner iterations”). The initial conditions for 

the simulation were developed by converging the full domain using a steady solver (SIMPLE) with the 

RANS equations and k- SST turbulence model. Then, the results of two precursor simulations were 

mapped onto the domain and the simulation was started. 

Precursor Simulations 

In order to hasten the development of a fully turbulent boundary layer within the recycling domain 

upstream of the film injection, the boundary layer simulation was started on a considerably smaller 

domain. The goal was merely to initialize the corresponding portion of the full simulations with a realistic 

turbulent boundary layer. Similarly, the upstream portion within the coolant tube was also developed first 

in a preliminary simulation before used to initialize the instantaneous velocities as fully developed pipe 

flow. The numerical grids for these precursor simulations were extracted from the numerical grid from the 

full simulation such that there was direct mapping to transfer the results of the precursor simulation to the 

initialized full simulation. 

Similarly, the first five diameters of the coolant pipe was extracted from the grid and simulated 

separately. This also was mapped directly onto the simulation initialized flow field. 

Domain and Boundary Conditions 

The extent of the boundary layer was 10 disturbance thicknesses in length (~7.5 diameters), 2.67 

boundary layer thicknesses normal to the wall (~2 diameters), and the full 3 diameters in the lateral 

direction with periodic boundary conditions. The grid corresponding to this region upstream of the hole 

was extracted from the full grid. Figure 4 shows the region extracted from the full domain. 

The inlet to the precursor domain followed the specification detailed in Appendix A for generating an 

inflowing spatially developing turbulent boundary layer. The recycle plane was placed 8.75 disturbance 

thicknesses (~70 displacement thicknesses) downstream of the inlet to avoid spurious accumulation of 

error, false periodicity, and unwanted interaction of streamwise-oriented coherent structures (see Jewkes 

et al. 2011). 

The boundary conditions for the sub-domains were specified to reflect the boundary conditions of the 

main solution as closely as possible. The boundary layer precursor domain inlet, wall and periodic 

conditions remained identical to the full domain. The free-stream boundary conditions were the same, but 

moved closer to the wall to reduce the computational cost of the precursor simulation. The outflow, 

likewise, was an advective outflow condition, as in the main simulation, but was moved necessarily to 

accommodate the boundary layer simulation. An advective outflow was added to the pipe flow domain, as 

necessary to close the domain. All other boundary conditions remained identical to the full simulation for 

the pipe flow precursor. 
NASA/TM—2012-217695 6 

Figure 4.—Two precursor simulation grid extracted from full domain grid. 

Grid and Spatial Discretization 

As previously mentioned, the grid was created for the precursor simulation simply by extracting the 

relevant domain from the full solution domain grid. Spatial discretization was the same as applied in the 

main simulation. 

Solution Technique and Temporal Discretization 

The solution method was similar to the full simulation. Some differences were applied due to the fact 

that the grid was very well-behaved compared to some regions of the main simulation near the lip of the 

coolant hole. The PISO method was applied with no outer iterations at each time step. Three pressure 

corrections were used within the PISO method to converge the solution at each time step. The semiimplicit Crank-Nicholson scheme was used for temporal discretization precursor simulations. Note that 

the backward scheme in time was used for the main simulation, due to stability concerns. The simple 

boundary layer and pipe flow simulations had no such stability problems. 

Initialization and Precursor Results 

The initial conditions for these simulations are discussed in Appendix B. Figure 5 shows the velocity 

magnitude at the initialized conditions. The results of the precursor simulations are subsequently shown in 

Figure 6. The difference is visually perceptible, with the second figure having obtain realistic turbulent 

structure, whereas the initialization is simply a superposition of random perturbations. 

These two results are then mapped onto the full domain simulation, after it has been initialized by 

converging the RANS equations with k-ω SST turbulent model. Figure 7 shows the final initialized full 

domain. 
NASA/TM—2012-217695 7 

Figure 5.—Initial velocity magnitude contours for precursor simulations (see Appendix B). 

Figure 6.—Final velocity magnitude contours for precursor simulations. 

