Vikram Shyam

Glenn Research Center, Cleveland, Ohio

Douglas Thurman

U.S. Army Research Laboratory, Glenn Research Center, Cleveland, Ohio

Phillip Poinsatte

Glenn Research Center, Cleveland, Ohio

Ali Ameri

The Ohio State University, Columbus, Ohio

Peter Eichele

Gilcrest Electric & Supply Company, Cleveland, Ohio

James Knight

Glenn Research Center, Cleveland, Ohio

Long Hole Film Cooling Dataset for

CFD Development 

Part 1: Infrared Thermography and Thermocouple Surveys

NASA/TM—2013-218086/PART1

November 2013
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Vikram Shyam

Glenn Research Center, Cleveland, Ohio

Douglas Thurman

U.S. Army Research Laboratory, Glenn Research Center, Cleveland, Ohio

Phillip Poinsatte

Glenn Research Center, Cleveland, Ohio

Ali Ameri

The Ohio State University, Columbus, Ohio

Peter Eichele

Gilcrest Electric & Supply Company, Cleveland, Ohio

James Knight

Glenn Research Center, Cleveland, Ohio

Long Hole Film Cooling Dataset for

CFD Development 

Part 1: Infrared Thermography and Thermocouple Surveys

NASA/TM—2013-218086/PART1

November 2013

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
Acknowledgments

The authors would like to acknowledge the support of the NASA Fundamental Aeronautics program’s Fixed Wing project under 

whose funding this work was performed. The authors also thank Dr. David Rigby (Vantage Partners) for his expert advice on 

grid generation.

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Level of Review: This material has been technically reviewed by technical management. 

This work was sponsored by the Fundamental Aeronautics Program 

at the NASA Glenn Research Center.
NASA/TM—2013-218086/PART1 1 

Long Hole Film Cooling Dataset for CFD Development

Part 1: Infrared Thermography and Thermocouple Surveys

Vikram Shyam

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135 

Douglas Thurman

U.S. Army Research Laboratory

Glenn Research Center

Cleveland, Ohio 44135 

Phillip Poinsatte

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135 

Ali Ameri

The Ohio State University

Columbus, Ohio 43210 

Peter Eichele

Gilcrest Electric & Supply Company 

Cleveland, Ohio 44135 

James Knight

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135 

Abstract

An experiment investigating flow and heat transfer of long (length to diameter ratio of 18) cylindrical 

film cooling holes has been completed. In this paper, the thermal field in the flow and on the surface of 

the film cooled flat plate is presented for nominal freestream turbulence intensities of 1.5 and 8 percent. 

The holes are inclined at 30° above the downstream direction, injecting chilled air of density ratio 1.0 

onto the surface of a flat plate. The diameter of the hole is 0.75 in. (0.01905 m) with center to center 

spacing (pitch) of 3 hole diameters. Coolant was injected into the mainstream flow at nominal blowing 

ratios of 0.5, 1.0, 1.5, and 2.0. The Reynolds number of the freestream was approximately 11,000 based 

on hole diameter. Thermocouple surveys were used to characterize the thermal field. Infrared thermography was used to determine the adiabatic film effectiveness on the plate. Hotwire anemometry was used to 

provide flowfield physics and turbulence measurements. The results are compared to existing data in the 

literature. The aim of this work is to produce a benchmark dataset for Computational Fluid Dynamics 

(CFD) development to eliminate the effects of hole length to diameter ratio and to improve resolution in 

the near-hole region. In this report, a Time-Filtered Navier Stokes (TFNS), also known as Partially 

Resolved Navier Stokes (PRNS), method that was implemented in the Glenn-HT code is used to model 

coolant-mainstream interaction. This method is a high fidelity unsteady method that aims to represent 

large scale flow features and mixing more accurately.
NASA/TM—2013-218086/PART1 2 

Introduction

Gas turbine engines produce thrust by burning fuel in the combustor and accelerating fluid through 

the engine. To burn efficiently in the combustor, air entering the engine inlet must be compressed to a 

high pressure ratio. The fluid exiting the combustor can reach temperatures well in excess of the 

component thermal limits in the high pressure turbine (HPT) section. High operating pressure ratios as 

well as increased turbine inlet total temperatures are expected to be the hallmarks of NASA’s N+3 

engines. In order to cool the surfaces of components in the HPT, part of the air form the compressor is 

bled away and fed through internal passages to the HPT where the relatively cooler fluid is injected 

through discrete holes onto the surface of the hot components. This cooling fluid is deprived of passage 

through the combustor and moreover, when fed into the rotor, has work done on it by the rotor instead of 

being used to turn the HPT blades. In general, an increase in compressor pressure ratio leads to higher 

engine efficiency. This is accompanied by an increase in the temperature exiting the combustor and 

entering the high pressure turbine. 

The motivation, then, of improving our understanding of film cooling (or thermal management in 

general) is to reduce the amount of coolant required to cool the surface of the hot components in the HPT 

by using it more efficiently and by understanding the optimal locations and quantities in which to 

distribute it. This is accomplished through experiment and CFD. Experiments allow testing of new 

cooling concepts and serve as benchmarks to compare computations to. Traditionally, computations have 

a tendency to under predict laterally averaged film cooling effectiveness data even for the simple case of a 

flat plate with a cooling hole. There are large errors associated with measurement techniques either due to 

small temperature gradients or invasive measurement methods. The experimental and computational 

uncertainty leads to excess coolant being used as a safety margin. Up to 20 percent excess cooling can be 

used over the mission of an aircraft jet engine. Presently, the cooling air used to cool turbine components 

comprises approximately 12 percent of the mass flow through the core. It is therefore important to 

perform simple experiments that isolate the various parameters that have an impact on film cooling to 

enable development of turbulence modeling that better predicts the physics of the flow and therefore the 

laterally averaged film cooling effectiveness. The parameters of note for flat plate film cooling using 

cylindrical holes are injection angle (α), blowing ratio (M), density ratio (DR), free stream turbulence 

intensity (Tu), surface roughness, length to diameter ratio (L/D) of the hole and hole spacing (pitch to 

diameter ratio, P/D). The blowing ratio is a ratio of coolant mass flow rate per unit area to freestream 

mass flow rate per unit area. The density ratio is the ratio of coolant to freestream density. The hole 

diameter is D or d. In addition, the manner in which the flow enters the cooling hole plays a significant 

role in determining the film effectiveness. For example, flow may enter the hole from a plenum in which 

case the direction of the flow in the plenum is important (same direction as free stream or at an angle to 

free stream.) 

Turbine blades often have compound angled holes although these are angled so that locally they are 

in the direction of the free stream so as to minimize loss. Typically turbine blades use holes inclined at 

30° to 35° to the local freestream and have L/D values ranging from 2 to 4. Blowing ratios are typically 

high and can approach 2.0 near the leading edge and on the suction side. The density ratios are typically 

around 2.0 due to the large temperature difference between coolant and freestream. Higher density ratio 

coolant jets can stay attached at larger blowing ratios due to their lower momentum ratios as compared to 

low density ratio coolant jets. However, the physics of jet-mainstream interaction is essentially the same. 

It is easier to use ambient air as a coolant for experiments and testing with a density ratio of 1.0 is 

therefore popular. Assuming that the surface is adiabatic, a film cooling effectiveness, η, is defined as

c

aw

TT

TT

−

− =η

∞

∞ (1) 
NASA/TM—2013-218086/PART1 3 

Here, Taw is the adiabatic wall temperature downstream of injection. T∞ is the freestream temperature. For 

low speed flows, recovery temperature, Trec can be used instead of freestream temperature. This is useful 

for finding the effectiveness using an optical method such as infrared thermography. Tc is the coolant exit 

temperature. Film effectiveness η, is a nondimensional measure of the surface temperature downstream of 

injection with maximum value of 1.0 for Taw = Tc and minimum value of 0.0 for Taw = T∞. The use of a 

plenum to feed the coolant tubes results invariably in conduction effects through the plate. If the coolant 

is supplied through tubes this conduction effect is almost nonexistent except in the immediate proximity 

of the hole. A good review of parametric effects on film cooling is provided by Bogard et al. (Ref. 1) so 

only a few relevant studies are cited in this review to provide the reader with some background to the 

work reported in this paper.

