February 2014 

NASA/TM2014-218171 

NESC-RP-11-00697 

 

Orion Multi-Purpose Crew Vehicle (MPCV) 

Capsule Parachute Assembly System (CPAS) 

Wake Deficit Wind Tunnel Testing 

James C. Ross 

Ames Research Center, Moffett Field, California 

David M Schuster/NESC 

Langley Research Center, Hampton, Virginia 
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February 2014 

NASA/TM2014-218171 

NESC-RP-11-00697 

 

Orion Multi-Purpose Crew Vehicle (MPCV) 

Capsule Parachute Assembly System (CPAS) 

Wake Deficit Wind Tunnel Testing 

James C. Ross 

Ames Research Center, Moffett Field, California 

David M Schuster/NESC 

Langley Research Center, Hampton, Virginia 
Available from: 

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Hanover, MD 21076-1320 

443-757-5802 

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National Aeronautics and Space Administration. 
NASA Engineering and Safety Center 

Technical Assessment Report 

Document #: 

NESC-RP11-00697 

Version: 

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Title: 

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NESC Request No.: TI-11-00697 

Orion Multi-Purpose Crew Vehicle (MPCV) Capsule Parachute 

Assembly System (CPAS) Wake Deficit Wind Tunnel Testing 

January 23, 2014 
NASA Engineering and Safety Center 

Technical Assessment Report 

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Version: 

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Title: 

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NESC Request No.: TI-11-00697 

Report Approval and Revision History 

NOTE: This document was approved at the January 23, 2014, NRB. This document was 

submitted to the NESC Director on February 13, 2014, for configuration control. 

Approved: Original Signature on File 2/13/14 

Acting NESC Director Date 

Version Description of Revision Office of Primary 

Responsibility Effective Date 

1.0 Initial Release Dr. David Schuster, NASA 

Technical Fellow for 

Aerosciences, LaRC 

1/23/14 
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Table of Contents 

Technical Assessment Report 

1.0 Notification and Authorization ........................................................................................ 7

2.0 Signature Page ................................................................................................................... 8

3.0 Team List ......................................................................................................................... 9

3.1 Acknowledgements ................................................................................................. 9

4.0 Executive Summary ........................................................................................................ 10

5.0 Introduction and Objectives .......................................................................................... 11

5.1 Boundary Layer Tripping Study ........................................................................... 12

5.2 Transonic Wind Tunnel Test ................................................................................ 12

6.0 Boundary Layer Tripping Study ................................................................................... 13

6.1 Test Cell 2 at the Ames Research Center (ARC) Fluid Mechanics Laboratory 

(FML) ................................................................................................................... 14

6.2 Trip Study at Texas A&M University Low-Speed Wind Tunnel ......................... 16

6.3 High-Reynolds Number Test of Orion CM .......................................................... 16

7.0 Transonic Wind Tunnel Test ......................................................................................... 17

7.1 Model Configuration ............................................................................................. 17

7.2 Model Mounting ................................................................................................... 17

7.3 CM Model Details................................................................................................. 20

7.4 Instrumentation ..................................................................................................... 25

8.0 Test Conduct.................................................................................................................... 38

8.1 IR Thermography .................................................................................................. 40

8.2 PSP Measurements................................................................................................ 47

8.3 PIV Measurements ................................................................................................ 50

8.4 Unsteady Pressure Measurements......................................................................... 58

8.5 Boundary Layer Measurements ............................................................................ 62

9.0 Findings, Observations, and NESC Recommendations .............................................. 64

9.1 Findings................................................................................................................. 64

9.2 Observations ......................................................................................................... 65

9.3 NESC Recommendations...................................................................................... 65

10.0 Alternate Viewpoint ........................................................................................................ 65

11.0 Other Deliverables .......................................................................................................... 65

12.0 Lessons Learned .............................................................................................................. 65

13.0 Recommendations for NASA Standards and Specifications ...................................... 65

14.0 Definition of Terms ......................................................................................................... 66

15.0 Acronyms List ................................................................................................................. 67

15.1 Nomenclature ........................................................................................................ 67
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16.0 References ....................................................................................................................... 68

17.0 Appendices ....................................................................................................................... 70

List of Figures 

Figure 6.1-1. MPCV CM Model Mounted in 32x48-inch FML Subsonic Wind Tunnel ........... 14

Figure 6.1-2. IR Image of the Orion Model Showing Transition on the Heatshield at the Trip 

Location ................................................................................................................ 15

Figure 6.1-3. Oil Flow Visualization on the Heatshield Corresponding to the IR Image in 

Figure 6.1-2 ........................................................................................................... 16

Figure 7.1-1. Definition of Smooth CM Geometry (dimensions are model-scale inches); 

x- and z-Axes are shown with the Origin at the Heatshield Apex ........................ 17

Figure 7.2-1. Oblique View of the Model Support Strut Arrangement ...................................... 18