Boundary Layer

Pipe Flow

Converged

RANS

Figure 7.—Initialized velocity for the full domain, with two piecewise mappings superposed on a 

converged RANS solution.
NASA/TM—2012-217695 8 

Results 

The simulation was advanced with a time step of 5e-6 sec (~0.0025 D/U∞). From the initial 

conditions, the simulation was run for 0.175 sec (~90 D/U∞) before beginning statistical sampling. 

Statistical samples were taken over a period of 0.3 sec (~150 D/U∞). 

Mean Profiles 

The evolution of the jet is shown in Figures 8 to 12 by non-dimensional temperature contour plots at 

constant x locations. Comparison with experimental data and RANS solution is shown at each location. 

The temperature contours best mark the presence of the cooling jet and are of highest interest in the film 

cooling scenario, since the goal is to provide thermal protection to the adiabatic surface. The streamwise 

locations are designated by distance downstream of the leading edge of the film hole. 

Figure 8 shows the temperature contours at x/d = 2. This is the streamwise location of the trailing 

edge of the hole. For this reason, the core of the jet remains intact and has only just begun to spread. The 

experimental results show slightly more initial spreading than either of the numerical results. Most 

notable is the proximity of the outer part of the coolant to the wall. The experiment shows that the coolant 

is much closer to the wall. 

At x/d = 3, Figure 9, the coolant shows noticeable lift-off behavior. The coolant core is also bent into 

an upside-down “U-shape”. This indicates the presence of a strong counter-rotating vortex in the wake of 

the jet. The numerical simulations still underpredict the mixing of the coolant, as the coolant core seems 

to remain more intact than in the experiment. Concerning trajectory, the main difference again is the 

proximity of the under-side of the jet to the surface. The trajectory of the upper-half of the jet is 

accurately predicted. Both RANS and LES predictions have trouble matching the experimental data at 

this point. 

The LES begins to show better predictions with the experimental data at the jet progresses to x/d = 4, 

Figure 10. Certainly, the LES predictions are better than that of the RANS model, which continues is 

under-prediction of jet’s mixing. The LES predicts a more defined upside-down “U” shape than the 

experiment and still does not show enough spreading back to the wall. Otherwise, the contours are largely 

in agreement with the experimental data. 

RANS k-ω SST LES Experiment

Figure 8.—Comparison of mean non-dimensional temperature contours at x/d = 2 (from the leading edge of the 

injection). 
NASA/TM—2012-217695 9 

RANS k-ω SST LES Experiment

Figure 9.—Comparison of mean non-dimensional temperature contours at x/d = 3. 

RANS k-ω SST LES Experiment

Figure 10.—Comparison of mean non-dimensional temperature contours at x/d = 4. 
NASA/TM—2012-217695 10 

At x/d = 6, the LES continues to show dramatically better agreement with experiment than the RANS 

model, Figure 11. The LES shows a slight general under-prediction of spreading in all directions, but this 

is still most notable on the underside of the jet. The “U” shape of the jet is now mostly blurred by the 

increased turbulent mixing within the jet and wake regions. The core region of the jet shows temperatures 

mid-way between those of the coolant and crossflow. 

The final streamwise station to be shown is x/d = 8, Figure 12. At this point, it is seen from the 

experiment that neighboring jets have begun to interact. This places the periodic boundary conditions of 

the LES in suspect, due to the possibility of enforcing unphysical periodicity. The temperature contours 

from the experiment show that the spreading of the jet has greatly increases since x/d = 6. The jump in 

spreading is not mirrored in the LES, though the general shape and position of the jet is well predicted. 

The RANS model clearly fails to predict the experiment, even qualitatively. The failure of the LES to 

keep pace with the experiment in the spreading of the jet is attributable to the spanwise boundary 

conditions. At this location, the adjacent jets have begun to merge and periodic boundary conditions 

enforce a false periodicity in the time-accurate sense. The experiment consisted of a row of just three film 

cooling jets, while the current simulation attempts to model an infinite row of jets. Therefore, once the 

effects of neighboring jets begin to play a noticeable role, the fidelity of the spanwise boundary 

conditions in matching the experiment is highly questionable. 

