John D. Wolter

Glenn Research Center, Cleveland, Ohio

Robert Childs

ELORET Corporation, Moffett Field, California

Mark P. Wernet

Glenn Research Center, Cleveland, Ohio

Andrea Shestopalov

ELORET Corporation, Moffett Field, California

John E. Melton

Ames Research Center, Moffett Field, California 

Flow Field Characterization of an Angled 

Supersonic Jet Near a Bluff Body

NASA/TM—2011-216958

March 2011

AIAA–2010–4829
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John D. Wolter

Glenn Research Center, Cleveland, Ohio

Robert Childs

ELORET Corporation, Moffett Field, California

Mark P. Wernet

Glenn Research Center, Cleveland, Ohio

Andrea Shestopalov

ELORET Corporation, Moffett Field, California

John E. Melton

Ames Research Center, Moffett Field, California 

Flow Field Characterization of an Angled 

Supersonic Jet Near a Bluff Body

NASA/TM—2011-216958

March 2011

AIAA–2010–4829

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

28th Applied Aerodynamics Conference

sponsored by the American Institute of Aeronautics and Astronautics

Chicago, Illinois, June 28–July 1, 2010
Available from

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Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2011-216958 1

Flow Field Characterization of an Angled 

Supersonic Jet Near a Bluff Body 

John D. Wolter

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Robert Childs*

ELORET Corporation

Moffett Field, California 94035

Mark P. Wernet

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Andrea Shestopalov†

ELORET Corporation

Moffett Field, California 94035

John E. Melton

National Aeronautics and Space Administration

Ames Research Center

Moffett Field, California 94035

Abstract

An experiment was performed to acquire data from a hot supersonic jet in cross flow for the purpose 

of validating computational fluid dynamics (CFD) turbulence modeling relevant to the Orion Launch 

Abort System. Hot jet conditions were at the highest temperature and pressure that could be acquired in 

the test facility. The nozzle pressure ratio was 28.5, and the nozzle temperature ratio was 3. These 

conditions are different from those of the flight vehicle, but sufficiently high to model the observed 

turbulence features. Stereo Particle Image Velocimetry (SPIV) data and capsule pressure data are 

presented. Features of the flow field are presented and discussed.

Nomenclature

AAPL Aero-Acoustic Propulsion Laboratory

AM Abort Motor

CEV Crew Exploration Vehicle

CFD Computational Fluid Dynamics

CX Axial force coefficient

HFJER High Flow Jet Exit Rig

LAV Launch Abort Vehicle

NATR Nozzle Aeroacoustic Test Rig

SPIV Stereo Particle Image Velocimetry

φ Azimuthal angle

 *

Currently with Science & Technology Corporation, Moffett Field, California 94035. †

Currently with Exa Corporation, Burlington, Massachusetts 01803.
NASA/TM—2011-216958 2

Figure 1.—Ares I during launch (artist's rendition).

1.0 Background

Accurate prediction of aerodynamic forces on the Ares I Launch Abort Vehicle (LAV) is important 

for determining the forces on the Orion crew and the trajectory of the vehicle. This report describes an 

experiment to improve the computational fluid dynamics (CFD) prediction capability for this type of 

problem.

The Ares I vehicle (Figure 1) is designed to deliver the Orion Crew Exploration Vehicle (CEV) to 

low-earth orbit. A launch abort system provides a means to quickly remove Orion from danger in the 

event of an emergency. The LAV consists of a tower and the Orion Crew Module but excludes the Orion 

Service Module. The tower includes an attitude control motor with eight nozzles, a solid-rocket Abort 

Motor (AM) with four nozzles, and a boost protective cover.

