G. James Van Fossen

Glenn Research Center, Cleveland, Ohio

Ronald S. Bunker

General Electric, Schenectady, New York

Augmentation of Stagnation Region Heat

Transfer Due to Turbulence From

a DLN Can Combustor

NASA/TM—2000-210241

July 2000

ASME 2000–GT–0215
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Glenn Research Center

Prepared for the

45th International Gas Turbine and Aeroengine Technical Congress

sponsored by the American Society of Mechanical Engineers

Munich, Germany, May 8–11, 2000

G. James Van Fossen

Glenn Research Center, Cleveland, Ohio

Ronald S. Bunker

General Electric, Schenectady, New York

NASA/TM—2000-210241 ASME 2000–GT–0215

Augmentation of Stagnation Region Heat

Transfer Due to Turbulence From

a DLN Can Combustor
Available from

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NASA/TM—2000-210241 1 

AUGMENTATION OF STAGNATION REGION HEAT TRANSFER DUE TO 

TURBULENCE FROM A DLN CAN COMBUSTOR

G. James Van Fossen 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio

Ronald S. Bunker 

General Electric 

Schenectady, New York

ABSTRACT 

Heat transfer measurements have been made in the 

stagnation region of a flat plate with a circular leading 

edge. Electrically heated aluminum strips placed 

symmetrically about the leading edge stagnation 

region were used to measure spanwise averaged heat 

transfer coefficients. The maximum Reynolds number 

obtained, based on leading edge diameter, was about 

100,000. The model was immersed in the flow field 

downstream of an approximately half scale model of a 

can-type combustor from a low NOx, ground based 

power-generating turbine. The tests were conducted 

with room temperature air; no fuel was added. Room 

air flowed into the combustor through six vane type 

fuel/air swirlers. The combustor can contained no 

dilution holes. 

The fuel/air swirlers all swirled the incoming 

airflow in a counter clockwise direction (facing 

downstream). A 5-hole probe flow field survey in the 

plane of the model stagnation point showed the flow 

was one big vortex with flow angles up to 36o

 at the 

outer edges of the rectangular test section. Hot wire 

measurements showed test section flow had very high 

levels of turbulence, around 28.5%, and had a 

relatively large axial-length scale-to-leading edge 

diameter ratio of 0.5. X-wire measurements showed 

the turbulence to be nearly isotropic. Stagnation heat 

transfer augmentation over laminar levels was around 

77% and was about 14% higher than predicted by a 

previously developed correlation for isotropic grid 

generated turbulence. 

NOMENCLATURE 

B wind tunnel blockage 

Cτ constant in fit of autocorrelation, sec.-1 

d leading edge diameter, cm 

E hot wire bridge voltage, volts 

e fluctuating component of bridge voltage, volts 

Fr Frossling number = Nusselt / (Re)1/2

Pr Prandtl number

R radius of model leading edge, cm 

r recovery factor 

R(τ) autocorrelation function 

Re Reynolds number 

s surface distance from stagnation point, cm 

T temperature, K 

Tu turbulence intensity 

U, V axial and cross-stream components of velocity, 

m/sec 

u', v' fluctuation components of U & V, m/sec 

Greek

Λ Length scale, cm 

κ wave number, m-1 

ρ Density, kg/m3 

τ Delay time in autocorrelation, sec. 

Subscripts

avg averaged over time 

d leading edge diameter 

RMS root mean square 

r recovery 

st static condition 

t total condition 

x axial direction 

∞ free stream condition 

INTRODUCTION

Firing temperatures of modern turbine engines 

often exceed the melting temperature of the 

downstream blade and vane alloy material thus 

requiring the hardware to be cooled by compressor 

discharge air or some other means. The design of these 

cooling systems requires a knowledge of the heat 

transfer coefficients between the hot gas and the blade 

or vane. Turbulence in the combustor exit stream plays 

a major role in determining the level of heat transfer. 

The stagnation region is an area of high heat transfer 

and is very sensitive to the level of freestream 

turbulence. The objective of the present work is to 

determine the level of stagnation region heat transfer 

augmentation due to turbulence produced by a scaled, 

can-type combustor from a ground-based, gas-fired, 
NASA/TM—2000-210241 2 

turbine operating at room temperature and compare the 

results with heat transfer augmentation from isotropic 

grid-generated turbulence. 