Figure 7.2-2. Side and Top Views of Model Installation showing the Lateral Offset Cable 

Arrangement (without cable fairings) and the Floor Plates and Feet at the Base of 

the Lower Struts .................................................................................................... 19

Figure 7.2-3. Interior Pivot and Pivot Cover Plates on the CM Model ...................................... 20

Figure 7.3-1. Computer-Aided Design (CAD) Image of the CM Mounted to Struts showing 

Cover Plates, Boundary Layer Probe, and Leveling Plate .................................... 21

Figure 7.3-2. Pin and Bolt Arrangements for the Three CM AoAs ........................................... 21

Figure 7.3-3. Back Shell Design showing Inner Details ............................................................ 22

Figure 7.3-4. Unsteady Transducer Insert .................................................................................. 23

Figure 7.3-5. Cutout in Heatshield Shoulder to Accommodate an Unsteady Transducer 

Insert ..................................................................................................................... 23

Figure 7.3-6. Rough Heatshield Surface ..................................................................................... 24

Figure 7.3-7. Dimples Machined into Rough Heatshield Surface .............................................. 24

Figure 7.4-1. Pressure Tap Layout on the Heatshield, K1-K40 are Unsteady Transducers ....... 26

Figure 7.4-2. Pressure Tap Layout on the Heatshield and Back Shell, K1-K44 are Unsteady 

Transducers ........................................................................................................... 27

Figure 7.4-3. Kulite® Transducer Holder ................................................................................... 28

Figure 7.4-4. Layout of Cameras and Lights around Test Section in Vinci Modeling Tool 

(view of right-hand wall; wind is right to left) ..................................................... 29

Figure 7.4-5. Illumination quality (left) and Camera View Angle Relative to Surface Normal 

(right) for PSP Cameras ........................................................................................ 29

Figure 7.4-6. Traversing Boundary Layer Total-Pressure Probe, Stanton Gauge, and Static 

Pressure Tap .......................................................................................................... 30

Figure 7.4-7. Boundary Layer Insert .......................................................................................... 30

Figure 7.4-8. Boundary Layer Insert Mounted in Rough Heatshield ......................................... 31

Figure 7.4-9. Strain Gauges on Left Upper Strut........................................................................ 32

Figure 7.4-10. Laser Sheet Paths in 11x11-ft Test Section .......................................................... 34

Figure 7.4-11. PIV Laser Sheets ................................................................................................... 35

Figure 7.4-12. PIV Imaging Areas for Both Model Positions ...................................................... 36
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Figure 8.1-1. IR Image of the Rough Heatshield at Mach 0.7, AoA 30 degrees, and Reynolds 

Number 107 ........................................................................................................... 41

Figure 8.1-2. PSP Image Compared to IR Image with Temperature Range to Show Stagnation 

Point; Rough Heatshield at Mach 0.3, AoA of 15 degrees, and Reynolds Number 

5.3x106 .................................................................................................................. 42

Figure 8.1-3. IR Image Showing Changes in Flow Structure with Increasing Mach 

Number ................................................................................................................. 42

Figure 8.1-4. IR Images of Heatshield taken with the Ceiling Mounted Camera at Mach 0.5, 

AoA of 30 degrees, and Reynolds Number 1.33x106 ........................................... 43

Figure 8.1-5. Difference in Roughness Depth Across Machining Seam .................................... 44

Figure 8.1-6. Top-view IR Image of Smooth Heatshield; M = 0.7, AoA of 30 degrees, and 

Reynolds Number 10x106

 ..................................................................................... 44

Figure 8.1-7. Side-view of IR Image; M = 0.7, AoA of 30 degrees, and Reynolds Number 

10x106 ................................................................................................................... 45

Figure 8.1-8. Map of Defects on Smooth Heatshield ................................................................. 46

Figure 8.1-9. IR Image of Smooth Heatshield Mach 1.02, AoA of 30 degrees, and Reynolds 

Number 1.33x106

 .................................................................................................. 47

Figure 8.2-1. Pressure Distributions on CM at AoA of 30 Degrees ........................................... 48

Figure 8.2-2. Axial Force Coefficient Variation with Mach Number ........................................ 50

Figure 8.3-1. Average Velocity in CM Wake; Mach 0.7, AoA of 14.0 degrees, and Reynolds 

Number 10x106 Color Contours of Vx, Every 10th Vector Shown ....................... 51

Figure 8.3-2. RMS of Vx in CM Wake; Mach 0.7, AoA of 14.0 degrees, and Reynolds Number 

10x106 ................................................................................................................... 52

Figure 8.3-3. Velocity Measurements from High-speed PIV; Mach 0.3, AoA of 14.0 degrees 