In all of the streamwise-normal planes, an over-prediction of temperature at the wall is evident. This 

seems to be caused by the LES model failing to spread back toward the wall at a rate equal to the 

experiment. A possible explanation is the near wall treatment of the LES. Near solid boundaries, the 

range of turbulent length scales is shortened due to proximity to impermeability. This puts the theory 

behind LES on weak grounds. Many attempts, such as DES and two-layer zonal methods, have been 

made to get around the deficiency of LES near a wall, but all are still under development. No clear 

solution to the LES near-wall problem has emerged, other than essentially running DNS in the near-wall 

region. The failure of the current model to resolve enough “large eddies” in the near-wall region is very 

likely the cause of its failure to properly predict spreading back toward the wall. Since the ‘large eddies’ 

in the near-wall region are much smaller than elsewhere in the flow, even with the resolution in the wallnormal direction, the other two dimensions could be filtering out eddies of size and influence crucial to 

the turbulent mixing in this region. The anisotropy, meanwhile could be preventing the dynamic 

Smagorinsky model (based on the cube root of the cell volume), from accurately predicting the 

unresolved turbulent stresses. 

Figure 13 shows the temperature contours along the center-plane of the jet. The main difference is 

again on the underside of the jet, where the temperature is over-predicted (under-prediction in 

effectiveness). The turbulent mixing is evidently not as strong compared to the lifting force of the 

counter-rotating vortices in the LES model as compared with experiment. Similarly, x-velocity contours 

are shown in Figure 14. The maximum velocity in the jet core near the exit of the hole is noticeably 

different between the LES and the experiment. This points to the experiment’s failure to adhere to the 

nominal conditions, hence a mismatch in blowing ratio between the LES and experiment. Also, the 

boundary layer is seen to be much thicker in the LES than in the experiment. Other evidence not shown 

indicates that the recycling boundary layer method upstream of the hole was thickened beyond the desired 

thickness due to the presence of the hole and this effect was recycled to the inlet. This is clearly 

undesirable, as a thicker boundary layer would provide less downward forcing on the jet. 

In theory, these two mismatches between experiment and simulation have opposite effects. The lower 

jet velocity in the LES would tend to give it less lift-off and a lower trajectory. Meanwhile, the thicker 

boundary layer would raise the “effective blowing ratio”. The plots shown so far indicate that the mean 

trajectory was accurately predicted by the LES, indicating that these two effects approximately canceled. 

The effect of these two considerations on the mixing and spreading of the jet is less certain and an 

additional consideration when interpreting the differences between experiment and simulation. 
NASA/TM—2012-217695 11 

RANS k-ω SST LES Experiment

Figure 11.—Comparison of mean non-dimensional temperature contours at x/d = 6. 

RANS k-ω SST LES Experiment

Figure 12.—Comparison of mean non-dimensional temperature contours at x/d = 8. 
NASA/TM—2012-217695 12 

(a) 

(b) 

Figure 13.—(a) LES and (b) experimental centerline mean temperature contours. 

(a) 

Figure 14.—(a) LES and (b) experimental centerline x-velocity contours. 
NASA/TM—2012-217695 13 

Unsteady Dynamics 

One of the main advantages of the LES approach to simulation over RANS models in the ability to 

look into the dynamics of the large scale turbulence. Such eddies and structures carry the most influence 

over the mean statistics, and hence give insight into the cause behind flow features. 

Figure 15 shows iso-surfaces of temperature at a time in the flow from different viewing angles. The 

upper right shows an isometric view, revealing that the breakdown of the jet into unsteady shear layer 

roller vortices does not happen immediately at the exit of the hole, but after a little while of interaction 

between the jet and crossflow. This is echoed in the other two views. The upper right view, from the side 

of the jet, shows the general detached nature of the temperature contours (hence effectiveness) of the 

coolant. While the coolant appears to spread rapidly in the upper half of the jet, the spreading back toward 

the surface is subdued. No large structures (from a temperature point of view) are seen in the lower edge 

of the jet. This reflects that the eddy length scales are damped near the wall and spreading is reduced. The 

bottom image in Figure 15, looking down on the jet from above, shows the spanwise spreading. The 

spreading seems to be hindered once the jet starts reaching the edge of the domain, despite the periodic 

boundary. The hypothesis from above, invoking a false-periodicity in the spanwise direction, could be 

cited for this phenomenon and the under-prediction of spreading in this region. Shear layer roller vortices, 

caused by the interaction of the jet with the slower crossflow, are evident on the top and sides of the jet. 