It has been known for a long time that a small 

increase in freestream turbulence from laminar levels 

can cause a large increase in heat transfer from the 

stagnation region of a bluff body such as a turbine 

vane or blade, Zapp (1950), Giedt (1951). Early 

attempts at correlation, e.g. Smith & Kuethe (1966) 

and Lowery & Vachon (1975), ignored the effect of 

turbulent length scale and tried to correlate heat 

transfer augmentation with only turbulence intensity 

and Reynolds number. These attempts resulted in 

large, unexplained scatter when compared to the data 

of other researchers. The works of Ames (1994) and 

Van Fossen, et al. (1994) showed that length scale was 

also an important parameter. 

Van Fossen, et al. (1994) studied the 

augmentation of stagnation region heat transfer due to 

isotropic turbulence from square-bar, square-mesh 

grids. A correlating parameter was developed from 

experimental data taken with several different grids 

which grouped the augmentation data very tightly, 

(±4%). The isotropic data of several other studies were 

compared to the correlation and found to agree within 

±10%. Data from an array of parallel, fine wires that 

generated non-isotropic turbulence were correlated by 

the parameter but showed a much higher level of 

augmentation than the isotropic turbulence. 

Ames (1994) fabricated turbulence generators that 

resembled a combustor liner having primary air inlets 

and dilution holes. Using the rapid distortion theory of 

Hunt and Graham (1978), he developed a correlating 

parameter based on intensity, Reynolds number and 

what he called an energy-scale, the average size of the 

energy containing eddies. The correlation fit his and 

other data quite well. 

Results to be presented from the present work 

include heat transfer measurements in the stagnation 

region of a flat plate with a circular leading edge, a 

5-hole probe flowfield survey, turbulence intensity 

survey using a single hot wire, and X-wire results 

downstream of the simulated combustor liner. 

TEST FACILITY, INSTRUMENTATION AND 

DATA ACQUISITION

Wind Tunnel—The experiments were carried out in 

the wind tunnel shown in Fig. 1. 

Air drawn from the test cell passed through the 

six-scaled combustor fuel/air swirlers into the 

32.4-cm diameter by 26.4-cm long can, through a 

constant area, 68.5-cm long, transition section and into 

the 15.2-cm by 54.1-cm by 77.5-cm long test section. 

Air leaving the test section then passed through 

another transition section to a 10-inch pipe, which 

contained an orifice for flow measurement and a 

butterfly valve for flow control. Flow through the 

swirlers could be choked so the control valve was set 

to supply the maximum test section pressure that gave 

the maximum flow rate. The ratio of test section static 

pressure to test cell barometric pressure was typically 

0.63. This gave a leading edge Reynolds number of 

about 100,000. A Kiel probe located in the test section 

was used to measure the total pressure for each run. 

Four chromel-alumel thermocouples near the inlet of 

the swirlers measured the total inlet temperature. A 2-

dimensional actuator system allowed hot wire and 5-

hole probe surveys in a plane perpendicular to the flow 

that passed through the stagnation line of the heat 

transfer model. The model was not present when the 

flow surveys were conducted. 

Figure 1.—Rig layout. 

Simulated Combustor—The combustor liner 

model is of the natural gas fired, Dry-Low-NOx 

(DLN), can-type having no dilution or film cooling air 

injection. Six identical, scaled DLN-type fuel-air 

swirlers are set into the upstream end, or headend, of 

the combustor liner. Each swirler, shown in fig. 2, was 

composed of a set of 12 large air swirl vanes (outer) 

and a 16-vane diffusion gas tip (inner). 

Figure 2.—DLN fuel-air swirler looking upstream. 

The air swirler vane sets were cast stainless steel 

with about 45-degrees of turning. The diffusion gas 

tips were fabricated from aluminum for this mockup. 

Swirl is in the same direction for each of the inner and 

outer vanes. Each swirler is also provided with a 

hardware. In this mockup, the headend has no 

additional flow orifices other than the swirlers. This 

configuration results in a cold flow combustor 
NASA/TM—2000-210241 3 

mockup, generating both swirl and large-scale 

turbulence typical of a land-based power turbine. 