Reynolds Number 5.3x106

, and V∞ = 103 m/s ..................................................... 53

Figure 8.3-4. RMS X-velocity Measurements from High-speed PIV; Mach 0.3, AoA of 14.0 

degrees, and Reynolds Number 5.3x106

 ............................................................... 53

Figure 8.3-5. Distribution of Vx Along Center of Wake at Mach 0.7, AoA of 13.9 degrees, 

and Reynolds Number 107 .................................................................................... 54

Figure 8.3-6. Variation in Total Velocity at Various Downstream Locations; Mach 0.7, 

AoA of 13.9 degrees, and Reynolds Number 107

 ................................................. 55

Figure 8.3-7. Contours of Vx in Far-field Wake (Mach 0.7, AoA of 15 degree) at an Instant 

Showing Low Velocity Across Approximate Drogue Parachute Diameter, 

V∞ = 234 m/s ......................................................................................................... 56

Figure 8.3-8. Contours of Vx in Far-field Wake (Mach 0.7, AoA of 15 degrees) at an Instant 

Showing High Velocity Across Approximate Drogue Parachute Diameter, 

V∞ = 234 m/s ......................................................................................................... 56

Figure 8.3-9. DLC from Current Study Compared to Results in Reference 1 ........................... 58

Figure 8.4-1. 1/3-octave SPL Levels for Unsteady Pressures; Mach 0.7, AoA of 29.25 degrees, 

and Reynolds Number 10x106 .............................................................................. 59

Figure 8.4-2. Narrow Band SPL Levels Showing Effect of Roughness on Shedding Frequency; 

Mach 0.7, AoA of 29.25 degrees, and Reynolds Number 10x106

 ........................ 60
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Figure 8.4-3. Narrow Band SPL Levels Showing Effect of Mach Number on Shedding 

Frequency; Mach 0.7, AoA 29.25 degrees, and Reynolds Number 10x106 ......... 61

Figure 8.4-4. Correlation Coefficient versus Time Referenced to Sensor K8; Mach 0.7, AoA of 

29.25 degrees, and Reynolds Number 10x106

 ...................................................... 62

Figure 8.5-1. Boundary Layer Profiles for Rough Heatshield.................................................... 63

Figure 8.5-2. Example of Oil Film Images for Measuring Skin Friction; Mach 0.7, AoA of 

15 degrees, and Reynolds Number 10x106, Skin Friction Coefficient = 0.0026 

directed 4.6 degrees from Vertical ........................................................................ 64

List of Tables 

Table 5.2-1. Runs Accomplished during AUPWT Test (numbers in the boxes indicate the 

Reynolds Number) ................................................................................................ 13

Table 7.4-1. Flow Statistics in Archived PIV Data Files .......................................................... 37

Table 8.0-1. List of Runs Accomplished During Test .............................................................. 38

Table 8.2-1. Integrated Forces and Moments from Surface Pressures ...................................... 49

 
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Technical Assessment Report 

1.0 Notification and Authorization 

Dr. David Schuster, NASA Technical Fellow for Aerosciences at the NASA Langley Research 

Center, was selected to lead this assessment. The assessment plan was approved by the NASA 

Engineering and Safety Center (NESC) Review Board on May 12, 2011. 

The key stakeholders for this assessment are the Orion Multi-Purpose Crew Vehicle Capsule 

Parachute Assembly System Project and the Crew Exploration Vehicle Aerosciences Project. 

The Agency’s wind tunnel test facilities are also a key stakeholder. 
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2.0 Signature Page 

Submitted by: 

Team Signature Page on File – 2/25/14 

Dr. David M. Schuster Date 

Significant Contributors: 

Dr. James C. Ross Date Mr. J. T. Heineck Date 

Dr. Gloria K. Yamauchi Date Mr. Theodore J. Garbeff Date 

Mr. Marvin Sellers Date Ms. Laura K. Kushner Date 

Ms. Nettie Halcomb Date Mr. Nathan J. Burnside Date 

Signatories declare the findings, observations, and NESC recommendations compiled in the 

report are factually based from data extracted from program/project documents, contractor 

reports, and open literature, and/or generated from independently conducted tests, analyses, and 

inspections. 

 
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3.0 Team List 

Name Discipline Organization 

Core Team 

David Schuster NASA Technical Fellow for 

Aerosciences 

LaRC 

James Ross Aerospace Engineer ARC 

Patricia Pahlavani MTSO Program Analyst LaRC 

J.T. Heineck Photographic Technologist ARC 

Gloria Yamauchi Aerospace Engineer ARC 

Theodore Garbeff Instrumentation Engineer ARC 

Marvin Sellers Senior Engineer Aerospace Testing 

Alliance 

Laura Kushner Senior Engineer Aerospace Computing, 

Inc. 