A popular dataset used for CFD validation is that of Sinha et al. (Ref. 2). They report span averaged 

and centerline adiabatic film cooling effectiveness on a flat plate with a row of 7 cooling holes inclined at 

35° for density ratios of 1.2, 1.6, and 2.0. The blowing ratios varied from 0.25 to 1.0. The data was taken 

using surface mounted thin ribbon thermocouples at several spanwise and axial locations. The cooling 

holes were spaced 3 diameters apart and were fed from a plenum in a direction normal to the flat plate. 

The L/D for the study was 1.75. The test plate used Styrofoam with a conductivity of 0.027 W/m2

K. For a 

density ratio of 1.2, a 50 K temperature difference between free stream and coolant was used. Uncertainty 

in temperature is reported to be 0.1 K.

The short hole lengths contribute to jet detachment at lower blowing ratios. A comparison with the 

data of Pederson et al. (Ref. 3) shows this trend for all but the lowest blowing ratio. Pederson et al. used 

long holes inclined at 35° to study the effect of density ratio on effectiveness. They used a mass transfer 

technique over a wide range of density and momentum ratios. 

Bons et al. (Ref. 4) performed experiments to study the impact of free stream turbulence on film 

cooling effectiveness downstream of discrete cooling holes injecting coolant into the freestream at an 

angle of 35° from the surface and in the downstream direction. Turbulence intensities from 0.9 to

17 percent were studied. Flow uniformity of 2.5 percent was achieved with free stream velocity of 16 m/s. 

Blowing ratios of 0.55 to 1.5 are reported. The density ratio for the study was 0.95. Surface temperature 

was measured using thermocouples under an Inconel foil that is not heated so that the wall is adiabatic. 

They found that at low blowing ratios, high turbulence intensity led to deterioration of cooling 

effectiveness while at high blowing ratio, high free stream turbulence improves laterally averaged 

effectiveness. This is consistent with what one would expect because high turbulence enhances mixing 

which is useful if the jet is detached but is a source of loss (leading to higher temperature at surface) if the 

jet is attached and already providing film cooling coverage. These results are supported by a study by 

Mayhew et al. (Ref. 5) who used Liquid Crystal Thermography to study adiabatic film cooling 

effectiveness for a configuration similar to that of Bons et al. at various blowing ratios. 

Lutum et al. (Ref. 6) studied the effect of length to diameter ratio on film cooling effectiveness for a 

35° injection hole and blowing ratios from 0.52 to 1.56. They found that the effectiveness increased with 

the length to diameter ratio although the variation is small in the range of 5<L/D<18. It is possible 

however that this result is dependent on plenum boundary conditions which were not taken into account. 

Thermochromic liquid crystal was used in conjunction with thermocouples to determine the adiabatic 

effectiveness.

Surface roughness effects can be a major influence on film effectiveness via its role in changing the 

heat transfer coefficient. At engine scale, roughness of several hundred microns can be on the order of 

magnitude of the hole diameter. Modeling roughness is then an added complexity for CFD. The effects of 

roughness are very similar to those of free stream turbulence in that increased roughness can be beneficial 

at high blowing ratios while it is detrimental at lower blowing ratios.

One of the consistent problems with the aforementioned experiments is that they required invasive 

techniques to determine the adiabatic effectiveness. A benchmark data set for film cooling should allow 

the computational modeling effort to focus on solving the mixing problem in the near-hole region without 

having to introduce conjugate heat transfer. Infrared thermography is a non-invasive technique thereby 
NASA/TM—2013-218086/PART1 4 

allowing CFD to model the experiment more closely. There is also a lack of uncertainty analysis related 

to laterally averaged film effectiveness. To eliminate the effect of the plenum on heat transfer and flow 

physics in the hole, it is desirable to use long holes without a plenum. 

Thurman et al. (Ref. 7) presented results from a study that used long holes (L/D~15) for blowing 

ratios of 1 and 1.9. The basis for the use of long tubes for the coolant holes was to yield a well-developed 

velocity profile at the hole exit. Span-averaged effectiveness values were not reported in the study of 

Thurman et al. and the focus was to qualitatively compare an anti-vortex hole geometry to a round hole 

geometry. Nusselt number distributions on the surface were shown using a Liquid Crystal technique.

Han et al. (Ref. 8) provide a review of measurement techniques used in studying film cooling. These 

range from surface mounted thermocouples for surface temperature measurements to Particle Image 

Velocimetry for full field flow visualization. Infrared cameras have now been used with some regularity 

for qualitative thermal imaging and for non-destructive evaluation of engine components. Infrared 

cameras measure radiation not just from the object being viewed but also radiation reflected by the object 

and emanating from other environmental sources in the vicinity. There is also a loss of radiation due to 

atmospheric absorption or due to the optical properties of the viewing window. These properties change 

with time and must therefore be accounted for during each measurement to obtain accurate temperatures. 

This requires the use of multiple reference temperatures. The parameters that influence the temperature 

measurement are emissivity of the object surface, the reflected apparent temperature, the distance 

between the object and the camera and the ambient or background temperature. Over short distances, the 

atmospheric transmission losses (influenced by relative humidity) are small enough to be neglected. 

However, the use of an optical window introduces another loss mechanism that must be accounted for. 

As early as 1979, Papell et al. (Ref. 9) used infrared images to study the variation in film cooling 

effectiveness due to the curvature of the coolant tubes. Blowing ratios of 0.37 to 1.25 were tested using a 

temperature difference of 25 K between coolant and tunnel inlet. Baldauf et al. (Ref. 10) used infrared 

images to show surface effectiveness on a flat plate with thermal conductivity of 0.3 W/m2

K. The hole 

length was six times the diameter and pitch to diameter ratios of 2, 3, and 5 were tested. Density ratios of 

1.2, 1.5, and 1.8 were reported. Ekkad et al. (Ref. 11) used a transient infrared thermography technique to 

determine simultaneously the surface effectiveness and heat transfer coefficients for a single hole on a 

half cylinder using a FLIR SC3000 camera. They used a reference thermocouple to measure the surface 

temperature without flow. The emissivity of the surface was varied to obtain a temperature on the infrared 

camera that matched the thermocouple temperature.

Wright et al. (Ref. 12) used pressure sensitive paint to obtain surface effectiveness distributions at a 

density ratio of 1.4. Blowing ratios were varied between 0.25 to 2.0 for free stream turbulence intensities 

between 1 to 12.5 percent for cylindrical and fan shaped holes inclined at 35° and spaced 4 hole diameters 

apart.

The present study includes detailed measurements of turbulence, velocity and temperature (presented 

in this report) using thermocouple surveys and infrared thermography for two different nominal free 

stream turbulence intensities: 1.5 percent, referred to as ‘low Tu’ and 8 percent, referred to as ‘high Tu’ or 

‘with turbulence grid’. The turbulent intensity at the CFD inlet plane location is estimated to be 2 percent

for the ‘low Tu’ case and 14 percent for the ‘high Tu’ case. Air that was chilled by passing it through an 

ice bath was used as coolant. The coolant was introduced into the main stream through long tubes 

(L/D>18). Both span averaged and centerline effectiveness are reported. Infrared thermography allows for 

full field thermal mapping of the test section. 

For film-cooling flows in the higher blowing ratio range (0.8 to 2.0), Reynolds Averaged Navier 

Stokes (RANS) simulations do not produce accurate solutions (Refs. 13 and 14). Argument is made that 

RANS models are not adequate for film-cooling flows because they involve unsteady interactions and 

large-scale mixing of the main stream and cooling stream. This needs to be better accounted for by direct 

modeling of these interactions. In recent years, attempts have been made to model film cooling flows 

using Large Eddy Simulation (LES) and Detached Eddy Simulation (DES). Some of these attempts have 

resulted in computations of cooling effectiveness while others have used LES to shed light on the physics 

of film cooling. Guo et al. (Ref. 15) studied blowing ratios of 0.1 and 0.48 for normal and 30° inclined 
NASA/TM—2013-218086/PART1 5 

angles using LES. Their purpose was to investigate the turbulent flow structure and the vortex dynamics 

for gas turbine blade film cooling. Their work showed that the turbulent flow downstream of the hole is 

highly anisotropic and that LES is well suited for such simulations. Tyagi and Acharya (Ref. 16) studied 

film cooling physics for a blowing ratio of 0.5. They showed good agreement with experiment for 

velocity distributions downstream of the hole and shed light on the physical features of the film cooling 

flow such as the resolution of hairpin vortices downstream of the holes. More recently, Fujimoto (Ref. 17) 

used surface-adjusted octree hexahedral meshes to perform LES computations of film cooling 

effectiveness. He showed a comparison of his results and the results of a multi-block LES computation 

with the experimental data of Sinha et al. (Ref. 2). He reported good agreement with the experimental 

data for a range of blowing ratios. 