The presence of an impermeable wall limits their presence on the underside of the jet. These structures are 

likely liable for the spreading rate of the jet. 

Another phenomenon uncovered in the instantaneous results is the occurrences of unsteady 

reattachment events on the side regions of the jet. Such an event is shown via temperature and shear stress 

contours at the wall, Figure 16. Figure 17 shows iso-surfaces of temperature from a view underneath the 

jet at the same time. It is seen that a structure of lower temperature, protruding from the jet, has made 

significant contact with the surface. 

Figure 15.—Iso-surfaces of temperature for the spreading jet. 
NASA/TM—2012-217695 14 

Figure 16.—Contours of temperature and skin friction on the adiabatic surface, showing a 

re-attachment event on the side of the jet. 

Figure 17.—View from underneath the jet of temperature is-surfaces, showing the 

re-attachment event. 
NASA/TM—2012-217695 15 

wake

jet

Figure 18.—Iso-surfaces of the lambda-2 criterion for vortex eduction from a top and side view of the jet. 

Figure 18 shows iso-surfaces of the lambda-2 criterion for vortex eduction (Jeong and Hussain 1995). 

The turbulent flow field is obviously very chaotic with a large range of resolved scales. The largest 

coherent structures are of the most interest, since they are likely most responsible for turbulent mixing 

and the spreading of the jet. The arrows attempt to link the location of shear layer structures in the two 

views. The top image, looking down on the jet, reveals spanwise coherent structures riding along the top 

of the jet. The side image reveals that the large structures are oriented normal to the wall. In fact, the two 

structures pointed with arrows in Figure 18 show the characteristics of both of these orientations. This 

evidence reveals an upside-down “U” shape pattern in the large-scale coherent structures. These are 

linked with the shear layer instability and roll-up seen in the temperature contours. It is interesting to note 

that these do not occur on the underside of the jet. Instead, smaller coherent structures are found. Indeed, 

there is a distinct boundary in the presence of these large structures in the upper half of the jet and 

absence in the lower half (wake region). 

Although unable to be visualized in this paper, motion pictures of the structures in Figure 18 reveal a 

second item of interest. Some structures are pulled, evidently by the motion of the counter-rotating vortex 

pair, into the wake region along the sides of the jet, where they dissipate on a collision course with the 

wall. These are in the same region as the unsteady reattachment events noted in the temperature isosurfaces. Thus, the evidence points to an unsteady separation of large scale shear rollups along the side of 

the jet from the “U”-shape and subsequent motion toward the wall, creating unsteady re-attachment 

events of increased adiabatic effectiveness and wall shear stress and higher speed fluid is force in close 

proximity to the wall. 
NASA/TM—2012-217695 16 

29

Approx. hole location

Shear Layer Wake

y=0.5D

Figure 19.—Y-vorticity at 0.5 diameters above the surface. 

A final note from the unsteady structure of the LES flow is to be made concerning the wall normal 

vorticity, Figure 19. As the jet emerges from the hole, the leading edge (which becomes the upper edge) 

of the jet is at a higher velocity than the crossflow. This causes vorticity wrapping away from the center 

of the jet in the shear layer at the top and side of the jet. Meanwhile, wake vortices typically take the form 

similar to those seen in the wake of a solid cylinder in crossflow. Such wake vorticity is opposite that of 

the shear layer. Such a pattern is indeed revealed in Figure 19. The figure shows a plane one-half diameter 

above the surface. In the streamwise direction, the shear layer y-vorticity becomes visible first, as the 

upper half of the jet passes through the plane. As the lower half passes and the wake reaches into the 

plane, the vorticity toward the centerline shows a pattern opposite that of the shear layer. 