CFD predictions have raised questions about the most appropriate turbulence model for such flows. A 

simulation by researchers at NASA Ames Research Center of the LAV 606-068 configuration was run at 

an angle of attack of 20°, and a flight Mach number of 0.3 with the abort motors firing at 100 percent 

power. All cases were run with the OVERFLOW code (Ref. 1). Three turbulence models were used: Lag

(Ref. 2), SST (Refs. 3 and 4), and Baldwin-Barth (Ref. 5). The cases run and their resulting axial force 

coefficients (Cx) are summarized in Table 1.
NASA/TM—2011-216958 3

TABLE 1.—CFD PREDICTIONS

Code version Turbulence

model

Cx

OVERFLOW 2.0aa “Lag_Robusta” Lag 1.8

OVERFLOW 2.0aa SST 0.8

OVERFLOW 2.0aa “Lag_Robusta” SST 0.7

OVERFLOW 2.0aa Baldwin-Barth 0.6

The difference between the two SST results is probably due to the difference in the turbulent kinetic

energy (TKE) production term, which is based on vorticity in the standard code and the strain rate in the 

Lag version. The large difference between the results from Lag turbulence model compared to those of 

other turbulence models raised questions about the accuracy of the predictions. Further investigation 

revealed that the mechanism for the higher force appears to be elevated turbulence transport in the shear 

layer of the AM plumes, resulting in energy transfer into the freestream which flows past the plumes and 

impinges on the LAV capsule. The mechanisms leading to elevated drag were dependent on jet 

temperature and hence velocity, for the fixed Mach number jet. These physical mechanisms were not 

implausible, and this raised the concern that jet temperature was an important parameter of the problem.

Furthermore, all wind tunnel tests of LAV aerodynamics are run with jet plumes that are substantially 

cooler than the flight vehicles, because hot jets are very difficult to create and work with in wind tunnels.

An experiment was needed to determine a “right answer” for this problem. Salient features of the 

experiment included: 1) a freestream flow, 2) a high-energy jet at an angle to the freestream, 3) a capsulelike shape on which pressures could be recorded, and 4) a means for measuring detailed flow field

characteristics of the plume and surrounding air. The AeroAcoustic Propulsion Laboratory (AAPL) at 

NASA Glenn Research Center was chosen as the site of this experiment based on its suitability for SPIV 

measurements and its capability to produce a high energy plume. However, the facility is not capable of 

pressures and temperatures as great as in the AM, so CFD studies were performed which confirmed that 

an experiment in this facility would provide an appropriate analog.

In many respects, this is a classic jet in crossflow problem. However, the high pressure and 

temperature of the jet, the angle of the jet, and the presence of the Orion capsule are novel aspects not 

found in the existing literature.

2.0 Equipment and Procedures

The experiment was conducted in the AeroAcoustic Propulsion Laboratory (AAPL) at the NASA 

Glenn Research Center between July and October of 2008. The AAPL is an acoustic test facility housed 

in a 130 ft diameter acoustically lined dome (Ref. 6). Three active test rigs are housed in the AAPL. For 

this experiment the Nozzle Aeroacoustic Test Rig (NATR) (Ref. 7) was used. The NATR is a 53-in.

diameter freejet rig capable of freestream velocities up to 0.3 Mach number. Installed within the NATR is 

the High Flow Jet Exit Rig (HFJER), a jet engine simulator (Ref. 8). Figure 2 depicts the NATR and 

HFJER within the AAPL facility.
NASA/TM—2011-216958 4

Figure 2.—Nozzle Aeroacoustic Test Rig (NATR) and High Flow Jet Exit Rig (HFJER).

Historically, the HFJER has supplied nozzles of 10 to 20 in.2 with heated flow at nozzle pressure 

ratios up to 4 and temperatures up to 1900 R at a flow rate up to 29 lbm/sec in single flow configuration.

For this test, the rig was modified to provide flow at a nozzle pressure ratio of 28.5 and temperature up to 

1460 R at a significantly lower flow rate of approximately 1 lbm/sec. Cold air was supplied to the HFJER 

continuously at up to 450 PSIG. The air was heated using a natural gas combustor.