One of the issues often confronted when computing film-cooling flows is the effect of the plenum. 

Modeling the plenum and the assumptions concerning such modeling is a source of uncertainty. Including 

the plenum in the numerical model is often done but the incompressible flow within the plenum reduces 

the convergence rate of the numerical schemes. This limitation has little to do with the actual engine 

conditions and is encountered because experimental setups use plena for film cooling. To realistically 

model the flow the plena have to be included in the computations. Earlier works in LES for film-cooling 

flows includes the work of Peet and Lele (Ref. 18) who used a combination of CFD codes namely an 

incompressible method for the plenum and the pipe and a compressible code for the cross-stream and the 

interaction zone. They computed turbulence statistics and effectiveness of the film-cooling for a low, 

blowing ratio case of M = 0.5. 

In the present study a Time-Filtered Navier-Stokes (TFNS) method is used to compare film cooling 

effectiveness computations of long-hole film cooling to infrared thermography and thermocouple surveys. 

This method is also referred to as Partially Resolved Navier Stokes or PRNS and the two terms (TFNS 

and PRNS) are used interchangeably. Flow over a flat plate and temperature distribution downstream of 

film cooling holes at a 30° inclination at blowing ratios of 0.5, 1.0, 1.5, and 2.0 for a single nominal 

density ratio of unity were simulated. The range of momentum ratios is thus the same as the blowing 

ratios. Some earlier computations (Refs. 13 and 14) performed at the blowing ratio of M = 2, showed the 

inadequacy of the k-ω turbulence model (Ref. 19) at high blowing ratios. In this study, several turbulence 

models within the NASA Glenn-HT code (Ref. 20) were used for an intermediate blowing ratio of unity. 

In addition, a commercial code, CFX and the models within it (SST and k-ε) were also applied and it was 

verified that high blowing ratio cases were not well reproduced. Near-hole results were especially of 

concern as the discrepancy between the predictions and data was large enough to require reconsideration 

of the method of data-acquisition in that locality. 

The unsteady approach called the TFNS, PRNS, or Very Large Eddy Simulation in earlier papers by 

the developer team at NASA) (Refs. 21 and 24), was developed at NASA Glenn Research Center. The 

method was developed for the simulation of large turbulent eddies requiring grid resolution comparable to 

what is often used in RANS simulations. The method is based on the concept of temporal filtering. This 

avoids issues associated with coarse grids that directly influence the solutions in LES due to the filter size 

being directly connected to the grid resolution. In TFNS, the larger time scales (or lower frequencies) of 

the turbulence are directly calculated and the effects of the unresolved time scales of the turbulence are 

modeled by a dynamic equation system. The contents of both the resolved and unresolved turbulence are 

regulated by a “Filtering Control Parameter” (FCP), which is related to the width of the temporal filter. 

The basic equations of TFNS and its subscale model, as presented by Shih and Liu (Ref. 23) are grid 

invariant, i.e., they do not have grid spacing as a parameter in their formal formulations. Therefore, it is 

possible to achieve a grid-independent numerical solution. Shih and Liu (Ref. 23) count this as a major 

point of departure from the traditional LES approach and its variants. Another distinction is that TFNS 

enables performance of unsteady RANS (URANS), LES, and even DNS in a unified way through the 

judicious selection of the value of FCP, and the appropriately refined grid. It should be noted that the 

TFNS approach is not a variant of the popular hybrid RANS/LES (e.g., Refs. 25 to 28). There is no 

enforced transition between the perceived RANS and LES domains.
NASA/TM—2013-218086/PART1 6 

The subfilter model is constructed to use a more general relationship between the unresolved 

turbulent stresses and the resolved turbulent flow field, Shih and Liu followed the analysis of the rational 

mechanics and obtained a general constitutive relationship for the Reynolds stresses following (Ref. 23). 

The constitutive relationship in addition to an eddy-viscosity term includes several other terms 

representing the anisotropy and the rotation effects due to the interactions between the resolved and 

unresolved turbulence. In addition to the dissipative and diffusive effects accounted for through the eddy 

viscosity, the effects of anisotropy and rotation are accounted for in the subscale model, when the 

simulation pertains to the very large eddy simulation. Source terms introduced in the momentum equation 

sustain the turbulent fluctuations in the calculated flow field. One important outcome of the use of this 

constitutive relationship is the ability of the TFNS to sustain turbulent fluctuations for a turbulent pipe 

flow in the lower range of applicable Reynolds numbers. The k-ε variant was not able to sustain a 

turbulent flow physically. The model adopted for the TFNS simulations in the Glenn-HT code employs 

this unique feature of the nonlinear subscale model. 

Setup of Experiment

A 30 times engine scale test facility was used to interrogate the complex, highly three-dimensional 

flow field associated with detached flows and to specifically resolve the shear layers and wake regions. A 

schematic of the flow and geometry for the test are shown in Figure 1. Experiments were carried out in 

the wind tunnel shown in Figures 2 to 4. This is an open loop tunnel with a temperature controlled coolant 

loop. The tunnel consists of an aluminum bellmouth, flow conditioning screens, square acrylic sections 

that are 0.208 m wide and 0.0191 m thick, the test section on the floor of the tunnel and a lid directly 

above it for either viewing or actuator support. Air was drawn from the room and passed through flow 

conditioning sections prior to entering the test section. 

Airflow was provided by a 5220W fan. The freestream velocity was 9.1 m/s and the Reynolds 

number based on freestream velocity and coolant hole diameter was 11,000. The test section is a square 

section of cross section measuring 0.2083 by 0.2083 and 0.8636 m in length. The freestream turbulence 

intensity measured approximately 1.5 percent without a grid and 8 percent with a square grid based on 

hotwire surveys. The boundary layer thickness at x/d = –0.5 is 0.0127 m and is taken to be the vertical 

distance from the wall at which the velocity is equal to 99 percent of the freestream velocity. 

Figure 1.—Schematic of cooling flow.
NASA/TM—2013-218086/PART1 7 

Figure 2.—SW-6 low speed wind tunnel.

Figure 3.—CAD model of SW-6 facility.

Figure 4.—Schematic of test setup.
NASA/TM—2013-218086/PART1 8 

The coolant flow was provided by blowing pressurized supply air through a heat exchanger, which 

consisted of a copper tube coiled inside an ice-water tank. The coolant was then fed through a manifold to 

three separate flow meters, and then to acrylic tubes with L/D = 20 connected to each cooling hole. The 

separate cooling tubes provided an opportunity to model the cooling flow in each hole as fully developed. 

The test section (Figs. 5 and 6) was a flat plate made of acrylic with three holes inclined at 30° and a 

hole diameter of 0.019 m. The hole spacing was y/d = 3. To determine the tunnel flow rate, a total 

pressure probe was placed upstream of the test section and static pressure taps were placed on the 

sidewalls. Freestream temperature was measured with an open-ball thermocouple located upstream of the 

holes near the total pressure probe. Coolant temperature was measured with open-ball thermocouples 

inside the coolant tubes. Temperature survey data was taken along the centerline plane of the tunnel and 

at several cross sectional planes with a small, open ball, type E thermocouple probe attached to an 

actuator above the test section. The adiabatic wall temperature was determined from the thermocouple 

probe surveys with the probe located near the floor of the test section. Various hotwire probes were used to 

obtain the three-dimensional velocity components and turbulent stresses along the centerline and at several 

streamwise planes.

Figure 5.—Film cooling test plate with long holes. 

Figure 6.—Bottom view of long holes. 
NASA/TM—2013-218086/PART1 9 

Description of Infrared (IR) Measurement Setup

To obtain the adiabatic film cooling effectiveness a FLIR SC655 infrared camera was used with 

multiple temperature ranges. For the work reported the range of –293.15 to 423.15 K was used. This 

corresponds to a wavelength of approximately 10 μm. A ZnSe window was used with a transmittance of 

0.96 in the 9 to 11 μm range. The resolution was approximately 81 pixels/in. or 3.2 pixels/mm. This 

corresponds to 108 pixels/D. The specified accuracy of the camera is 2 percent of the measured 

temperature. The camera is power linear with object signal voltage described by the FLIR user manual 

(2010) to be

obj tot ref atm

11 1 UU U U

ετ

τ− − εε− − ετ = (2) 

This equation can be rewritten as 

( )( ) ( )( ) totobj ref obj atm obj UU 1 UU 1 −τ−−−ε−τ−= UU (3) 

It is clear that if the emissivity is high and if the reflected temperature and atmospheric temperatures 

are close to the temperature being measured, the object signal output can be approximated by the total 

signal. The temperature range being examined is approximately between 285 and 300 K corresponding to 

an object signal range of 117 to 137. The maximum error due to neglecting reflection and background 

corrections is 1.49 percent of object signal or ±1 K. 