While this cold flow combustor mockup does not 

achieve the pressure and temperature levels of an 

actual engine, nor does it have the reaction and 

combustion chemistry, previous work of Zimmerman 

(1979) and Geobel et al. (1993) indicates that the 

inclusion of the major geometry and flow features in 

such a cold flow device does result in the generation of 

freestream turbulence and swirl representative of a 

pressurized and fired system. 

Five Hole Pressure Probe—Mean flow angle and 

velocity were measured with a standard, commercially 

available, 5-hole pressure probe. Pressures were 

measured with strain gage type transducers and 

recorded using the Laboratory steady state data 

recording system. 

Hot wire instrumentation—Turbulence measurements were obtained using a 2-channel constant 

temperature anemometer system. The turbulence 

intensity survey in the plane of the model stagnation 

point was conducted with a single hot wire oriented 

parallel to the model spanwise direction with the 

model removed. Bridge voltage was converted to mean 

and RMS components with special purpose analog 

meters whose output was recorded with the Laboratory 

steady state data acquisition system. Length scale and 

isotropy measurements were made with a cross-flow 

X-wire probe. The wires were oriented perpendicular 

to the model spanwise direction. Bridge voltage was 

passed through an 8 pole, 6 zero, constant delay antialiasing filter set to begin roll off at 6 kHz. 220 samples 

were recorded at an 18 kHz sampling frequency with a 

high-speed, 12-bit, digital to analog conversion (DAC) 

board and stored in a personal computer. In order to 

obtain good resolution of the fluctuating component of 

the signal it was necessary to subtract a known DC 

voltage from the raw signal then amplify the result 

before passing it to the DAC. For all hot wire 

measurements the frequency response was estimated to 

be about 90 kHz using the standard square wave test. 

Heat Transfer Model—The heat transfer model 

was a 65-cm long flat plate with a circular leading 

edge with a radius of 3.30-cm. The model spanned the 

15.2-cm test section width. A 52-cm long wedge 

shaped afterbody designed to eliminate vortex 

shedding extended downstream into the outlet 

transition section as shown in Fig. 1. The model 

contained 19 heat flux gages consisting of aluminum 

strips 6.6-cm long by 0.48-cm wide and 0.32-cm deep 

with a thin foil electric heater glued to the rear and a 

chromel-alumel thermocouple embedded in a groove. 

Each gage averaged the heat transfer over its 6.6-cm 

span. The farthest downstream gages were used as 

guard heaters to minimize losses in the streamwise 

direction. A guard heater behind the gages prevented 

conduction to the interior of the model. The average 

gap between the aluminum strips was 0.25 mm and 

was filled with epoxy. The aluminum strips were 

maintained at a constant temperature (±0.25 o

C) by a 

specially designed control circuit (Van Fossen et al. 

1984). 

DATA REDUCTION 

Flow Field—The 5-hole probe was calibrated in an 

atmospheric jet over a range of pitch and yaw angles 

using the technique described in Giel et al. (1996). 

Yaw and pitch angles were least square curve fit with 

fourth order polynomials. Symmetry was not assumed 

in any of the curve fits. 

Turbulence

Intensity—Hot wire probes were calibrated in 

an atmospheric jet. A fourth order polynomial was 

used to fit bridge voltage versus density-velocity 

product. Density-velocity product was used because 

the test section pressures were below atmospheric thus 

calibration density was different than test section 

density. The mean and RMS voltages recorded for the 

intensity survey with the steady state recording system 

were converted to density velocity product and 

turbulence intensity was calculated as 

Where E is the bridge voltage, eRMS is the RMS 

component of the bridge voltage and the derivative is 

the slope of the calibration curve at E. 

 Integral Length Scale—Length scale was 

calculated from the autocorrelation function by fitting 

an exponential curve 

between the arbitrary limits of 1>R(t)>0.33. 

Integrating R(t) between zero and infinity and 

multiplying by the mean velocity gives the integral 

length scale, 

∫ (3) 

∞

Λ = =

0

( ) Cτ

U U R t dt x

τ

τ τ C R e− ( ) =

U

e

dE

d U

Tu

RMS

E

ρ

ρ





 





=

( )

(1) 

(2) 
NASA/TM—2000-210241 4 

Although not a problem with this data set, this method 

of computing the length scale eliminates inconsistent 

results from some data which contain low frequency 

noise, see Van Fossen et al. (1994). 