The Stereo Particle Image Velocimetry (SPIV) system enabled cross-stream measurements of the 

three-component velocity field from the test article. The SPIV system employed two high-resolution 

(4008x2672 pixel) cameras to provide a 600- by 600-mm field of view. The measurement plane was 

illuminated using a 400 mJ/pulse Nd:YAG laser system. The flow was seeded with alumina seed particles 

in the nozzle core flow, and a propylene glycol based liquid seed material was used in the ambient flow 

stream. The large field of view was required in order to be able to capture the angled nozzle plume over a 

range of 20+ diameters downstream from the nozzle exit plane. In addition to the cross-stream planes, the 

SPIV system was reconfigured on the final test day to acquire axial plane slices of the 25° and 40° nozzle 

flow. The field of view of the SPIV system was nominally the same for the axial plane measurements as 

that used in the cross-stream configuration. The SPIV image data were processed to yield velocity vector 

maps with 1 mm spatial resolution. A sequence of 400 velocity vector maps were acquired at each 

measurement station and ensemble averaged to provide first- and second-order statistics over the entire 

measurement plane. All of the processed SPIV data were placed in the nozzle coordinate system and 

transferred to NASA Ames for comparison to CFD. Wernet et al. (Ref. 9) describe the SPIV data 

collection and results in greater detail. CFD validation studies are occurring at NASA Ames Research 

Center, and will be reported by A. J. Shestopalov et al. at a future date.

HFJER

NATR
NASA/TM—2011-216958 5

Figure 3.—85-AA Model in the NATR.

As described earlier, the problem to be modeled featured a hot, high-speed jet issuing at an angle 

relative to a freestream flow. The jet traveled past a capsule shape, not striking the capsule, but passing 

close enough that freestream flow striking the capsule interacted with the jet shear layer before reaching 

the capsule.

Figure 3 shows an illustration and photograph of the model configuration. The end of the HFJER was

capped and an extension from this cap served as a support for the capsule. A “mini-strut” and “ministing” were attached to the HFJER sting, to supply heated air to the nozzle. Three interchangeable nozzles 

were built, to vector the jet 0°, 25°, and 40° from the freejet axis. Figure 4 shows details of the nozzles’ 

internal flowpaths. Design exit Mach number of the nozzles was 2.78, and design NPR was 26.3. Nozzle 

instrumentation comprised a static pressure tap and a thermocouple upstream of the nozzle throat.

A support strut extended from the base of the HFJER to support the capsule model. The design of the 

support strut allowed for the repositioning of the capsule as needed. Two sets of bolt holes at the 

connection between the capsule and the support strut allowed the capsule to be mounted with its axis at 0°

(for the 25° nozzle) and 15° (for the 40° nozzle) relative to the freejet axis. Thus, the capsule rotates with 

the nozzle in these cases. These rotational positions represent the vehicle at different angles of attack. A 

series of 21 bolt holes in the vertical support strut allowed the capsule’s vertical position to be adjusted in 

1⁄2 in. increments. Capsule position was designated by a number from 1 (lowest position) to 19 (highest 

position). Figure 5 shows the capsule positions used during this experiment. The red portions of the figure 

represent the 25° vectored nozzle and the capsule mounted with 0° angle of attack in either position 10 or 

11. The blue portions of the figure represent the 40° vectored nozzle and the capsule mounted with 15°

angle of attack in either position 18 or 19. The axial position and angle of the capsule relative to the 

nozzle in these configurations were representative of the Ares I configuration; the vertical position was 

not. Instead, the vertical position was adjusted to match the capsule surface pressure characteristics 

observed in the original CFD problem, as described in Section 3.2.
NASA/TM—2011-216958 6

Figure 4.—Geometric details for 0°, 25°, and 40° nozzles.

Figure 5.—Capsule positions.
NASA/TM—2011-216958 7

Figure 6.—Capsule tap locations.

The capsule was a 4 percent scale model of the LAV without the umbilical cutout or the tower. A 

total of 63 static pressure taps were arranged on the capsule surface as shown in Figure 6.