Figure 7 shows a plot of object signal versus temperature. A quadratic fit was found to be the 

appropriate choice although the linear fit is also a good approximation for this temperature range. Based 

on the data fit the temperature is related to camera output signal as 

aUT mU ++= C 2 (4) 

where T is temperature in Kelvin and U is the camera output signal. The R2 value for the fit is 1. The 

constants a, m, and C are evident form the equations in Figure 7. 

Figure 7.—Relationship of camera indicated temperature to object signal. 
NASA/TM—2013-218086/PART1 10

Figure 8.—Infrared thermograph showing regions of interest. 

Film cooling effectiveness can be calculated as

rec c

rec aw

TT

TT

−

− 1method =η (5) 

where each value on the right hand side of the equation is a surface measurement from a single IR image.

Figure 8 shows an example of an IR image with regions of interest marked for reference. 

The thermocouple location for reference measurements is marked in the lower left corner as 

‘Thermocouple Location’. The recovery temperature, Trec, is measured using a 9-pixel average at the 

upper left location marked ‘Trec’. This marker is located at a spanwise distance of 0.75 p/d from the left 

coolant hole leading edge. Coolant temperatures are measured using three different regions of interest. An 

elliptical region consisting of 4598 pixels, labeled ‘Left Coolant Average’, is used to measure average 

coolant temperature over the visible portion of the left hole wall. A 9-pixel marker labeled ‘Coolant Exit 

Temperature’ measures the exit temperature on the wall of the center cooling hole nearest to the cooling 

hole exit. It is located at a distance of 1.7 D upstream of the hole trailing edge and along the centerline. 

This corresponds to a location of y/d = –0.1 at which the coolant temperature from CFD is extracted.

A rectangular averaging marker labeled ‘coolant average’ measures coolant exit temperature over 

612 pixels. The temperatures measured by these three different markers are within 0.1 K. For 

effectiveness, the temperature measured by the ‘Coolant Exit Temperature’ marker is used to match 

results from CFD. The line marker ‘eta sweeper’ is used to sweep the surface between the left and right 

symmetry markers shown in Figure 8 between x/d = 0 and x/d = 6.0. The marker width is equal to 1 pitch. 

The skewness of the image is compensated for in the calculations. The error in location coordinates arises 

from converting pixels to physical units. In this case, the primary reference for measurement is two 

marker points spaced 0.0381 m (1.5 in.) apart and shown in Figure 8 as ‘1.5 in. Between Markers’. In 

addition, the distance between two pairs of metal tapes is also known with the distance specified by 

‘Between Tapes X’ being 0.127 m (5 in.) and ‘Between Tapes Z’ being 10.4775 m. The corresponding 

pixel locations are used to determine the ratio of pixels to physical length. The hole diameter is then 

ascertained based on this conversion and is found to be accurate to 0.5 pixels or 0.0001524 m for all 
NASA/TM—2013-218086/PART1 11

cases. These calculations are all performed using an image exported from the infrared camera and 

processed by in-house written MATLAB (The Mathworks, Inc., Natick, MA) routines. It is evident that 

as the dimension of the hole shrinks, the uncertainty of measurement location as a function of hole 

diameter will increase.

If it is assumed that the viewing area is small enough that the reflected apparent temperature does not 

vary significantly over it and that the surface emissivity and optical transmittance are uniform, the 

camera’s total output temperature can be used instead of the true object temperature. This allows for rapid 

effectiveness calculations without the need to adjust reflected temperature settings during each 

measurement. The uncertainty introduced to effectiveness is less than ±0.002 based on an emissivity 

range of 0.9 to 0.98 and a transmittance range of 0.95 to 0.975. Reflected apparent temperature may vary 

between 273.15 and 295 K. The actual uncertainty is much smaller because the reflected temperature is 

much closer to the ambient temperature. For the effectiveness calculated using method 1, the reflected 

temperature was set to ambient temperature. The emissivity was set to 0.95 and the transmittance to 0.96 

based on the material data sheet for the ZnSe window that was used. The uncertainty for the effectiveness 

values reported here is ±0.002 based on the definition of effectiveness used. 

The effectiveness can also be calculated as, 

c

aw

TT

TT

−

− =η

∞

∞ 2method . Here, the tunnel inlet temperature, 

T∞, is used instead of the recovery temperature. This temperature is obtained using the type E

thermocouples described in the previous section. The coolant temperature is also obtained using 

thermocouples. In this case, the emissivity and reflected temperature are varied until the temperature 

measured by the IR camera at the reference thermocouple location is within ±0.05 °C of the temperature 

indicated by the reference thermocouple and the coolant exit temperature matches the thermocouple 

measured coolant temperature within ±0.05 °C. The uncertainty in transmission is ±0.01. 

Setup of Numerical Domain

Film cooling flow computations were undertaken using the NASA Glenn-HT (Ref. 13) code. The 

code is a multi-block structured-grid N-S solver. The computational domain extended 12 cooling hole 

diameters upstream of the cooling hole and 16 diameters downstream. Spacing between the holes in the 

experiment was equal to three hole-diameters and there were three holes. For the computations a period of 

3 diameters was imposed. It was however understood that due to flow unsteadiness, imposition of 

periodicity is a simplification that is done for the sake of reducing resource requirements for the 

computations. The free-stream was placed at 5 diameters from the wall. A number of grid blocks were 

assigned to individual CPUs in a balanced manner to allow parallel computing. To reduce the 

communications overhead, smaller blocks were further consolidated before the grouping was 

implemented. For the present computations, initially 120 groups (CPUs) were used but by refining the 

blocking, it was possible to increase the number of CPUs to 1200 for better parallel capability. The grid 

itself contained approximately 8 million cells. Particular attention was paid to refinement of the grid in 

areas near the no-slip walls and near the cooling hole outlet and especially downstream of the cooling 

hole. The grid cells in the core of the film cooling flow were constructed to be uniform and near cubic. 

The overall grid topology, the symmetry plane of the computational domain, as well the near-hole grid 

shown from two views are presented in Figures 9(a) and (b). 

The simulations were conducted according to the conditions provided by the experiments. Geometric 

scaling was performed to raise the flow Mach number while maintaining the Reynolds number based on 

the film cooling hole. The density ratio of the flow was held to near unity by adjusting the total 

temperature at the hole inlet. Blowing ratios of 0.5, 1.0, 1.5, and 2.0 were computed. The cases were run 

primarily using the TFNS model with the Filtering Control Parameter, 0.16 T

T

  ∆

  =   which produced a 

factor of fi = 0.3 (Ref. 23).
NASA/TM—2013-218086/PART1 12

Figure 9.—Computational grid showing views of the grid near the hole exit (a) and (b), the computational domain (c) 

and the grid blocking for parallel computing (d).

As mentioned earlier, experience suggests that the film cooling effectiveness cannot be predicted with 

RANS when the blowing ratio is relatively high (>~1). The computations at high blowing ratios presented 

here, attempt to model the film cooling flow with a TFNS model to verify that a better film cooling 

effectiveness prediction may be obtained. 

Results and Discussion

In this section results from CFD, thermocouple surveys and infrared thermography are shown. The 

results are divided into sections based on nominal blowing ratio. For blowing ratio of 2.5 only infrared 

thermography data is presented. 

Blowing Ratio = 0.5 

Figure 10 shows the surface effectiveness values computed according to method 1 described in the 

previous section as measured by the infrared camera. The values range from 0 (indicating no cooling) to 

1.0 (indicating the coolant temperature or maximum effectiveness.) For this case the image shows some 

asymmetry in the surface temperature downstream of the holes located at y/d = 3 and y/d = –3. The flow 

exiting the central hole is very similar to that exiting the hole at y/d = 3. It is possible that the flow exiting 

the hole at y/d = –3 is warmer due to an imperfect seal formed by the thermocouple in the coolant tube. 