Conclusion 

A wall-resolved large-eddy simulation (WR-LES) of a film cooling flow with hig blowing ratio 

(M = 1.7) was completed to complement the experimental study of El-Gabry (2010). Mean velocity and 

temperature profiles show good agreement with experimental data, far superior to predictions by the k-ω

SST model. Particularly, the spanwise spreading at the sides of the jet and the wall-normal spreading at 

the upper shear layer of the jet were well-predicted by the LES in the near-hole region. In the region 

where the jets begin to merge, the LES prediction of spreading does not keep up with the experimental 

data due to the growing infidelity of the spanwise periodic boundary condition. The region below the jet 

suffers the most disagreement between the LES predictions and experimental results. It is speculated that 

the resolution of the grid (size and anisotropy of the implicit filter) was not sufficient for the applicability 

of a zero-equation sub-grid model in a region of such fine-scale structures near the wall and in the region 

of high turbulent kinetic energy below the jet. Analysis of the unsteady structures reveal an upside-down 

“U” shaped tendency around the three free shear layers of the jet. The counter-rotating pair of vortices act 

in an apparently random manner to pull some of these structures around the side of the jet and toward the 

surface to create reattachment events on either side of the centerline. Wall-normal wake structures 

underneath the jet are shown similar to those of Fric and Roshko (1994), with opposite rotation to the 

shear layer roller vortices on the outer sides of the jet. 
NASA/TM—2012-217695 17 

Appendix A—Turbulent Boundary Layer Inflow 

A common difficulty for LES simulations of engineering wall bounded flows is the difficulty of 

specifying time-accurate boundary conditions for an approaching turbulent boundary layer. One popular 

family of techniques is the recycling-rescaling method first introduced by Lund et al. (1994). This method 

has been updated over the years and many variants have been introduced in open literature, most recently 

by Jewkes et al. (2011). In these methods, an auxiliary simulation is performed alongside the main 

simulation in which a plane of the velocity field is extracted and scaled using a blend of inner and outer 

equilibrium boundary layer scaling. The scaled velocity data is reintroduced at the inlet of the auxiliary 

simulation and given enough time, a realistic spatially-developing turbulent boundary layer is produced. 

At this point, velocity planes are extracted from the auxiliary domain and introduced at the inlet of the 

main simulation. 

In the present implementation, the method of Lund et al. (1994) is adopted with some of the 

modifications of Jewkes et al. (2011). Specifically, the displacement thickness is used instead of the 

disturbance thickness and the initialization suggestions of Jewkes et al. (2011) are followed. In addition, a 

modification suggested but not implemented by Lund et al. (1994) is used. This modification is the 

elimination of the need for the auxiliary solution. Instead, the recycling plane is simply taken from the main 

simulation itself, scaled, and reintroduced at the inlet. In this way, the need for an auxiliary domain is 

eliminated. To decrease the cost of the start-up transients in the recycling method, however, a smaller piece 

upstream of the film cooling hole is extracted from the main grid and the solution is first advanced on this 

smaller grid until a realistic turbulent boundary layer is produced (Figs. 4 and 6). The velocity field from 

this smaller simulation is then mapped onto the full grid to initialize the main simulation (Fig. 7). For 

mathematical details of the method, the reader is referred to the two sources mentioned above. 

NASA/TM—2012-217695 19 

Appendix B—Turbulent Pipe Inflow 

The other flow inlet did not present as much of a challenge because a spatially developing inlet 

condition was not deemed necessary. The experimental holes were very long ( 15 D) so the goal of the 

inlet to the holes was to produce fully developed pipe flow. While the experimental holes may not have 

been perfectly fully developed in 15 diameters (an elbow and t-joint were upstream of the holes), for 

unsteady simulation purposes, the most reliable and accurate method of reproducing the experiment was 

the specification of fully developed flow at uniform temperature and adiabatic walls. 