For each configuration, the HFJER was operated long enough to reach thermal equilibrium, and then 

SPIV data were acquired. Steady-state data, such as test conditions and capsule pressures, were sampled

over a one second period. A plane of SPIV data required approximately two minutes of sampling to 

obtain a valid statistical population. A recording of the steady-state data was taken for each plane of SPIV 

data recorded. During operations, the primary test conditions (NPR, jet temperature, freejet Mach 

number) were held to no more than 0.5 percent root mean square error from the desired test condition. 

For the configurations with the capsule present, data was acquired in a two-step process. First, the 

model was run using cold air, and capsule pressures were recorded for several vertical positions of the 

capsule. From these runs a capsule position was selected for a run using heated air. The capsule position 

was selected based on evidence of impingement of vortical structures in the plume wake on the capsule, 

as observed in the CFD studies. For details of this selection process, see Section 3.2.

A total of 8 configurations were tested, as shown in Table 2. The test matrix included all of the 

available nozzles, with various SPIV configurations to capture different parts of the flow. Configurations 

with the capsule absent were intended to provide a baseline case with which to compare the 

configurations with the capsule present. In addition, configurations with the capsule absent served as a 

hedge against the possibility that laser flare off of the capsule would greatly reduce the quality of the 

SPIV data for configurations with the capsule present. Fortunately, this was not observed in this 

experiment.

TABLE 2.—CONFIGURATIONS TESTED

Configuration 

name

Nozzle

angle

Capsule

present?

SPIV configuration

Cancun 0 No Cross-stream

Bahamas 25 No Cross-stream

Cozumel 25 No Cross-stream, small field of view

Vail 25 No Streamwise

Galapagos 25 Yes Cross-stream

Puerto Vallarta 40 No Cross-stream

Vegas 40 No Streamwise

Rio de Janeiro 40 Yes Cross-stream
NASA/TM—2011-216958 8

3.0 Results

3.1 SPIV Measurements

Figure 7 and Figure 8 show velocity contours for the 25° and 40° nozzles, respectively, as viewed 

from the side. The SPIV measurement process and results from this experiment are more thoroughly 

discussed in a separate paper (Ref. 8). That paper also compares SPIV results to CFD of the experiment 

using the SST turbulence model with and without a compressibility correction. It concludes that neither 

CFD solution matches the SPIV data in all details (e.g., shock spacing, cross-stream vorticity, plume 

dispersion), with the SPIV data generally falling between the CFD solutions with and without 

compressibility correction.

Figure 7.—Velocity contours for 25° nozzle, NPR=28.5, T=1350 R.

Figure 8.—Velocity contours for 40° nozzle, NPR=28.5, T=1350 R.
NASA/TM—2011-216958 9

Figure 9 through Figure 14 show velocity results from the SPIV axial measurements. In these figures, 

axial location is normalized by the nozzle exit diameter. Note that “axial” in this context refers to planes 

normal to the freejet axis, not the jet axis. Figure 9 shows velocities for the 0° nozzle. Axial locations 

normalized by the nozzle diameter in the figure are: (top row) 0.55, 1.43, 2.31, 4.06, (middle row) 5.83, 

7.59, 9.79, 12.97, (bottom row) 17.37, 23.53, 32.00.

Figure 10 shows velocity contours for the 25° nozzle. Normalized axial locations in the figure are: 

(top row) 0.76, 2.35, 3.95, 5.54, (middle row), 7.14, 9.19, 12.04, 16.03, (bottom row) 21.55, 29.26.

Figure 9.—Velocity contours for 0° nozzle, no capsule, NPR=28.5, T=1350 R.

Figure 10.—Velocity contours for 25° nozzle, no capsule, NPR=28.5, T=1260 R
NASA/TM—2011-216958 10

Figure 11 and Figure 12 show velocity contours for the 40° nozzle at two different temperatures.