The region of interest however is downstream of the central hole located at y/d = 0.0 so the variation in 

temperature exiting the adjacent hole is a second order effect. There is a necking of the surface coverage 

immediately downstream of the hole followed by a rise in effectiveness further downstream up to x/d = 1 
NASA/TM—2013-218086/PART1 13

followed by a steady drop in effectiveness. The region between y/d = 1 and y/d = 2 sees little cooling 

downstream of the hole. The near hole region however appears to be cooled through the action of the 

horseshoe vortex and mixing of the coolant with the freestream. The contours in the cooling hole do not 

show a uniform effectiveness of 1.0 over the area of the hole because the IR camera shows the 

temperature on the surface of the coolant tube as a projection. A key point to note therefore is that the 

coolant exit temperature indicated by the IR camera is different from the temperature measured by the 

thermocouples in the coolant tube. 

Figure 10.—Film cooling effectiveness from (a) infrared 

thermography and (b) CFD for M = 0.5, DR = 1.0, Tu =

1.5 percent. 
NASA/TM—2013-218086/PART1 14

Figure 11.—Film effectiveness from thermocouple surveys for M = 0.5, 

DR = 1.0, Tu = 1.5 percent.

Figure 12.—Q-criterion showing Kelvin-Helmholtz instability at M = 0.5, 

DR = 1.0, Tu = 1.5 percent.

The temperature rise from coolant measurement location to hole exit would shift effectiveness 

measured using thermocouples by up to 0.04 higher. However, it is impractical to place a thermocouple 

however small at the coolant hole exit due to its influence on the jet. It is important therefore to treat the 

thermocouple measurements independently of the infrared measurements when comparing to future or 

existing datasets. Figure 11 shows effectiveness measured by thermocouple surveys and processed using 

Tecplot. In Figure 11, planes at x/d locations 0.133, 2.133, 4.133, and 6.133 downstream of the cooling 

hole are shown along with a plane at the centerline. Flow is from left to right along the x/d axis. The 

contours show the coolant core sticking close to the surface and gradually dissipating while providing a 

wide area of cooling. 

Figure 12 shows iso-surfaces of the Q-criterion (Ref. 29) which represents locations in the flow where 

the rotation dominates the strain and shear. The cooling fluid with the higher density is shown to 

propagate downstream with periodic mixing patterns appearing first at a location slightly upstream of the 
NASA/TM—2013-218086/PART1 15

hole and staying regular some distance downstream before it becomes more random. Q-criterion serves to 

highlight vortical structures within the turbulent flow. The patterns indicate definite ring vortices that are 

formed around the film column and convect downstream. These rings are quite regular and appear to be 

of the Kelvin-Helmholtz instability type. Further downstream of the hole the patterns appear to have 

dissipated due to turbulent mixing. The surfaces are colored by non-dimensional temperature, θ = 1–η. 

Thus, the blue regions show cooling flow while the red regions are at freestream conditions. The coherent 

vortical structures persist for several hole diameters downstream of the hole and then break up. This 

coincides with the rapid deterioration in effectiveness seen in Figure 10 especially in the spanwise 

direction (seen by sharp narrowing of the color bands.) This is akin to waves breaking up and churning up 

the ocean below. It is possible that the prediction of the breakup of these vortices could be the key to 

accurate prediction of film cooling effectiveness far downstream of the hole. 

Figure 13 shows film effectiveness at the jet centerline (y/d = 0) from thermocouple surveys. A 

curvature of the coolant jet is seen at x/d = 1 that is coincident with the rise in effectiveness observed in 

Figure 10 at the same location. This indicates the migration of the jet core toward the surface due to the 

influence of the freestream. The influence of the jet extends to z/d = 0.3 at the hole trailing edge and to 

z/d = 1.1 at x/d = 5.25. 

Figure 14 shows contours of effectiveness at x/d = 0.133 (top), 2.133, 4.133, and 6.133 (bottom) 

downstream of the cooling hole. The spanwise extent (y/d) of the figures is 1 pitch (3 hole diameters.) 

The core is seen to stick close to the surface and the jet spreads out gradually as it mixes with the 

freestream. The effectiveness of the core drops from 1.0 at the hole exit to 0.7 at x/d = 6.133. The cooling 

jet spreads gradually downstream of the hole in agreement with the IR image form Figure 10. The vertical 

extent grows from 0.7 to 1.3 y/d from x/d = 0.133 to x/d = 6.133. The coldest region of the jet also 

changes from elliptical with a spanwise major axis to elliptical with major axis in the vertical direction.

Figures 15 and 16 show span-averaged and centerline effectiveness for M = 0.5, respectively. The 

cyan line shows effectiveness from IR thermography for the high Tu case (8 percent). The results from IR 

thermography for low turbulence lie within the error bands of the thermocouple data for low turbulence. 

The black dashed lines bound the results from IR measurements with the red dashed line representing an 

IR measurement that has been calibrated using a surface mounted thermocouple and a thermocouple in 

the coolant tube.

This proves that at the small temperature ranges in use for this experiment, as long as effectiveness is 

computed using values from a single image, no significant error or penalty is incurred. However the same 

assumption does not extend to larger temperature differences that would require careful calibration using 

several surface mounted thermocouples at a range of temperatures that include the coolant and free stream 

temperature.

It is interesting to note that the CFD results show a relatively high effectiveness maintained much 

further downstream of hole than the experiment. The IR results shown in Figure 16, show that higher 

turbulence intensity is initially beneficial to mixing the jet at the hole exit but then causes a deterioration 

in centerline effectiveness downstream as the near-surface coolant is forced to mix with the freestream 

more rapidly and spread out. The span-averaged effectiveness however is higher for the case of high 

turbulence. This is consistent with the results of Brown et al. (Ref. 30) who found that above a blowing 

ratio of 0.7 higher turbulence intensity caused higher span averaged effectiveness but reduced centerline 

effectiveness. Figure 11 shows a ‘necking’ of the effectiveness contours downstream of the hole and this 

is indicative of the lift-off tendency of the jet. At higher density ratio, this lift-off would occur at a lower 

blowing ratio due to the increased momentum of the jet. Figure 14 shows that CFD predicts the KelvinHelmholtz structures persist to approximately x/d = 4 at which point they begin to break up. It would 

suggest that this coherence is the cause of the relatively high effectiveness shown by CFD for low 

blowing ratios. At higher blowing ratios, it is conceivable that this same effect would have deleterious 

consequences. Downstream of the breakup region, the TFNS model, causes a rapid mixing of the jet and 

freestream resulting in an increased rate of decline of effectiveness.
NASA/TM—2013-218086/PART1 16

Figure 13.—Centerline effectiveness from (a) CFD, (b) thermocouple at 

M = 0.5, DR = 1.0, Tu = 1.5 percent.
NASA/TM—2013-218086/PART1 17

Figure 14.—Effectiveness contours from (a) CFD, (b) thermocouples at planes of constant x/d for

M = 0.5, DR = 1.0, Tu = 1.5 percent. 
NASA/TM—2013-218086/PART1 18

Figure 15.—Spanwise-averaged film effectiveness for M = 0.5, DR = 1.0. 

Figure 16.—Centerline film effectiveness for M = 0.5, DR = 1.0. 
NASA/TM—2013-218086/PART1 19

Blowing Ratio = 1.0 

Surface effectiveness (Fig. 18) for a blowing ratio of 1.0 from IR thermography for the low 

turbulence intensity case. The maximum effectiveness downstream of the hole has dropped to 

approximately 0.45 and occurs at approximately 2.5 diameters downstream of the hole trailing edge. In 

addition, the contours appear to be pinched and aligned with the flow direction. This indicates that the jet 

core may already be lifted away from the surface. 

The contours in Figure 17 indicate that there is uniform cooling that seems to extend far downstream 

and does not drop off as quickly as for M = 0.5. The coverage appears only to extend to y/d = ± although 

there is a gradual spreading of the jet as it progresses downstream. There is a sharp neck at x/d = 0.5, 

Figure 17.—Film cooling effectiveness computed from (a) infrared 

thermography and (b) CFD for M = 1.0, DR = 1.0, Tu = 1.5 percent. 
NASA/TM—2013-218086/PART1 20

although, it is not as pronounced as it is for M = 0.5 because the jet trajectory is now directed away from 

the surface. Near the hole exit where the coolant temperature is measured, the coolant is influenced by the 

jet mixing that causes a drop in effectiveness. The contour rings around the hole are a combined effect of 

the horseshoe vortex and to a lesser extent, localized conduction due to the finite thickness of the flat 

plate. Figure 18 shows effectiveness calculated from thermocouple surveys and confirms that the jet is in 

fact lifting from the surface. Planes at x/d locations of 0.133, 2.133, 4.133, and 6.133 downstream of the 

cooling hole are shown along with a plane at the centerline. Flow is from left to right along the x/d axis. 