Isotropy—The X-wire probe was calibrated in 

the same air jet used to calibrate the single wire probe. 

The procedure was to set the probe angle so both wires 

were as close as possible 45o

 to the flow and perform a 

velocity calibration as was done for the single wire. 

This produced a curve fit of bridge voltage versus jet 

velocity for each wire. A directional calibration was 

accomplished by setting the jet density-velocity 

product at the value expected in the test section and 

recording bridge voltage versus jet velocity over a 

range of probe angles between –40o

 and +40o. Using 

the velocity curve fit for each wire, a density-velocity 

product was calculated for each probe angle and 

normalized by the actual jet density-velocity product. 

These data sets were then curve fit with polynomials. 

Figure 3 shows the resulting calibration curves with 

the data points superimposed. 

It can be seen in the directional calibration that 

wire 2, the outer wire, looks like a cosine curve—

normal behavior for an angled hot wire. Wire 1 

however looks like the cosine for angles between –40o

and 0o

 but at angles greater than 0o does not. This is a 

problem inherent to the cross-flow probe design; at 

angles greater than 0o

 wire 1 is in the wake of the 

prongs that support wire 2. Several features of this 

type of probe outweigh this inconvenience: 1) the wire 

angle can be changed by simply rotating the probe 

stem without a change in the streamwise or crossstream position, 2) the probe can be inserted through a 

small (6.4-mm) hole in an actuator or tunnel wall 

without having other access to the test section such as 

a window or door to install the probe tip. This 

shortcoming of the crossflow probe can be overcome 

by limiting the flow angles relative to the probe to the 

range –40o

 to 0o. The curve fits from fig. 3 were used 

to construct a lookup table of the ratio ρV1/ρV2 versus 

probe angle. The data reduction procedure was then: 

1) calculate ρV1 and ρV2 using the curve fits of 

velocity versus bridge voltage. 2) using the ratio of 

ρV1/ρV2 and the lookup table, find the angle of the 

velocity vector relative to the probe, 3) using the curve 

fit of ρV2/ρVjet versus angle, compute ρVjet (ρVjet is 

now the magnitude of the mass-velocity vector). 

Heat Transfer

An energy balance was used to determine the 

Frossling number for each gage: 

Probe Angle

ρVn / ρVjet

-40 -20 0 20 40 0.0

0.4

0.8

1.2

1.6

Inner Wire, V1

Outer Wire, V2

Fit

ρVjet, kg / m2

-sec

Bridge Voltage

0 20 40 60 80 100 1.0

1.5

2.0

2.5

Inner Wire, V1

Outer Wire, V2

Fit

Figure 3.—Cross-flow X-wire calibration curves. 

Where qEI is the heat added to the gage by the electric 

heater, qrad is the heat lost by radiation, qgap is the heat 

conducted away from the gage to the epoxy gap and 

the unguarded ends of the heaters, A is the exposed 

gage surface area, Tw is the gage temperature, Tr(s) is 

the local recovery temperature, k is the thermal 

conductivity of air, and Red is the Reynolds number 

based on leading edge diameter, d. Corrections for 

radiation heat loss, qrad, were made assuming gray 

body radiation to black surroundings and an emissivity 

of 0.05 for the aluminum gage. 

An estimate of the gap loss, qgap, can be obtained 

from an exact solution for two-dimensional heat 

conduction in a rectangle. See Van Fossen, et al. 

(1984) for details. The gap loss was about 10 percent 

of the total heat flow while the radiation heat loss was 

on the order of 0.2 percent. 

The recovery temperature was calculated as 

Where Tst,: is the static temperature upstream of 

the model. 

( ) ( )( ) r = st,∞ + Tt − Tst,∞ T s T r s (5) 

( ) [ ] W r d

EI rad gap

A T T s k

q q q d

Fr s R

( ) Re

( ) / −− − = (4) 
NASA/TM—2000-210241 5 

The local recovery factor, r(s), was calculated as 

The local mass flow ratio, ρU(s)/(ρU):, was found 

from a numerical solution of flow over the model that 

included the tunnel walls (Rigby et al., 1992). 