Nettie Halcomb Aerospace Engineer ARC 

Nathan Burnside Aerospace Engineer ARC 

Administrative Support 

Diane Sarrazin Project Coordinator LaRC/AMA 

Linda Burgess Planning and Control Analyst LaRC/AMA 

Christina Williams Technical Writer LaRC/AMA 

3.1 Acknowledgements 

The staff at the Ames Research Center Unitary Plan Wind Tunnel provided outstanding support 

for the test, keeping the test on track. Management instituted customized shift hours that 

matched the unusual nature of the test and made test execution more efficient. The Facility 

Manager, John Holmberg, was invaluable in solving many problems during the testing. The Test 

Manager, Jennifer Everett, and Test Engineer, Ross Flach, were key contributors in balancing the 

needs of the test with the practical aspects of completing the test in the allotted time. 
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4.0 Executive Summary 

Sizing of the parachute system for the Multi-Purpose Crew Vehicle (MPCV) is critical for safe 

operation. In particular, the drogue parachutes need to be large enough to ensure reliable 

deployment and inflation, but making them too large unnecessarily increases the load on the 

vehicle, system weight, and packing volume. The wake deficit at the drogue parachute location 

behind the capsule is the primary parameter that determines the size of the parachutes. Empirical 

wake deficit estimates are available in the literature, but their validity for all configurations is not 

well documented. A computational method of determining the wake deficit would be preferred, 

but the available computational fluid dynamics (CFD) tools have not been reliable for predicting 

bluff body separation and wake development at the Mach numbers expected for drogue 

parachute deployment. This test provides data at a variety of relevant flight conditions to 

validate the computational tools. The data provided the Orion Capsule Parachute Assembly 

System (CPAS) team with a measured wake deficit value to verify the value selected for their 

drogue parachute design. 

The model tested in this assessment was an approximately 8 percent-scale generic/axisymmetric 

Orion crew module (CM). Measurements included the surface pressure distribution, skin friction 

at a point on the heatshield, separation locations around the heatshield shoulder, and velocity 

measurements in the CM wake to the drogue parachute location. The test was conducted at 

Mach numbers of approximately 0.3, 0.5, 0.7, 0.9, and 1.05, and angles of attack (AoA) of 15 

and 30 degrees. A brief study was done to examine the effect of heatshield roughness, but the 

majority of the data was acquired using a heatshield model that approximated the post-entry 

roughness of an MPCV Avcoat heatshield. 

The acquired data demonstrated the separated wake behind the CM was highly unsteady for all 

conditions tested, and the level of unsteadiness increased as the capsule AoA decreased and the 

heatshield was pointed directly into the flow (i.e., from 30 towards 0 degrees). In general, the 

testing confirmed that handbook wake persistence estimates were conservative, but these 

estimates do not effectively account for the unsteadiness of the wake flow or motion, which 

could be important to parachute design. Heatshield roughness was found to be an important 

contributor to the wake flowfield character and should be accounted for in future testing of these 

classes of vehicles. 

The surface and off-body flowfield data acquired and archived in this test are suitable for CFD 

code validation and comprise a unique dataset characterizing the unsteady flow behind a blunt 

capsule body. With the accelerating development of CFD tools for predicting unsteady separated 

flows, this dataset should be an important contributor to the validation of these future 

computational methods. 
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5.0 Introduction and Objectives 

During descent after re-entry into the Earth’s atmosphere, the Orion CM deploys its drogue 

parachutes at approximately Mach 0.7. Accurately predicting the dynamic pressure experienced 

by the drogue parachutes at deployment is critical to properly designing the parachutes. Previous 

recovery system designs made extensive use of data manuals to estimate the dynamic pressure at 

the parachute location [ref. 1]. The Orion CPAS team wanted improved estimates and employed 

CFD methods to model the flow around the capsule at the drogue parachute deployment 

condition. Unfortunately, this flight condition proved difficult to accurately model with CFD. 

The flow in the wake was difficult to measure in a wind tunnel since the model is usually 

mounted on a sting located in the wake region. The prediction and measurement shortcomings 

led to uncertainty in the Orion parachute system design. This shortcoming affected the CM 

aerodynamic predictions during descent leading to a mismatch between the aerodynamic 

database and the capsule aerodynamics measured during the Pad Abort 1 flight test. 

The task of predicting and measuring the MPCV CM aerodynamics was complicated by 

uncertainty concerning the boundary layer state on the heatshield. Based on the Reynolds 

number for the descent trajectories, the heatshield boundary layer is expected to be turbulent. 

After entering the atmosphere from the Earth’s orbit or lunar return, the heatshield surface is 

roughened due to Avcoat ablation from the high temperatures encountered. The roughness 

forces the boundary layer to be turbulent from the stagnation point with additional turbulent 

kinetic energy generated by the flow over the rough surface. The flow also experiences a 

favorable pressure gradient, which tends to stabilize the boundary layer and delay transition, 

particularly at the lower Reynolds numbers encountered in wind tunnel testing. During a 

potential launch abort (i.e., pad or ascent), the CM may experience lower Reynolds numbers at 

drogue-chute deployment with a smooth heatshield, resulting in larger areas of laminar boundary 

layer. Therefore, simulating the desired boundary layer state in the wind tunnel becomes even 

more problematic for this type of vehicle than for typical aircraft, which the team has a much 

broader experience base. This test, and other similar experiments, addresses the issue of 

modeling a rough heatshield and employs innovative techniques to simulate the proper boundary 

layer state on capsules in a wind tunnel. 