Figure 19 shows the local effectiveness at a plane located at y/d = 0.0 based on thermocouple surveys. 

The jet influence extends to approximately z/d = 1.4 at x/d = 1.5. The jet can be thought of as following a 

parabolic trajectory with a peak at x/d = 1 centered at y/d = 0.4. Notice the region between x/d = 0.5 to 

2.0 where there appears to be a region of relatively lower effectiveness. The coherent hairpin vortices that 

are lifted off the surface allow the hot mainstream gas to be entrained under the jet. As the coherent 

structures breakup, coupled with the downward trajectory of the jet, the effectiveness increases near the 

surface again. Figure 20 shows contours of effectiveness at planes of constant x/d. The kidney shape seen 

at x/d = 2.133 is seen to become less resolved at x/d = 4.133 as the jet loses coherence.

Figures 21 and 22 show span-averaged and centerline effectiveness for M = 1.0 at DR = 1.0. The 

thermocouple measurements show a consistently higher effectiveness in the near hole region as the KelvinHelmholtz vortices are broken up and promote mixing. In the downstream region, the thermocouple results 

match well with IR. The CFD result starts out over predicting effectiveness because as seen from Figure 12,

the hairpin vortices propagate far downstream before they break up. The pink line shows effectiveness with 

the turbulence grid installed and the trends follow the discussion in the previous section.

Figure 18.—Film effectiveness from thermocouple surveys for M = 1.0, DR = 1.0, 

Tu = 1.5 percent. 
NASA/TM—2013-218086/PART1 21

Figure 19.—Centerline effectiveness from (a) CFD, (b) thermocouple survey at 

M = 1.0, R = 1.0, Tu = 1.5 percent. 
NASA/TM—2013-218086/PART1 22

Figure 20.—Effectiveness contours from (a) CFD, (b) thermocouples at planes of constant x/d for

M = 1.0, DR = 1.0, Tu = 1.5 percent. 
NASA/TM—2013-218086/PART1 23

Figure 21.—Spanwise-averaged film effectiveness for M = 1.0, DR = 1.0. 

Figure 22.—Centerline film effectiveness for M = 1.0, DR = 1.0. 
NASA/TM—2013-218086/PART1 24

Blowing Ratio = 1.5 

Figure 23(a) shows surface effectiveness for a blowing ratio of 1.5 from IR thermography for the low 

turbulence intensity case (Tu = 1.5 percent). Figure 23(b) shows surface effectiveness from the TFNS 

simulation. The maximum effectiveness downstream of the hole has dropped to approximately 0.25 and 

occurs between x/d = 2 and x/d = 3. The jet is completely detached. The CFD shows negligible 

effectiveness in the region of detachment and this could be due to a lack of prediction of mixing of the 

coherent hairpin structures with the main flow. 

Figure 23.—Film cooling effectiveness from (a) infrared 

thermography and (b) CFD for M = 1.5, DR = 1.0, Tu = 

1.5 percent.
NASA/TM—2013-218086/PART1 25

Figure 24.—Film effectiveness from thermocouple surveys for M = 1.5, DR = 1.0, 

Tu = 1.5 percent.

Figure 24 shows effectiveness calculated from thermocouple surveys. Planes at x/d locations of 

0.133, 2.133, 4.133, and 6.133 downstream of the cooling hole are shown along with a plane at the 

centerline. Flow is from left to right along the x/d axis. Figure 25 shows Q-criterion for M = 1.5 from the 

TFNS simulation. The same features seen in Figure 12 for M = 0.5 are seen but the jet has now penetrated 

further into the freestream in the spanwise and vertical direction. The Kelvin-Helmholtz vortices persist 

up to the same downstream extent as in the M = 0.5 case. 

Figure 26(a) shows the effectiveness at a plane located at y/d = 0.0 based on TFNS while 

Figure 26(b) shows effectiveness for the same region from thermocouple surveys. The jet influence 

extends to approximately z/d = 2.0 at x/d = 1.5 for the thermocouple survey whereas CFD predicts the jet 

to extend to only z/d = 1.5. The thermocouple survey seems to indicate that the jet is mixing out rapidly 

and provides fairly constant cooling effectiveness as it travels downstream. The CFD solution has a ‘dead 

zone’ of effectiveness between x/d = 0.2 and x/d = 2.0. The coolant is securely carried within the coherent 

vertical structures and delivered downstream when the coherence is broken. Figure 27 shows contours of 

effectiveness at planes of constant x/d. The kidney shape seen at x/d = 2.133 is seen to become less 

resolved at x/d = 4.133 as the jet mixes with the freestream.
NASA/TM—2013-218086/PART1 26

Figure 25.—Isosurfaces of Q-criterion showing Kelvin-Helmholtz instability at

M = 1.5, DR = 1.0, Tu = 1.5 percent.

Figure 26.—Centerline effectiveness from (a) CFD and (b) 

thermocouple at M = 1.5, DR = 1.0, Tu = 1.5 percent.
NASA/TM—2013-218086/PART1 27

Figure 27.—Effectiveness contours from (a) CFD, (b) thermocouples at planes of 

constant x/d for M = 1.5, DR = 1.0, Tu = 1.5 percent.
NASA/TM—2013-218086/PART1 28

Figures 28 and 29 show span-averaged and centerline effectiveness for M = 1.5 at DR = 1.0. The 

thermocouple measurements show a consistently higher effectiveness in the near hole region as the 

Kelvin-Helmholtz vortices are broken up and promote mixing. In the downstream region, the 

thermocouple results match well with IR. The CFD result starts out over predicting effectiveness because 

as seen from Figure 12, the hairpin vortices propagate far downstream before they break up. The pink line 

shows effectiveness from IR thermography with the turbulence grid installed and the trends follow the 

discussion in the previous section. The centerline effectiveness degrades more rapidly due to the 

underlying freestream fluid in the near-hole region. 

Figure 28.—Spanwise-averaged film effectiveness for M = 1.5, DR = 1.0. 

Figure 29.—Centerline film effectiveness for M = 1.5, DR = 1.0. 
NASA/TM—2013-218086/PART1 29

Blowing Ratio = 2.0 

Figure 30(a) shows surface effectiveness for a blowing ratio of 2.0 from IR thermography for the low 

turbulence intensity case (Tu = 1.5 percent). Figure 30(b) shows surface effectiveness from the TFNS 

simulation. The maximum effectiveness downstream of the hole has dropped to approximately 0.125 and 

occurs between x/d = 4 and x/d = 5. The jet is completely detached as in the case of M = 1.5. The CFD 

shows similar trends to those of the M = 1.5 case. 

Figure 30.—Film cooling effectiveness from (a) infrared 

thermography and (b) CFD for M = 2.0, DR = 1.0, 

Tu = 1.5 percent.
NASA/TM—2013-218086/PART1 30

Figure 31 shows effectiveness calculated from thermocouple surveys. Planes at x/d locations of 

0.133, 2.133, 4.133, and 6.133 downstream of the cooling hole are shown along with a plane at the 

centerline. Flow is from left to right along the x/d axis. Figure 32 shows Q-criterion for M = 1.5 from the 

TFNS simulation. The same features seen in Figure 12 for M = 0.5 are seen but the jet has now penetrated 

further into the freestream in the spanwise and vertical direction. 

Figure 33(a) shows the effectiveness at a plane located at y/d = 0.0 based on TFNS while 

Figure 33(b) shows effectiveness for the same region from thermocouple surveys. The jet influence 

extends to approximately z/d = 2.4 at x/d = 1.5 for the thermocouple survey whereas CFD predicts the jet 

to extend to only z/d = 1.7. The thermocouple survey seems to indicate that the jet is mixing out rapidly 

and provides fairly constant cooling effectiveness as it travels downstream. The ‘dead zone’ of 

effectiveness is now larger and lies between x/d = 0.2 and x/d = 2.5. The coolant is securely carried 

within the coherent vertical structures and delivered downstream when the coherence is broken.

Figures 34(a) and (b) show contours of effectiveness at planes of constant x/d from CFD and 

thermocouple surveys, respectively.