Following the results of Rigby et al. (1992), the 

thermal conductivity and viscosity of air were 

evaluated at the free-stream total temperature from 

equations given in Hillsenrath et al. (1955). 

The Reynolds number, Red, was based on the 

diameter of the leading edge, d, and the mass-velocity 

averaged between the flow area with maximum model 

blockage and the unblocked flow area, that is, 

Where the blockage, B, is the ratio of maximum model 

thickness to tunnel height. Blockage was 0.122 for this 

test. 

UNCERTAINTY ANALYSIS 

 

Frossling number—Estimates of the bias error of 

each measuring instrument were made and combined 

by the method of Kline and McClintock (1953). 

Estimates of the precision of each measurement were 

calculated from 20 samples of each steady-state 

measurement and combined by the same method. The 

maximum Frossling number uncertainty occurred at 

the stagnation gage and was 5.7%. 

Flowfield measurements—Estimates of error in 

the hot wire and 5-hole probe measurements were 

made using a method suggested by Yavuzkurt (1984). 

Estimated error in turbulence intensity was 12.9% and 

in length scale was 18.3%. 

RESULTS AND DISCUSSION 

Flow Field—Results of the 5-hole probe survey are 

shown in fig. 4. Shown on the figure are streamlines 

and vectors that represent the components of velocity 

perpendicular to the axial direction. The length of the 

vector at the top of the figure is representative of the 

magnitude of the axial velocity. It can be seen that 

vortices from the six individual swirl nozzles add 

together to form a single vortex at the inlet to the test 

section. The center of the vortex is displaced from the 

tunnel centerline. Flow angles of up to 36o

 were 

measured at the extreme spanwise positions and angles 

near 0o

 were recorded at midspan. 

CROSS-STREAM, cm

SPANWISE, cm

-8 -6 -4 -2 0 2 4 6 8

-6

-4

-2

0

2

4

6

REFERENCE VECTOR

AVERAGE AXIAL VELOCITY =

Figure 4.—Five-hole probe survey; cross-stream vector 

and streamline plot. 

Figure 5 shows the mean flow velocity 

components in the axial and cross-stream directions 

measured with the X-wire. 

The axial component varies about 17% in the 

spanwise direction with a minimum near mid-span. 

The cross-stream component varies linearly with span 

with zero velocity corresponding with the mid-span 

position. 

Spanwise, cm

U, m/sec

V, m/sec

-6 -4 -2 0 2 4 6 0

10

20

30

40

50

60

-30

-20

-10

0

10

20

043099 30

043099

050499

052599

060499a

060499b

070299

U

V

Figure 5.—X-wire traverse: mean axial flow and 

cross-stream velocity components. 

2(1 )

(2 ) ( ) ( ) B

B U avg U −

− ρ = ρ ∞ (7) 

(6) (1 Pr) ( )

( ) ( ) 1

2

− 













 = − U ∞

U s

r s

ρ

ρ
NASA/TM—2000-210241 6 

Turbulence—A survey of turbulence intensity 

measured with the single hot wire is shown in fig. 6. 

Intensity was stratified in the spanwise direction with a 

maximum of around 30% at mid-span and tapering off 

to around 21% at the full test section span. Intensity 

averaged over the span of the stagnation heat transfer 

gage was 28.5%. 

22

21

22

22

23

23

23

23

24

24

24

25 25

24

25

26 27

26

26

26

27 27

27

27

27

27

28

28

28

28

28

28

28

27

28

28

29

29

29

29

29

29

30

29

30

30

28

29

30 23

25

27

29

30

28

26

27

29

24

28

28

30

29

29

30 30

CROSS-STREAM, cm

SPANWISE, cm

-8 -6 -4 -2 0 2 4 6 8

-6

-4

-2

0

2

4

6 TURBULENCE INTENSITY IN PERCENT

Figure 6.—Turbulence intensity contour plot. 

Figure 7 gives the fluctuating components of axial and 

cross-stream velocity. The axial component is nearly 

constant while the cross-stream component has a dip 

near mid-span. This, at first, appears contrary to the 

results of the single wire survey but closer examination 

shows that it is consistent. The single wire was 

oriented in the spanwise direction and thus would 

respond to axial and cross-stream components as: 

The zero velocity of the cross-stream mean 

component, V, causes the intensity to peak in the 

center of the span. 