This assessment was designed to provide a complete set of flowfield measurements on and 

around an idealized Orion CM shape with the most appropriate wind tunnel simulation of the 

Orion flight conditions prior to parachute deployment. The first part of the study was to examine 

potential boundary layer tripping strategies to best simulate flight conditions. The bulk of the 

effort was spent performing a high-fidelity wind tunnel test in the Ames Unitary Plan Wind 

Tunnel (AUPWT). The test covered a range of Mach numbers, Reynolds numbers, and AoAs. 

The data includes full-body pressure distributions using pressure sensitive paint (PSP) anchored 

to approximately 60 discrete surface static pressure taps; infrared (IR) thermography to capture 

boundary layer transition and separation locations; particle image velocimetry (PIV) 

measurements of the velocity in the CM model wake; unsteady pressure measurements on the 
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heatshield shoulder and back shell; boundary layer surveys at one heatshield location; and skin 

friction at a point adjacent to the boundary layer survey. 

The CM model was a smooth body to avoid obscuring important flow physics with local flow 

disturbances from protuberances. The model was mounted in the wind tunnel so as to minimize 

the influence of the mounting struts on the model wake flow in a downstream vertical plane 

where air velocity was measured across the full wake. 

5.1 Boundary Layer Tripping Study 

Before running the more expensive test in the AUPWT, a smaller effort was undertaken to study 

the effect of various boundary layer tripping strategies on the CM aerodynamics. For most wind 

tunnel tests, the boundary layer is tripped using roughness elements. These elements can be 

made from wires or tape, lines of grit or small adhesive dots, or grit distributed over the surface. 

There are established criteria for determining the size of these roughness elements based on the 

flight conditions to be simulated and the conditions in the wind tunnel for aircraft models 

[refs 2, 3]. The criteria are different for wings and fuselages, but are established for both types 

of flows. Most aircraft manufacturers have refined their own sizing methods to ensure tripping 

the flow from laminar to turbulent with minimal drag increments due to the drag of the 

roughness elements themselves. 

The flow on the blunt entry capsule heatshield is different than found on most aircraft so it may 

be that those methods do not apply directly or may need modification. Two wind tunnel tests 

were included in this effort to determine what an appropriate tripping strategy should be for a 

capsule after re-entry. 

5.2 Transonic Wind Tunnel Test 

The comprehensive test of the CM model was performed in the 11x11-foot (ft) transonic wind 

tunnel test section of the AUPWT [ref. 4]. This test facility can obtain Mach numbers from 0.3 

to 1.4. The total pressure in the facility can be varied from 0.5 to 2.0 atmospheres providing a 

range of Reynolds numbers. The first priority of the transonic wind tunnel test was to document 

the velocity deficit into which the Orion drogue parachute is deployed at Mach 0.7. A longer 

term objective was to gather aerodynamic and flowfield data for the purpose of validating CFD 

methods. The measurements made in the wind tunnel included velocity distributions in the 

wake, time-averaged surface pressure distribution over the entire model, unsteady pressures 

around the heatshield shoulder, transition location on the heatshield, separation lines on the 

heatshield shoulder region, and skin friction on the heatshield. Time-averaged flowfield velocity 

measurements were made in a vertical plane downstream of the CM model from as close to the 

model as possible to approximately 6 model diameters downstream. A second set of high-rate 

(i.e., 1900 frames/second) velocity measurements were made in a smaller horizontal plane in the 

model near wake. 

The surface measurements were made at Mach numbers of approximately 0.3, 0.5, 0.7, 0.9, and 

1.05, and AoAs of 15 and 30 degrees. A limited number of measurements were acquired at 
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0 degree AoA. The PIV measurements were made only for 15 degrees AoA and Mach numbers 

of 0.3, 0.5, and 0.7. These conditions cover the expected Orion parachute deployment and lower 

Mach numbers to allow more comprehensive CFD tool validation. The test conditions and types 

of data acquired are shown in Table 5.2-1. 