Figures 35 and 36 show span-averaged and centerline effectiveness for M = 2.0 at DR = 1.0. The 

thermocouple measurements show a consistently higher effectiveness in the near hole region as the 

Kelvin-Helmholtz vortices are broken up and promote mixing. In the downstream region, the 

thermocouple results match well with IR. The TFNS result starts out over predicting effectiveness 

because as seen from Figure 12, the hairpin vortices propagate far downstream before they break up. 

Unlike the RANS result (Fig. 35) the TFNS-computed effectiveness follows the trend of the IR 

thermography data. The RANS simulation predicts complete jet lift-off with little mixing downstream. 

The pink line shows effectiveness from IR thermography with the turbulence grid installed.

Figure 31.—Film effectiveness from thermocouple surveys for M = 2.0, DR = 1.0, 

Tu = 1.5 percent. 
NASA/TM—2013-218086/PART1 31

Figure 32.—Isosurfaces of Q-criterion showing at M = 2.0, DR = 1.0, Tu = 1.5 percent.

Figure 33.—Centerline effectiveness from (a) CFD and (b) 

thermocouple at M = 2.0, DR = 1.0, Tu = 1.5 percent. 
NASA/TM—2013-218086/PART1 32

Figure 34.—Effectiveness contours from (a) CFD, (b) thermocouples at planes of 

constant x/d for M = 2.0, DR = 1.0, Tu = 1.5 percent.
NASA/TM—2013-218086/PART1 33

Figure 35.—Spanwise-averaged film effectiveness for M = 2.0, DR = 1.0.

Figure 36.—Centerline film effectiveness for M = 2.0, DR = 1.0. 
NASA/TM—2013-218086/PART1 34

Blowing Ratio = 2.5 

Figure 37 shows surface effectiveness for a blowing ratio of 2.5 from IR thermography for the low 

turbulence intensity case (Tu = 1.5 percent). The maximum effectiveness downstream of the hole has 

dropped to approximately 0.06 and occurs downstream of x/d = 5. Figures 38 and 39 show span-averaged 

and centerline effectiveness for M = 2.5 at DR = 1.0 from infrared thermography. 

Figure 37.—Film cooling effectiveness computed from infrared 

thermography for M = 2.5, DR = 1.0, Tu = 1.5 percent.

Figure 38.—Spanwise-averaged film effectiveness for M = 2.5, DR = 1.0, Tu = 1.5 percent.
NASA/TM—2013-218086/PART1 35

Figure 39.—Centerline film effectiveness for M = 2.5, DR = 1.0, Tu = 1.5 percent.

Comparison With Prior Work

This section compares the results from the present study to data found in the literature. In order to 

enable better comparison of the varying configurations, the momentum flux ratio, I, (the square of 

blowing ratio divided by density ratio) is used to compare cases with different density ratios. For cases 

where blowing ratio is not the same as for the present study, the blowing ratio is noted in addition to the 

momentum flux ratio. Figures 40 and 42 show centerline effectiveness for I = 0.25, and 1.0, respectively. 

Figures 41, 43, 44, and 45 show span-averaged effectiveness for I = 0.25, 1.0, 2.25, and 4.0, respectively. 

Unless otherwise specified, the angle of inclination of the holes to the freestream direction is 30°. The 

data presented here was extracted from online publications using a plot digitizer. 

Dhungel (Ref. 32) studied the effect of sister hole or ‘anit-vortex’ configurations and used round 

holes inclined at 30° as the baseline. The holes had P/D = 3.0 and L/D = 5.0. The Reynolds number based 

on hole diameter was 11000 and the freestream turbulence intensity was 2 percent. The holes were fed 

through a plenum. Johnson et al. (Ref. 33) used PIV and PSP to investigate film cooling from a row of 

circular holes inclined at 30° to the freestream. The hole spacing was 10 which is fairly large and is 

expected to lead to non-interacting coolant jets. The L/D was 6 and the turbulence intensity was under 

1 percent. Density ratios of 0.97, 1.0, and 1.53 and blowing ratios in the range 0.4 to 1.7 were studied. An 

attempt was made to collapse data based on density ratios, momentum ratios and blowing ratios. Lawson 

and Thole (Ref. 34) conducted an experiment to study the effect of deposition on film cooling 

effectiveness due to holes angled at 30°. Turbulence intensities of 4.6 and 12.3 percent were considered. 

Holes were spaced three hole diameters apart with L/D = 3.0. Blowing ratios of 0.49, 0.73, and 1.01 were 

studied with nominal density ratio of 1.06. Rallabandi et al. (Ref. 35) investigated the effect of an 

upstream step on flat plate effectiveness. They used simple angled, compound angled, cylindrical and fanshaped holes. Hole inclinations, α, of 30° and 45° were used with P/D = 3 and L/D = 7.5. The range of 

blowing ratios was 0.3 to 1.5 with unit density ratio. Turbulence intensity is estimated to be below 

0.5 percent. 
NASA/TM—2013-218086/PART1 36

Figure 40.—Centerline effectiveness for I = 0.25 (corresponds to DR = 1.0,

M = 0.5 in present study).

Figure 41.—Span-averaged effectiveness for I = 0.25 (corresponds to DR = 1.0, 

M = 0.5 in present study). 
NASA/TM—2013-218086/PART1 37

Figure 42.—Centerline effectiveness for I = 1.0 (corresponds to DR = 1.0, M = 1.0 

in present study). 

Figure 43.—Span-averaged effectiveness for I = 1.0 (corresponds to DR = 1.0,

M = 1.0 in present study).
NASA/TM—2013-218086/PART1 38

Figure 44.—Span-averaged effectiveness for I = 2.25 (corresponds to DR = 1.0, M = 1.5 in 

present study). 

Figure 45.—Span-averaged effectiveness for I = 4.0 (corresponds to DR = 1.0, M = 2.0 in 

present study). 

Conclusions

Infrared thermography of a flat plate cooled by long holes inclined at 30° to the freestream was 

conducted. The film effectiveness obtained from IR thermography was found to match well with 

thermocouple surveys and is thought to have less uncertainty than the thermocouple measurements. 

Kelvin Helmholtz structures were observed in the Time-Filtered Navier Stokes simulations that appeared 

to persist far downstream of the cooling hole. The breakup of these coherent structures gives improved 

effectiveness downstream of the hole especially for high blowing ratio. However, the delayed breakup 

predicted by CFD results in the predicted effectiveness being far higher than that shown by IR 

thermography. It also appears that thermocouple surveys are possibly influencing the rapid mixing of the 

jet at high blowing ratios causing a fairly constant level of span averaged effectiveness. Furthermore, the 

effectiveness computed by the thermocouples is at a height of z/d = 0.03 above the surface. Thus the 
NASA/TM—2013-218086/PART1 39

thermocouple yields consistently higher effectiveness than infrared thermography. The surface 

effectiveness predicted by CFD does follow the same trend as the effectiveness deduced from the infrared 

imaging. This could imply that modification of the filtering parameter or increased refinement 

downstream of the hole could yield more satisfactory results. It is also necessary to look at the turbulent 

stresses in the near-hole region to determine the reason for the coherence of the Kelvin-Helmholtz 

structures. Results from hotwire surveys will be analyzed and compared to the Reynolds stresses 

predicted by CFD. 

References

1. Bogard, D., and Thole, K., 2006, “Gas Turbine Film Cooling,” Journal of Propulsion and Power, 

22(2), 249–270, DOI:10.2514/1.18034. 

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

Single Row of Holes with Variable Density Ratio,” Journal of Turbomachinery, Vol. 113, pp. 442-

449. 

3. Pedersen, D.R., Eckert, E., and Goldstein, R., “Film Cooling with Large Density Differences Between 

the Mainstream and the Secondary Fluid Measured by the Heat-MassTransfer Analogy,” Journal of 

Heat Transfer, Vol. 99, Q18 1977, pp. 620–627. 

4. Bons, J.P., Rivir, R.B., Mac, A., Charles D., “The Effect of High Freestream Turbulence on Film 

Cooling Effectiveness,” WL-TR-96-2097, 13-16 June 1994.

5. Mayhew, J., Baughnb, J., Byerleyc, A., “The effect of freestream turbulence on film cooling adiabatic 

effectiveness,” International Journal of Heat and Fluid Flow, Volume 24, Issue 5, October 2003, 

pp. 669-679. 

6. Lutum, E., and Johnson, B.V., 1998, “Influence of the Hole Length to Diameter Ratio on Film 

Cooling With Cylindrical Holes,” ASME J. Turbomachinery, 121, pp. 209-216. 