Normalized axial locations in Figure 11 are: (top row) 0.87, 1.53, 2.22, 3.56, (middle row) 4.91, 6.26, 

7.98, 10.40, (bottom row) 13.77, 24.95. Normalized axial locations in Figure 12 are: (top row) 0.87, 1.53, 

2.19, 2.85, (middle row) 3.51, 4.17, 4.83, 5.60, (bottom row) 6.28, 7.03.

Figure 11.—Velocity contours for 40° nozzle, no capsule, NPR=28.5, T=1350 R.

Figure 12.—Velocity contours for 40° nozzle, no capsule, NPR=28.5, T=1460 R.
NASA/TM—2011-216958 11

Figure 13 shows velocity contours for the 25° nozzle. Normalized axial locations in the figure are: 

(top row) 0.76, 2.35, 3.95, 5.54, (middle row) 7.14, 9.12, 11.98, 13.84, (bottom row) 15.93, 18.46, 21.54, 

29.24.

Figure 14 shows velocity contours for the 25° nozzle. Normalized axial locations in the figure are: 

(top row) 0.87, 2.19, 3.56, 4.91, (bottom row) 6.16, 8.00, 10.40, 13.74.

Figure 13.—Velocity contours for 25° nozzle with capsule, NPR=28.5, T=1350 R.

Figure 14.—Velocity contours for 40° nozzle with capsule, NPR=28.5, T=1350 R.
NASA/TM—2011-216958 12

Figure 15.—Capsule pressures, 25° nozzle at three jet conditions.

3.2 Capsule Pressures

Figure 15 through Figure 22 depict measured capsule surface pressures for the configurations in 

which the capsule was present.

Figure 15 shows capsule surface pressures for the 25° nozzle at three jet conditions: off, cold, and hot. 

Pressures are from the row of pressure taps closest to the jet (the φ=0° position in Figure 6). The pressures 

rise from the nose of the capsule until the break at x=24.7 in., then decrease as the flow accelerates 

around the bulk of the capsule. This behavior is reflected in all of the configurations where the capsule 

was present. This effect is enhanced in the cold jet as compared to jet off, and in the hot jet as compared 

to the cold jet. The enhancement is likely due to the entrainment of freestream flow by the jet flow.

Figure 16 through Figure 23 show capsule pressures at various positions and flow conditions. Refer to 

Figure 5 for capsule positions. In these pressure plots, the pressures are plotted by axial station. For 

Figure 16 through Figure 18 and Figure 20 through Figure 22, axial stations are denoted by color, with 

the image of the capsule to the right denoting the location of each station. 

Figure 16 and Figure 17 show capsule pressures for cold flow for the 25° nozzle configuration at 

capsule positions 10 and 11 respectively. While the two figures are similar, the position 11 plot shows a 

much deeper “dip” in pressure at the x=27.5 in. position. This dip is caused by vorticity from the 

interaction of the plume with the freestream. This dip was observed in the original CFD predictions and 

was considered an indicator that flow conditions were similar to those in the original CFD. Observations 

of the plume indicated that moving the capsule to position 12 would cause the jet to impinge directly on 

the capsule, which was not desired. Therefore, position 11 was chosen for hot flow studies.

Figure 18 shows capsule pressures for heated flow through the 25° nozzle with the capsule at position 

11. Figure 19 shows the data in Figure 17 and Figure 18 compared. The dip in the hot case is more 

pronounced for x=27.5 in. and aft, suggesting stronger vorticity.