The fully developed pipe flow was realized numerically by inserting a recycling plane 5 D 

downstream of the pipe inlet. No re-scaling was necessary because spatially developing flow was not 

desired. The velocity profiles were simply recycled to the inlet at each time step. This reliably produced 

the desired effect with the initialization procedure below. The recycling place was 5 D upstream of the 

hole exit. 

NASA/TM—2012-217695 21 

Appendix C—Initialization Methods 

Initialization for the upstream boundary layer and coolant pipe inlet recycling sub-domains was 

accomplished through a procedure similar to that described by Jewkes et al. (2011). The mean flow given 

by the Spalding law was first applied to the boundary layer thickness and pipe radius. Then random 

fluctuations were superposed on the mean profile. The random fluctuations were applied in a piece-wise 

fashion as suggested by Jewkes. Gaussian distributions with standard deviations based on percent of the 

local velocity were applied and clipped at two standard deviations from the mean. The resulting velocity 

values were also clipped at the free-stream velocity. The piecewise initialization was similar to those 

suggested by Jewkes: from 0.05  y/(delta)  0.25, standard deviations of 40% for u', 25% for v', and 30% 

for w', from 0.25  y/(delta)  0.5, standard deviations of 20% for u', 12.5% for v', and 15% for w', and 

from 0.5  y/(delta)  1.0, standard deviations of 10% for u', 6.25% for v', and 7.5% for w'. Figure 5 

shows the results of this initialization for both sub-domains. 
NASA/TM—2012-217695 22 

References 

Acharya, S., Tyagi, M., and Hoda, A., 2001, “Flow and Heat Transfer Predictions for Film Cooling,” 

Heat Transfer in Gas Turbine Systems, Ann. N.Y. Acad. Sci., Vol. 934, pp. 110–125. 

Brittingham, R.A., and Leylek, J.H., 2000, “A Detailed Analysis of Film-Cooling Physics: Part IV – 

Compound-Angle Injection with Shaped Holes,” ASME Journal of Turbomachinery, Vol. 122, 

pp. 102-112. 

Bunker, R.S., 2005, “A Review of Shaped Hole Turbine Film-Cooling Technology,” Journal of Heat 

Transfer, 127, pp. 441-453. 

Bunker, R.S., 2001, “A Method for Improving the Cooling Effectiveness of a Gaseous Coolant Stream,” 

US Patent 6,234,755. 

Bunker, R.S., 2002, “Film Cooling Effectiveness due to Discrete Holes within a Transverse Trench,” 

Proceedings of the ASME Turbo Expo 2002, Amsterdam, The Netherlands, GT-2002-30178. 

El Gabry, L., 2011, “Turbulence and Heat Transfer Measurements in an Inclined Large Scale Film 

Cooling Array – Part I, Velocity and Turbulence Measurements,” ASME Turbo Expo, GT2011-

46491. 

El Gabry, L., 2011, “Turbulence and Heat Transfer Measurements in an Inclined Large Scale Film 

Cooling Array – Part II, Temperature and Heat Transfer Measurements,” ASME Turbo Expo, 

GT2011-46498. 

Eriksen, V.L., and Goldstein, R.J., 1974, “Heat Transfer and Film Cooling Following Injection Through 

Inclined Circular Tubes,” ASME Journal of Heat Transfer, Vol. 96, pp. 239-245. 

Fric, T. F., and Roshko, A., 1994, “Vortical Structure in the Wake of a Transverse Jet,” J. Fluid Mech., 

Vol. 279, pp. 1–47. 

Germano, M., Piomelli, U., Moin, P., and Cabot, W.H., 1991, “A Dynamic Subgrid-Scale Eddy Viscosity 

Model,” Phys. Fluids, Vol. 3, No. 7, pp. 1760-1765. 

Goldstein, R.J., Eckert, E.R.G., and Ramsey, J.W., 1968, “Film Cooling with Injection Through Holes: 

Adiabatic Wall Temperatures Downstream of a Circular Hole,” ASME Journal of Engineering for 

Power, Vol. 90, pp. 384-395. 

Goldstein, R.J., Eckert, E.R.G., Eriksen, V.L., and Ramsey, J.W., 1970, “Film Cooling Following 

Injection Through Inclined Circular Tubes,” Israel Journal of Technology, Vol. 8, pp. 145-154. 