Table 5.2-1. Runs Accomplished during AUPWT Test (numbers in the boxes indicate the Reynolds 

Number) 

6.0 Boundary Layer Tripping Study 

Two studies were performed in preparation for the main test in the AUPWT. The first was to 

apply IR thermography on a capsule model to ensure the flow features of interest would be 

visible and to obtain experience with the IR camera. This practice was in preparation for the 

second part of the study at Texas Agricultural and Mechanical (A&M) University to look at the 

effects of different tripping strategies on the CM aerodynamics and flow structure. In addition, 

the IR image quality was examined for the models fabricated of aluminum and from nylon using 

selective laser sintering (SLS). 
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6.1 Test Cell 2 at the Ames Research Center (ARC) Fluid Mechanics 

Laboratory (FML) 

The first test was performed in the ARC FML Test Cell 2 32x48-inch subsonic wind tunnel. The 

model used in the study is shown in Figure 6.1-1, mounted on a sting and balance. The SLS 

material is white and fairly rough when manufactured so the model was filled and painted to give 

a smooth heatshield. The black color provided good contrast with the white oil (Figure 6.1-1) 

used for flow visualization with IR imagery. 

Figure 6.1-1. MPCV CM Model Mounted in 32x48-inch FML Subsonic Wind Tunnel 

Note: For flow visualization, white oil is spread on the model heatshield prior to tunnel operation. 

This study generated IR imagery showing detail about the flow over the model surface. Since 

this is a low-speed indraft wind tunnel, the flow temperature was nearly the ambient temperature 

and essentially the same as the model. Therefore, prior to each run, the model was heated above 

the ambient by approximately 20 degrees. This was done to enhance the flow-induced 

temperature differences on the models surface using the IR camera. The images in Figures 6.1-2 

and 6.1-3 show an IR image and the corresponding oil flow visualization. In this case, a 

boundary layer trip was applied to the heatshield consisting of a band of course grit (~60 grit). 

The oil was applied from the trip outboard on the heatshield and showed separation occurring 

near the start of the heatshield shoulder (Figure 6.1-3). The IR image in Figure 6.1-2 was taken 

through the only available open hole in the tunnel wall, which was perpendicular to the 

heatshield shoulder. This image shows transition occurring at the trip location and the separation 

at the same location as in the oil flow visualization. 
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The images obtained during this test demonstrated the IR images would show transition from 

laminar to turbulent flow and possibly show the separation line. The main concern for this 

technique was maintaining a temperature difference between the airstream and the interior model 

temperature. Since the technique had been successful when used in the AUPWT 11x11-ft test 

section, the FML results indicated the IR cameras had sufficient sensitivity to show the desired 

flow features. 

Figure 6.1-2. IR Image of the Orion Model Showing Transition on the Heatshield at the Trip Location 

Laminar Transition/trip 

location
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Figure 6.1-3. Oil Flow Visualization on the Heatshield Corresponding to the IR Image in Figure 6.1-2 

Note: Streaky oil may be coincident with the streaks in IR image; separation is indicated by the white band around 

the shoulder. 

6.2 Trip Study at Texas A&M University Low-Speed Wind Tunnel 

The CEV CPAS team performed a study of wake flows and parachute loads downstream of the 

Orion CM in the 7x10-ft Low-Speed Wind Tunnel at Texas A&M University [ref. 5]. The test 

used an aluminum, 24-inch diameter, smooth CM model mounted on a six-component balance. 

Model drogue parachutes were flown in the CM wake on load cells to measure the loads. The 

first portion of the test was used to determine the most appropriate method of tripping the 

boundary layer to simulate higher Reynolds number flight. 

This test resulted in interesting data, but was limited due to the time constraints. The main 

conclusion from the test was that distributed roughness using carborundum grit produced 

different aerodynamics than trip rings made from the standard Mylar® trip dots. 

6.3 High-Reynolds Number Test of Orion CM 

The Orion Aeroscience team performed a test in the National Transonic Facility at LaRC to 

determine the effect of Reynolds number on the CM subsonic/transonic aerodynamics [ref. 6]. 

The test was run at low and at flight-relevant Reynolds numbers. In addition, the effect of 

heatshield roughness was examined. The test showed there was a significant effect of roughness 

(i.e., representative of post-entry Avcoat) on the subsonic aerodynamics. In particular, the rough 

heatshield increased the CM drag coefficient by approximately 10 percent at a Mach number of 

0.3 and Reynolds number of 24 million based on CM diameter. As the Mach number increased, 
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the change in drag due to roughness decreased until Mach 0.8. Above Mach 0.8, there was no 

measurable difference due to roughness. 

7.0 Transonic Wind Tunnel Test 

7.1 Model Configuration 

The model was an 8.08 percent-scale (16-inch diameter) representation of the smooth, 

axisymmetric Orion CM shape. It did not include any of the protuberances or other details on 

the back shell or heatshield. The geometry definition and the x- and z-axes are shown in 

Figure 7.1-1 where the y-axis is directed into the page. The origin of the model coordinate 

system is as shown, at the apex of the heatshield. 