Figure 20 and Figure 21 show capsule pressures for cold flow through the 40° nozzle with the capsule 

at positions 19 and 18, respectively. The differences between these two plots are much more subtle than 

in previous comparisons. Again the plume would impinge on the capsule at the next position, so position 

18 was chosen for hot flow.

x, inches

Pre

s

s

ure, P

SIA

22 24 26 28 30 13.4

13.6

13.8

14

14.2

14.4

14.6

14.8

15

15.2 Rdg 98 No Jet

Rdg 102 Cold Jet

Rdg 119 Hot Jet

CONFIG = GALA

CAPSULE POS = 11

MFJ = 0.30

PHI = 0 DEG
NASA/TM—2011-216958 13

Figure 22 depicts capsule pressures for heated flow through the 40° nozzle with the capsule at 

position 18. Figure 23 shows the data in Figure 21 and Figure 22 compared. Here, the difference between 

hot and cold flow is less pronounced than in the 25° case.

Figure 16.—Capsule pressures for 25° nozzle, position 10, cold.

Figure 17.—Capsule pressures for 25° nozzle, position 11, cold.

Phi, Degrees

Pre

s

s

ure, P

SIA

-150 -100 -50 0 50 100 150 13.4

13.6

13.8

14

14.2

14.4

14.6

14.8

15

Rdg99_X21.00

Rdg99_X24.70

Rdg99_X27.50

Rdg99_X28.50

Rdg99_X29.50

Capsule Surface Pressures

READING 99

CONFIG = GALA

CAPSULE POS = 10

NPR = 27.1

TT7 = 526.

MFJ = 0.30

Phi, Degrees

Pre

s

s

ure, P

SIA

-150 -100 -50 0 50 100 150 13.4

13.6

13.8

14

14.2

14.4

14.6

14.8

15

Rdg102_X21.00

Rdg102_X24.70

Rdg102_X27.50

Rdg102_X28.50

Rdg102_X29.50

Capsule Surface Pressures

READING 102

CONFIG = GALA

CAPSULE POS = 11

NPR = 28.6

TT7 = 522.

MFJ = 0.30
NASA/TM—2011-216958 14

Figure 18.—Capsule pressures for 25° nozzle, position 11, hot.

Figure 19.—Capsule pressures for 25° nozzle, position 11, cold 

and hot compared.

Phi, Degrees

Pre

s

s

ure, P

SIA

-150 -100 -50 0 50 100 150 13.4

13.6

13.8

14

14.2

14.4

14.6

14.8

15

Rdg119_X21.00

Rdg119_X24.70

Rdg119_X27.50

Rdg119_X28.50

Rdg119_X29.50

Capsule Surface Pressures

READING 119

CONFIG = GALA

CAPSULE POS = 11

NPR = 28.5

TT7 = 1348.

MFJ = 0.30

Phi, Degrees

Pre

s

s

ure, P

SIA

-150 -100 -50 0 50 100 150 13.4

13.6

13.8

14

14.2

14.4

14.6

14.8

15

Capsule Surface Pressures

READING 102

CONFIG = GALA

CAPSULE POS = 11

NPR = 28.6

TT7 = 522.

MFJ = 0.30

READING 119

CONFIG = GALA

CAPSULE POS = 18

NPR = 28.5

TT7 = 1348.

MFJ = 0.30
NASA/TM—2011-216958 15

Figure 20.—Capsule pressures, 40° nozzle, position 19, cold.

Figure 21.—Capsule pressures, 40° nozzle, position 18, cold.

Phi, Degrees

Pre

s

s

ure, P

SIA

-150 -100 -50 0 50 100 150 13.4

13.6

13.8

14

14.2

14.4

14.6

14.8

15

Rdg104_X21.00

Rdg104_X24.70

Rdg104_X27.50

Rdg104_X28.50

Rdg104_X29.50

Capsule Surface Pressures

READING 104

CONFIG = RIO

CAPSULE POS = 19

NPR = 28.5

TT7 = 521.

MFJ = 0.30

Phi, Degrees

Pre

s

s

ure, P

SIA

-150 -100 -50 0 50 100 150 13.4

13.6

13.8

14

14.2

14.4

14.6

14.8

15

Rdg107_X21.00

Rdg107_X24.70

Rdg107_X27.50

Rdg107_X28.50

Rdg107_X29.50

Capsule Surface Pressures

READING 107

CONFIG = RIO

CAPSULE POS = 18

NPR = 28.5

TT7 = 518.