Figure 8 shows the ratio of the two fluctuating 

velocities, u’/v’. The average ratio over the span is 

equal to 0.97 and varies only about ±10% indicating 

that the flow is nearly isotropic. 

Typical autocorrelation functions used to 

determine the length scale are shown in fig. 9. Also 

shown on the figure are the exponential curve fits used 

to determine the length scale. These figures are typical 

of autocorrelations for all the data taken. The axial 

direction autocorrelation is closely approximated by 

the exponential curve fit while the cross-stream 

component has a large area of negative correlation. 

Spanwise, cm

u', m/sec

v', m/sec

-6 -4 -2 0 2 4 6 4

5

6

7

8

9

7

8

9

10

11

12

043099

050499

052599

060499a

060499b

070299

u'

v'

Figure 7.—X-wire traverse: RMS fluctuating axial and 

cross-stream velocity components. 

Spanwise, cm

u'/v'

-6 -4 -2 0 2 4 6 0.7

0.8

0.9

1.0

1.1

1.2

1.3

043099

050499

052599

060499a

060499b

070299

Figure 8. X-wire traverse: ratio of RMS fluctuating 

axial and cross-stream velocity components. 

The spanwise distribution of length scale is shown 

on fig. 10. The average axial length scale was 3.34-cm 

while the cross-stream length scale was 2.64-cm. 

making the ratio of axial to cross-stream length scale 

about 1.26. 

Axial and cross-stream components of the power 

spectrum in wave number space are shown in fig. 11. 

Both components exhibit the –5/3 slope indicative of 

the inertial subrange. The rolloff in the spectrum at 

wave numbers greater than about 1500 is caused by the 

anti-aliasing filter. The spectra shown in fig. 11 are 

typical of all spanwise locations. 

( )

( ' ' ) 2 2

2 2

U V

u v Tu

+

+ = (8) 
NASA/TM—2000-210241 7 

τ, sec.

Rx(τ)R(τ)y

0 0.01 0.02 0.03

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

1.8

2.0

2.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Data

Fit

R(τ)x

R(τ)y

Figure 9.—Typical axial and cross-stream autocorrelation 

functions with curve fits. 

Spanwise, cm

Λx, cmΛy, cm

-6 -4 -2 0 2 4 6 0

1

2

3

4

5

1

2

3

4

5

6 043099

050499

052599

060499a

060499b

062599

Λ 070299 x

Λy

Figure 10.—X-wire traverse: axial ancd cross-stream 

length scales. 

κ=2πf / U, m-1

Ex(κ)=UEx(f)/2π, m3/sec2Ey

(κ)=UEy(f)/2π, m3/sec2

10-1 100 101 102 103 104 10-6

10-5

10-4

10-3

10-2

10-1

100

101

10-5

10-4

10-3

10-2

10-1

100

101

102

Ex(κ)

Ey(κ)

-5/3 SLOPE

Figure 11.—Typical axial and cross-stream 

wave number spectrum. 

Heat Transfer—The heat transfer distribution in the 

stagnation region is shown in fig. 12 where Frossling 

number (Nusselt number/(Red)

1/2) is plotted versus the 

ratio of surface distance to leading edge radius. For 

comparison, a numerical solution for the cylindrical 

leading edge for a laminar free stream obtained using 

the PARC-2D code taken from Rigby and Van Fossen 

(1992) is shown. 

Also shown are the PARC-2D code results 

multiplied by the factor 1.77 in order to match the 

experimental data at the stagnation point. The 

augmentation of stagnation heat transfer due to the 

DLN turbulence is about 77% above the laminar case. 

Frossling numbers for heat flux gages off the 

stagnation line generally follow the laminar 

distribution multiplied by the augmentation factor until 

s/R of roughly ±0.9 (~±52o

). Beyond this value of s/R, 

the augmentation is greater than the 1.77 of the 

stagnation region. 

Figure 13 shows a comparison of the DNL 

turbulence augmented stagnation heat transfer 

compared with a correlation for stagnation region heat 

transfer taken from Van Fossen et al. (1995). 