Figure 7.1-1. Definition of Smooth CM Geometry (dimensions are model-scale inches); x- and z-Axes 

are shown with the Origin at the Heatshield Apex 

7.2 Model Mounting 

The model was mounted on a pair of struts attached to the floor of the 11x11-ft wind tunnel test 

section as shown in Figure 7.2-1. The struts were designed to locate the model at two axial and 

two lateral positions. One axial location placed the model so the vertical wake measurement 

plane covered the area from the heatshield shoulder to ~3 model diameters downstream while the 

other axial location allowed the same measurement plane to cover a region to ~6 model 

diameters downstream. The two lateral locations were on the test section centerline and 
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4 inches to the left of the centerline (y/D = -0.25). Data were only acquired for the forward and 

aft positions on the tunnel centerline due to time constraints.

The model was mounted on struts in the wind tunnel to minimize the influence of the model 

support on the wake flow, particularly in a vertical plane downstream where the PIV 

measurements were made. Figure 7.2-2 shows the model and struts mounted in the wind tunnel. 

Model motions were minimized by stabilizing the upper mounting struts with cables to reduce 

the uncertainty in the various optical measurements. The cables were attached to the test section 

floor upstream and outboard of the strut attachment points. The upper struts were instrumented 

with strain gauges near their attachment to the lower struts and on the model side of the cable 

attachment fairings. The gauges were calibrated with known loads in the lift and drag directions 

and monitored during the test for safety. The resulting support was very rigid and the model 

moved less than 0.25 inches (0.015D) due to the mean aerodynamic load. The unsteady model 

motion was less than 0.1 inch (~0.006D). 

Figure 7.2-1. Oblique View of the Model Support Strut Arrangement 

Sidewall windows 

(in both walls)

Instrumentation 

cables

Right, lower 

strut

Right, upper 

strut

Floor windows 

(also in ceiling)
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Figure 7.2-2. Side and Top Views of Model Installation showing the Lateral Offset Cable Arrangement 

(without cable fairings) and the Floor Plates and Feet at the Base of the Lower Struts 

The struts are untwisted, tapered 654-021 airfoil sections truncated at the 95 percent chord 

location. The fairings at the upper-to-lower strut junctions provided more bearing area for the 

bolted joint. The fairing at the junction of the cables and upper struts provided a better 

aerodynamic shape for the junction to minimize any disturbance to the wake downstream. 

The cables were intended to have pivoting fairings to minimize their wakes. During the initial 

runs of the wind tunnel test the fairings did not pivot into the wind as intended. They were 

removed and the test was completed without them. The large sweep angle of the cables relative 

to the flow helped reduce the cable wake.

The model was pivoted at the ends of the mounting struts to AoAs of 0, 15, and 30 degrees. 

Removable filler pieces around the strut/model intersection maintained the correct model 

geometry and made a tight seal around the strut tips. The model set at 0 degree AoA was 

positioned in the wind tunnel by placing the model origin (heatshield apex) at x = 151.7, y = 0.0, 

and z = -4.9 inch (wind-tunnel coordinates). The rotation center is shown in Figure 7.1-1. The 
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CM pivot is shown on the model in Figure 7.2-3. The AoA was set by matching specific pin and 

bolt holes on the strut flange and the CM model. 

Figure 7.2-3. Interior Pivot and Pivot Cover Plates on the CM Model 

The major objective of the test was to document the wake flowfield up to 6-model diameters 

downstream. The resulting area of interest was too large to image using the available cameras so 

either the model or PIV system would have to move to cover the entire region. PIV systems are 

difficult to set up, align, and calibrate, so the model was moved relative to the measurement 

plane to acquire velocity data in the near wake and far downstream. The floor plates were drilled 

and tapped with two bolt patterns for the lower strut attachments, spaced 4 ft apart in the 

streamwise direction. After all the measurements were made, with the model in the aft position, 

the lower struts were unbolted from the floor plates and the model was moved to the upstream 

position and bolted to the second set of threaded holes. The yellow cables in Figure 7.2-1 are the 

bundles of instrumentation cables that ran through holes bored through the struts with sufficient 

slack to move the model to the second mounting location. 

7.3 CM Model Details 

Figure 7.3-1 shows the CM mounted to the struts with the leveling plate and the boundary layer 

survey probe installed. The strut attachments were covered with a split cover plate that restores 

the proper back shell shape around the strut penetration. Three sets of cover plates were required 

for the test (i.e., one for each AoA). The cover plates were manufactured using polycarbonate 

SLS that was tailor fit around the upper strut tips and flush to the surrounding back shell 

surfaces. The leveling plate was used to mount a precision inclinometer to establish the installed 

AoA in the wind tunnel. 
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Figure 7.3-1. Computer-Aided Design (CAD) Image of the CM Mounted to Struts showing Cover 

Plates, Boundary Layer Probe, and Leveling Plate 

Figure 7.3-2 shows the pin and bolt arrangement for changing the CM AoA. A single pin and 

three bolts on one side of the model were sufficient to carry the pitching moment CM with the 

required factor of safety. The test was run with a pin and at least five bolts on each side. 