MFJ = 0.30
NASA/TM—2011-216958 16

Figure 22.—Capsule pressures, 40° nozzle, position 18, hot.

Figure 23.—Capsule pressures, 40° nozzle, position 18, cold and 

hot compared.

 

Phi, Degrees

Pre

s

s

ure, P

SIA

-150 -100 -50 0 50 100 150 13.4

13.6

13.8

14

14.2

14.4

14.6

14.8

15

Rdg116_X21.00

Rdg116_X24.70

Rdg116_X27.50

Rdg116_X28.50

Rdg116_X29.50

Capsule Surface Pressures

READING 116

CONFIG = RIO

CAPSULE POS = 18

NPR = 28.5

TT7 = 1353.

MFJ = 0.30

Phi, Degrees

Pre

s

s

ure, P

SIA

-150 -100 -50 0 50 100 150 13.4

13.6

13.8

14

14.2

14.4

14.6

14.8

15

Capsule Surface Pressures

READING 107

CONFIG = RIO

CAPSULE POS = 18

NPR = 28.5

TT7 = 518.

MFJ = 0.30

READING 116

CONFIG = RIO

CAPSULE POS = 18

NPR = 28.5

TT7 = 1353.

MFJ = 0.30
NASA/TM—2011-216958 17

4.0 Concluding Remarks

An experiment was performed to obtain detailed flow information for the validation of Computational 

Fluid Dynamics (CFD) modeling of the Ares I Launch Abort System. The experiment focused on a better

understanding of the flow field details in the abort motor plume. The geometry and conditions of the 

experiment were designed not to exactly replicate those of the launch abort problem, but to provide a 

useful analog. Stereo Particle Image Velocimetry (SPIV) and surface pressure measurements were made.

A high quality data set has been created for hot, high-energy, angled jets in a cold cross flow with and 

without a bluff body. It is hoped that comparison of these data to calculations will lead to future 

improvements in turbulence modeling. At the time of this writing, the future of the Constellation program 

and the Ares I vehicle is uncertain. Regardless, the data from this experiment will be relevant to a variety 

of problems.

References

1. Liou, M.S. and Buning, P.G., “Contribution of the Recent AUSM Schemes to the OVERFLOW 

Code: Implementation and Validation,” AIAA–2000–4404, 18th AIAA Applied Aerodynamics 

Conference, Denver, CO, Aug. 2000. 

2. Olsen, M.E., and Coakley, T.J., “The Lag Model, a Turbulence Model for Non-equilibrium Flows,” 

AIAA–2001–2564, 2001. 

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the CEV Hot Abort Motor Plume for CFD Validation,” AIAA–2010–1031, 2010.
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01-03-2011 

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4. TITLE AND SUBTITLE

Flow Field Characterization of an Angled Supersonic Jet Near a Bluff Body 

5a. CONTRACT NUMBER

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5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Wolter, John, D.; Childs, Robert; Wernet, Mark, P.; Shestopalov, Andrea; Melton, John, E. 

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

WBS 644423.02.39.03.02 

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National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field 

Cleveland, Ohio 44135-3191 

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E-17563 

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14. ABSTRACT

An experiment was performed to acquire data from a hot supersonic jet in cross flow for the purpose of validating computational fluid 

dynamics (CFD) turbulence modeling relevant to the Orion Launch Abort System. Hot jet conditions were at the highest temperature and 

pressure that could be acquired in the test facility. The nozzle pressure ratio was 28.5, and the nozzle temperature ratio was 3. These 

conditions are different from those of the flight vehicle, but sufficiently high to model the observed turbulence features. Stereo Particle 

Image Velocimetry (SPIV) data and capsule pressure data are presented. Features of the flow field are presented and discussed.

15. SUBJECT TERMS

Computational fluid dynamics; Plumes; Particle image velocimetry; Turbulence models; Shear layer 

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