The stagnation heat transfer due to the DLN 

turbulence is about 14% higher than the correlation 

predicts. The correlation was derived using turbulence 

generated with square-bar, square-mesh grids that 

produced nearly isotropic turbulence; the data was all 

contained within the ±4% bands shown in fig. 13. Van 

Fossen et al. (1995) showed that non-isotropic 

turbulence could produce augmentation levels higher 

than that of isotropic turbulence. It was suspected that 

this was the cause of the increased augmentation by 

the DLN combustor turbulence but the turbulence 

measurements show that this is not the case; the DLN 

turbulence in nearly isotropic. The cause of this 

increased augmentation above isotropic grid 

turbulence levels remains unknown. One possibility 

could be the swirl in the mean flow causing an 

effective angle of attack on the leading edge; positive 

on one half of the span and negative on the other half. 

This has not been investigated. 
NASA/TM—2000-210241 8 

s/R

Fr(s/R)

-1.5 -1.0 -0.5 0.0 0.5 1.0 1.5

0.4

0.6

0.8

1.0

1.2

1.4

1.6

1.8

Red=96400

PARC-2D

PARC-2D*1.77

Figure 12.—Heat transfer distribution over the 

model leading edge. 

TuRe0.8 (Λx / d)-0.574

Fr(0)Tu / Fr(0)Lam

0 5000 10000 15000 0.5

1.0

1.5

2.0

2.5

Isotropic Grid Turbulence

Range of Data Scatter (±4%)

DLN Turbulence

Figure 13.—Comparison of stagnation heat transfer 

augmentation due to DLN turbulence and correlation 

derived from square bar grid data. 

CONCLUSIONS 

Spanwise averaged heat transfer measurements 

were made in the stagnation region of a flat plate with 

a 3.3-cm radius circular leading edge. The heat transfer 

model was placed downstream of a scaled, simulated, 

gas fired, combustor liner from a ground based, power 

turbine. Conclusions drawn from this study are: 

1. The DLN combustor used in the present test 

with the swirl vanes from each of the six 

nozzles oriented in the same direction 

produced a single large swirl entering the test 

section. Flow angles up to 36o

 from axial were 

recorded at the outer edge while flow angles of 

nearly zero were found at the center. 

2. Very high levels of free stream turbulence 

were generated. Spanwise averaged turbulence 

intensity was found to be 28.5%. 

3. The turbulence generated appeared to be 

isotropic. Fluctuating components of velocity 

in the axial and cross-stream direction were 

nearly identical in magnitude. Wave number 

spectrum for both components demonstrated 

the –5/3-slope characteristic of an inertial 

subrange. 

4. Average length scale in the axial direction 

was found to be 3.34-cm while the crossstream direction length scale was 2.64-cm. 

5. Stagnation heat transfer in the turbulent 

exhaust of the DLN combustor liner was 

found to be higher than that for a laminar free 

stream by a factor of 1.77. 

6. Stagnation heat transfer augmentation was 

found to be 14% higher than predicted by a 

correlation developed from isotropic grid 

generated turbulence. 

REFERENCES 

Ames, F.E. (1994), “Experimental Study of Vane Heat 

Transfer and Aerodynamics at Elevated levels of 

Turbulence,” NASA CR–4633. 

Giedt, W.H., (1951), “Effect of Turbulence Level of 

Incident Air Stream on Local Heat Transfer From 

Cylinders,” J. of Aeronaut. Sci., vol. 18, no.11, 

pp. 725–730. 

Giel, P.W., Thurman, D.R., Lopez, I., Boyle, R.J., 

Van Fossen, G.J., (1966), “Three-Dimensional Flow 

Field Measurements in a Transonic Turbine Cascade,”

ASME Paper 96–GT–113. 

Gobel, S., Abuaf, N., Lee, C.-P., (1993), 

“Measurements of Combustor Velocity and 

Turbulence Profiles,” ASME Paper 93–GT–228. 

Hillsenrath, J., Beckett, C.W., Benedict, W.S., 

Fano, L., and Hobe, H.J., (1955), “Tables of Thermal 

Properties of Gases,” NBS Circular 564. 

Hunt, J.C.R., and Graham, J.M.R., (1978), “FreeStream Turbulence Near Plane Boundaries,” J.Fluid 

Mech., vol. 84, pp. 209–235. 