7. Thurman, D.G., El-Gabry, L.A., Poinsatte, P.E., and Heidmann, J.D., 2011, “Turbulence and Heat 

Transfer Measurement in and Inclined Large Scale Film Cooling Array – Part II, Temperature and 

Heat Transfer Measurements,” ASME paper GT2011-46498, ASME Turbo Expo 2011, Vancouver.

8. Han, J.C. and Rallabandi, A.P., 2010, “Turbine Blade Film Cooling Using PSP Technique,” Frontiers 

Heat Mass Transfer, 1 (1), pp. 1–21. 

9. Papell, S.S., Graham R.W., Cageao, R.P., “Influence of coolant tube curvature on film cooling 

effectiveness as detected by infrared imagery,” NASA Technical Paper 1546, November 1979. 

10. Baldauf, S., A. Schulz, and S. Wittig. “High-resolution measurements of local effectiveness from 

discrete hole film cooling,” Journal of Turbomachinery Vol. 123.4, 2001, pp. 758-765. 

11. Ekkad, S. v., Ou, S., and Rivir B.R., “A Transient Infrared Thermography Method for Simultaneous 

Film Cooling Effectiveness and Heat Transfer Coefficient Measurements From a Single Test,” 

Journal of Turbomachinery, Vol. 126, October 2004, pp. 597-603. 

12. Wright, L., McClain, S., and Clemenson, M., “Effect of Density Ratio on Flat Plate Film Cooling 

With Shaped Holes Using PSP,” Journal of Tubomachinery, Vol. 133 October 2011, pp. 041011-1 – 

041011-11. 

13. Sgarzi, O., and Leboeuf, F., 1997, “Analysis of Vortices in Three-Dimensional Jets Introduced in a 

Cross-Flow Boundary Layer,” ASME Paper No. 97-GT- 517. 

14. El-Gabry, L., Heidmann, J. and Ameri, A.A., “Numerical Analysis of Fill Cooling at High Blowing 

Ratio,” NASA/TM—2009-215517. 

15. X. Guo, X., Schröder, W., Meinke, M., “Large-eddy simulations of film cooling flows,” Computers & 

Fluids 35 (2006) 587-606. 

16. Tyagi, M. and Acharya, S., 2003, “Large eddy simulation of film cooling flow from an inclined 

cylindrical jet,” ASME Journal of Turbomachinery, Vol. 125, pp. 734-742. 

17. Fujimoto, S. “Large Eddy Simulation of Film Cooling Flows Using Octree Hexahedral Meshes,” 

GT2012-70090 June 11-15, 2012, Copenhagen, Denmark.
NASA/TM—2013-218086/PART1 40

18. Y. Peet and S.K. Lele (2008) “Computational Framework for Coupling Compressible and Low Mach 

Number Codes,” AIAA Journal, vol. 46, No. 8, pp. 1990-2001. 

19. Wilcox, D. C., 1994, “Simulation of Transition With a Two-Equation Turbulence Model,” AIAA 

Journal, Vol. 32, No. 2, pp. 274-255. 

20. Steinthorsson, E., Liou, M.S., and Povinelli, L.A., 1993, “Development of an Explicit 

Multiblock/Multigrid Flow Solver for Viscous Flows in Complex Geometries,” AIAA-93-2380; also 

NASA TM-106356. 

21. Liu, N.-S., and Shih, T.-H., “Turbulence Modeling for Very Large-Eddy Simulation,” AIAA Journal,

Vol. 44, No. 4, 2006, pp. 687-697. 

22. Shih, T.-H., and Liu, N.-S., 2009, “Modeling of Internal Reacting Flows and External Static Stall 

Flows Using RANS and PRNS,” Flow, Turbulence and Combust (2008) 81:279-299. 

23. Shih, T.H., and Liu, N.S., A Nonlinear Dynamic Subscale Model for PRNS/VLES of Internal 

Combustor Flows, AIAA-2009-0467. 

24. Liu, Nan-Suey, Wey, Thomas, Shih, Tsan-Hsing, 2013, “Time-Filtered Navier-Stokes Approach and 

Emulation of Turbulence-Chemistry Interaction,” AIAA-2013-0707, 51st AIAA Aerospace Sciences 

Meeting Including the New Horizons Forum and Aerospace Exposition. January.

25. Davidson, L., and Peng, S.H., “Hybrid LES-RANS modeling: a one-equation SGS model combined 

with a k–ω model for predicting recalculating flows,” International Journal for Numerical Methods in 

Fluids, Vol. 43, 2003, pp. 1003-1018. 

26. Larsson, J., Lien, F.S. and Yee, E., “The artificial buffer layer and the effect of forcing in hybrid 

LES/RANS,” International Journal of Heart and Fluid Flow, Vol. 28, 2007, pp. 1443-1459. 

27. Spalart, P.R., Jou, W.H., and Allmaras, S.R., “Comments on the feasibility of LES for wings and on a 

hybrid RANS/LES approach,” Advances in DNS/LES, 1997, Greyden Press. 

28. Fasel, H.F., and von Terzi, D.A. and Sandberg, R.D., “A methodology for simulating compressible 

turbulent flows,” Journal of Applied Mechanics, Vol. 73, pp. 405-412. 

29. Hunt, J.C.R., Wray, A. & Moin, P. 1988, “Eddies, stream, and convergence zones in turbulent flows,” 

Center for Turbulence Research Report CTR-S88. 

30. Brown, A. and Saluja, C.L., “Film Cooling From a Single Hole and a Row of Holes of Variable Pitch 

to Diameter Ratio”. Int. J. Heat and Mass Transfer, Vol. 22, pp. 525-533, 1979. 

31. Dhungel, A., “Film Cooling From A Row Of Holes Supplemented With Anti Vortex Holes,” M.S. 

Thesis, Louisiana State University, 2007. 

32. Johnson, B., Zhang, K., Tian W. and Hu, Hui, “Experimental study of film cooling effectiveness by 

using PIV and PSP techniques,” AIAA 2013-0603, Proceedings of the AIAA Aerosciences Meeting, 

07-10 January 2013, Grapevine Texas. 

33. Lawson, S.A., and K.A. Thole. “Effects of simulated particle deposition on film cooling,” Journal of 

Turbomachinery 133.2 (2011): 021009. 

34. Rallabandi, A.P., Grizzle, J., Han, J-C. 2011. “Effect of Upstream Step on Flat Plate Film-Cooling 

Effectiveness Using PSP,” Journal of Turbomachinery. Vol. 133, pp. 041024-1–8. 

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Long Hole Film Cooling Dataset for CFD Development 

Part 1: Infrared Thermography and Thermocouple Surveys

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6. AUTHOR(S)

Shyam, Vikram; Thurman, Douglas; Poinsatte, Phillip; Ameri, Ali; Eichele, Peter; Knight, 

James 

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WBS 473452.02.03.05.04.02.03 

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14. ABSTRACT

An experiment investigating flow and heat transfer of long (length to diameter ratio of 18) cylindrical film cooling holes has been 

completed. In this paper, the thermal field in the flow and on the surface of the film cooled flat plate is presented for nominal freestream 

turbulence intensities of 1.5 and 8 percent. The holes are inclined at 30° above the downstream direction, injecting chilled air of density ratio 

1.0 onto the surface of a flat plate. The diameter of the hole is 0.75 in. (0.01905 m) with center to center spacing (pitch) of 3 hole diameters. 

Coolant was injected into the mainstream flow at nominal blowing ratios of 0.5, 1.0, 1.5, and 2.0. The Reynolds number of the freestream 

was approximately 11,000 based on hole diameter. Thermocouple surveys were used to characterize the thermal field. Infrared thermography was used to determine the adiabatic film effectiveness on the plate. Hotwire anemometry was used to provide flowfield physics and 

turbulence measurements. The results are compared to existing data in the literature. The aim of this work is to produce a benchmark dataset 

for Computational Fluid Dynamics (CFD) development to eliminate the effects of hole length to diameter ratio and to improve resolution in 

the near-hole region. In this report, a Time-Filtered Navier Stokes (TFNS), also known as Partially Resolved Navier Stokes (PRNS), method 

that was implemented in the Glenn-HT code is used to model coolant-mainstream interaction. This method is a high fidelity unsteady 

method that aims to represent large scale flow features and mixing more accurately. 

15. SUBJECT TERMS

Infrared imagery; Nondestructive tests; Temperature distribution; Temperature measurement; Thermal mapping; Film cooling; 

Turbomachinery 

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