GridPro. Product Development Company, www.gridpro.com 

Guo, X., Schröder, W., and Meinke, M, 2006, “Large-eddy Simulations of Film Cooling Flows,” 

Computers and Fluids, Vol. 35, pp. 587–606. 

Haven, B.A., and Kurosaka, M., 1997, “Kidney and Anti-Kidney Vortices in Crossflow Jets,” J. Fluid 

Mech., Vol. 352, pp. 27–64. 

Haven, B.A., Yamagata, D.K., Kurosaka, M., Yamawaki, S., and Maya, T., 1997, “Anti-Kidney Pair of 

Vortices in Shaped Holes and Their Influence on Film Cooling Effectiveness,” Proceedings of ASME 

Turbo Expo 1997, Orlando, FL, USA, 97-GT-45. 

Hyams, D.G., and Leylek, J.H., 2000, “A Detailed Analysis of Film-Cooling Physics: Part III – 

Streamwise Injection with Shaped Holes,” ASME Journal of Turbomachinery, Vol. 122, pp. 122-132. 

Iourokina, I.V., and Lele, S.K., 2006a, “Large Eddy Simulation of Film-Cooling Above the Flat Surface 

with a Large Plenum and Short Exit Holes,” Proceedings of the AIAA Aerospace Sciences Meeting 

and Exhibit, 2006, Reno, Nevada, USA. 

Iourokina, I.V., and Lele, S.K., 2006b, “Large Eddy Simulation of Film Cooling Flow Above a Flat 

Surface from Inclined Cylindrical Holes,” Proceedings of the ASME Joint U.S.-European Fluids 

Engineering Summer Meeting, 2006, Miami, Florida, USA. 

Jeong, J., and Hussain, F., 1995, “On the Identification of a Vortex,” J. Fluid Mech., vol. 285, pp. 69-94. 

Jewkes, J.W., Chung, Y.M., and Carpenter, P.W., 2011, “Modification to a Turbulent Inflow Generation 

Method for Boundary-Layer Flows,” AIAA J. Technical Note, Vol. 49, No. 1, pp. 247-250. 
NASA/TM—2012-217695 23 

Johnson, P.L., Nguyen, C.Q., Ho, S.H., and Kapat, J.S., 2010, “Sensitivity Analysis of Domain 

Considerations for Numerical Simulations of Film Cooling,” Proceedings of the 14th ASME 

International Heat Transfer Conference, IHTC14-23241. 

Johnson, P.L., Shyam, V., Hah, C., 2011, “Reynolds-Averaged Navier-Stokes Solutions to Flat Plate Film 

Cooling Scenarios,” NASA/TM—2011-217025. 

Kelso, R.M., Lim, T.T., and Perry, A.E., 1996, “An Experimental Study of Round Jets in Cross-flow,” 

Journal of Fluid Mechanics, Vol. 306, pp. 111-144. 

Lavrich, P.L., and Chiappetta, L.M., 1990, ‘‘An Investigation of Jet in a Cross Flow for Turbine Film 

Cooling Applications,” United Technologies Research Center, UTRC Report No. 90-04. 

Leedom, D.H., and Acharya, S., 2008, “Large Eddy Simulations of Film Cooling Flow Fields From 

Cylindrical and Shaped Holes,” Proceedings of ASME Turbo Expo 2008, Berlin, Germany, GT2008-

51009. 

Ligrani, P.M., Wigle, J.M., Ciriello, S. and Jackson, S.W., 1994a, “Film-Cooling from Holes with 

Compound Angle Orientations: Part I – Results Downstream of Two Staggered Rows of Holes with 

3d Spanwise Spacing,” ASME Journal of Heat Transfer, Vol. 116, pp. 353-362. 

Ligrani, P.M., Wigle, J.M., and Jackson, S.W., 1994b, “Film-Cooling from Holes with Compound Angle 

Orientations: Part II – Results Downstream of a Single Row of Holes with 6d Spanwise Spacing,” 

ASME Journal of Heat Transfer, Vol. 116, pp. 353-362. 