Kline, S.J., and McClintock, F.A., (1953), “Describing 

Uncertainties in Single-Sample Experiments,” Mech. 

Engineering, vol. 75, pp. 3–8. 

Lowery, G.W., and Vachon, R.I., (1975), “Effect of 

Turbulence on Heat Transfer From Heated Cylinders,”

Int. J. of Heat Mass Trans., vol. 18, no. 11, 

pp. 1229–1242. 
NASA/TM—2000-210241 9 

Rigby, D.L., Van Fossen, G.J., (1992), “Increased 

Heat Transfer to Elliptical Leading Edges Due to 

Spanwise Variations in the Freestream Momentum: 

Numerical and Experimental Results,” AIAA paper 

92–3070. 

Smith, M.C., Kuethe, A.M., (1966), “Effects of 

Turbulence on Laminar Skin Friction and Heat 

Transfer,” Phy. of Fluids, vol. 9, no. 12, pp. 2337–

2344. 

Van Fossen, G.J., Simoneau, R.J., Olsen, W.A., 

Shaw, R.J., (1984), “Heat Transfer Distributions 

Around Nominal Ice Accretion Shapes Formed on a 

Cylinder in the NASA Lewis Icing Research Tunnel,”

Presented at the AIAA 22nd Aerospace Sciences Mtg., 

Jan. 9–12, Reno, NV (also NASA TM–83557). 

Van Fossen, G.J., Simoneau, R.J., Ching, C.Y., (1994), 

“Influence of Turbulence Parameters, Reynolds 

Number, and Body Shape on Stagnation Region Heat 

Transfer, NASA TP–3487. 

Van Fossen G.J., Simoneau, R.J., Ching, C.Y., (1995), 

“Influence of Turbulence Parameters, Reynolds 

Number, and Body Shape on Stagnation-Region Heat 

Transfer,” J. of Heat Transfer, vol. 117, pp. 597–603. 

Yavuskurt, S., (1984), “A Guide to Uncertainty 

Analysis of Hot-Wire Data,” J. of Fluids Eng., 

vol. 106, pp. 181–186. 

Zapp, G.M., (1950), "The Effect of Turbulence on 

Local Heat Transfer Coefficient Around a Cylinder 

Normal to an Air Stream," M.S. Thesis, Oregon State 

College, Corvalis, OR. 

Zimmerman, D.R., (1979), “Laser Anemometer 

Measurements at the Exit of a T63-C20 Combustor,”

NASA CR–159623. 
This publication is available from the NASA Center for AeroSpace Information, (301) 621–0390.

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NASA TM—2000-210241

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G. James Van Fossen and Ronald S. Bunker

Heat transfer; Gas turbine; Turbulence

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Prepared for the 45th International Gas Turbine and Aeroengine Technical Congress sponsored by the American Society

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Augmentation of Stagnation Region Heat Transfer Due to Turbulence From

a DLN Can Combustor

Heat transfer measurements have been made in the stagnation region of a flat plate with a circular leading edge.

Electrically heated aluminum strips placed symmetrically about the leading edge stagnation region were used to

measure spanwise averaged heat transfer coefficients. The maximum Reynolds number obtained, based on leading edge

diameter, was about 100,000. The model was immersed in the flow field downstream of an approximately half scale

model of a can-type combustor from a low NOx, ground based power-generating turbine. The tests were conducted with

room temperature air; no fuel was added. Room air flowed into the combustor through six vane type fuel/air swirlers.

The combustor can contained no dilution holes. The fuel/air swirlers all swirled the incoming airflow in a counter

clockwise direction (facing downstream). A 5-hole probe flow field survey in the plane of the model stagnation point

showed the flow was one big vortex with flow angles up to 36° at the outer edges of the rectangular test section. Hot

wire measurements showed test section flow had very high levels of turbulence, around 28.5 percent, and had a relatively large axial-length scale-to-leading edge diameter ratio of 0.5. X-wire measurements showed the turbulence to be

nearly isotropic. Stagnation heat transfer augmentation over laminar levels was around 77 percent and was about 14

percent higher than predicted by a previously developed correlation for isotropic grid generated turbulence.