Lilly, D.K., 1992, “A Proposed Modification to the Germano Subgrid-Scale Closure Method,” Phys. 

Fluids, Vol. 4, No. 3, pp. 633-635. 

Lund, T.S., Wu, X., and Squires, K.D., 1998, “Generation of Turbulent Inflow Data for Spatially 

Developing Boundary Layer Simulations,” J. Comp. Phys., Vol. 140, pp. 233-258. 

McGovern, K.T., and Leylek, J.H., 2000, “A Detailed Analysis of Film-Cooling Physics: Part II – 

Compound-Angle Injection with Cylindrical Holes,” ASME Journal of Turbomachinery, Vol. 122, 

pp. 113-121. 

OpenFOAM®, www.openfoam.com 

Peet, Y.V., and Lele, S.K., 2008, “Near Field of Film Cooling Jet Issued Into a Flat Plate Boundary 

Layer: LES Study,” GT2008-50420, Proceedings of ASME Turbo Expo 2008, Berlin, Germany. 

Renze, P., Schroder, W., Meinke, M., 2008a, “Large-Eddy Simulation of Film Cooling Flows at Density 

Gradients,” International Journal of Heat and Fluid Flow, Vol. 29, pp. 18-34. 

Renze, P., Schroder, W., Meinke, M., 2008b, “Large-Eddy Simulation of Film Cooling Flows with 

Variable Density Jets,” Flow Turbulence and Combustion, Vol. 80, 119-132. 

Renze, P., Schroder, W., and Meinke, M., 2008c, “Large-Eddy Simulation of Film Cooling Flow Ejected 

in a Shallow Cavity,” Proceedings of the ASME Turbo Expo 2008, Berlin, Germany, GT2008-50120. 

Renze, P., Schroder, W., and Meinke, M., 2009, “Large-Eddy Simulation of Interacting Film Cooling 

Jets,” Proceedings of the ASME Turbo Expo 2009, Orlando, Florida, USA, BT2009-59164. 

Schmidt, D.L., Sen, B., and Bogard, D.G., 1996, “Film Cooling with Compound Angle Holes – Adiabatic 

Effectiveness,” ASME Journal of Turbomachinery, Vol. 118, 807-813. 

Sen, B., Schmidt, D.L., and Bogard, D.G., 1996, “Film Cooling with Compound Angle Holes – Heat 

Transfer,” ASME Journal of Turbomachinery, Vol. 118, pp. 800-806. 

Sinha, A.K., Bogard, D.G., and Crawford, M.E., 1991, “Film-Cooling Effectiveness Downstream of a 

Single Row of Holes with Variable Density Ratio,” ASME Journal of Turbomachinery, Vol. 113, 

pp. 442-449. 

Tyagi, M., and Acharya, S., 2003, “Large Eddy Simulation of Film Cooling Flow from an Inclined 

Cylindrical Jet,” J Fluid Mech, Vol. 125, pp 734-742. 

Walters, D.K., and Leylek, J.H., 2000, “A Detailed Analysis of Film-Cooling Physics: Part I – 

Streamwise Injection with Cylindrical Holes,” ASME Journal of Turbomachinery, Vol. 122, 

pp. 102-112. 
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Large Eddy Simulation of a Film Cooling Flow Injected From an Inclined Discrete 

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14. ABSTRACT

A Large Eddy Simulation (LES) is performed of a high blowing ratio (M = 1.7) film cooling flow with density ratio of unity. Mean results 

are compared with experimental data to show the degree of fidelity achieved in the simulation. While the trends in the LES prediction are a 

noticeable improvement over Reynolds-Averaged Navier-Stokes (RANS) predictions, there is still a lack a spreading on the underside of the 

lifted jet. This is likely due to the inability of the LES to capture the full range of influential eddies on the underside of the jet due to their 

smaller structure. The unsteady structures in the turbulent coolant jet are also explored and related to turbulent mixing characteristics.

15. SUBJECT TERMS

Film cooling; Turbulence model; Heat transfer 

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