Figure 7.3-2. Pin and Bolt Arrangements for the Three CM AoAs 
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The CM back shell was fabricated using normal machining techniques from a billet of 

7075 aluminum. Figure 7.3-3 shows the details of the inside of the part. The two pivot points 

and the associated bolt/pin holes are evident in this view as is the mount at the bottom for the 

AoA sensor. Steel inserts were used for the pivot bearing surface. The strut pocket holes were 

for quick disconnects that attached the pressure tap tubing from the strut cover plates. A circular 

flat plate closes the hole of the truncated CM apex. 

Figure 7.3-3. Back Shell Design showing Inner Details 

The heatshield was a two-layer part with an inner layer of turned 7075 aluminum to which a 

0.25-inch outer layer of SLS polycarbonate was bonded prior to final machining of the outer 

surface. This was done to improve the IR imaging based on the prior testing done in the ARC 

FML using the plastic model, and at the Texas A&M’s 7x10-ft low-speed wind tunnel using the 

aluminum model. The plastic model in the preliminary tests provided a better IR imagery. 

Two heatshields were made for the test: one smooth and the other machined with a hexagonal 

dimple pattern to simulate the expected roughness of the Orion heatshield after atmospheric 

entry. Because of the expense and lead time to acquire the unsteady pressure transducers, four 

inserts were made to hold ten transducers each that could be installed in the heatshield. The 

inserts were made from polycarbonate SLS and machined with the smooth heatshield outer 

surface to get the correct shape. An image of one insert is shown in Figure 7.3-4 and its 

corresponding recess in the smooth heatshield is shown in Figure 7.3-5. 
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Figure 7.3-4. Unsteady Transducer Insert 

Figure 7.3-5. Cutout in Heatshield Shoulder to Accommodate an Unsteady Transducer Insert 

Note: The red material is epoxy to fill a crack in the polycarbonate material. 

The machined surface of the rough heatshield is shown in Figures 7.3-6 and 7.3-7. The fit of the 

unsteady pressure transducer insert is shown in Figure 7.3-6 (flush with the tops of the dimpled 

pattern). The details of the pattern are shown in Figure 7.3-7 with the dimensions, which were 

determined by physically scaling the expected roughness pattern of the Orion Avcoat heatshield 
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material after atmospheric entry [ref. 7]. The pink color of the dimpled heatshield was due to the 

PSP coating that was applied prior to the test. 

Figure 7.3-6. Rough Heatshield Surface 

Figure 7.3-7. Dimples Machined into Rough Heatshield Surface 
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7.4 Instrumentation 

7.4.1 Pressure Measurements 

The CM model had 63 static pressure taps at the locations listed in Appendix A. These were 

primarily intended to check/calibrate the PSP results. Since PSP can be temperature sensitive, 

the various types of surfaces on the model needed to have taps to help correct for temperature 

effects. The pressure taps were arranged as shown in Figures 7.4-1 and 7.4-2. The static 

pressure taps were clustered in regions of large pressure gradients. There were taps on the back 

shell, pivot cover plates, and the flat end of the CM. These areas were expected to have different 

flow features and associated pressure distributions, but also had different materials and 

thicknesses. These differences could affect their temperature during the test, potentially 

affecting the PSP images/data. 

The discrete static pressure measurements were acquired using a 15-pounds per square inch 

(psid), electronically-scanned pressure module (Pressure System, Inc. Model 64HD0701021000) connected to the AUPWT data acquisition system. This module had 64 pressure 

transducers and the table of static pressure showed 65 static taps. Standard practice for using 

these kinds of pressure modules assigns 1 or 2 of the ports on a module to a known reference 

pressure, meaning that only 62 ports were available for test measurements. The difference is 

made up by disconnecting three of the back shell pressure taps and using them for the pressure 

measurements on the boundary layer survey when it was installed (see TAP ID 63-65 in 

Appendix A). 

The unsteady pressure measurements were made using 44 Kulite® Semiconductor Products, Inc., 

model XCL-32-072-15D transducers. Forty of these transducers were installed in the four inserts 

described in Section 7.3. The sensor spiral arrangement around the heatshield shoulder was 

designed so no sensor was in the wake of another sensor. Four additional transducers were 

located on the back shell to provide information on the unsteady pressure field that might affect 

the MPCV pressure altimeter. The locations of the unsteady pressure transducers are shown in 

Figures 7.4-1 and 7.4-2. 
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Figure 7.4-1. Pressure Tap Layout on the Heatshield, K1-K40 are Unsteady Transducers 
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Figure 7.4-2. Pressure Tap Layout on the Heatshield and Back Shell, K1-K44 are Unsteady 

Transducers 

The unsteady transducers were mounted in the model using custom-fit holders (Figure 7.4-3). 

The holders served two functions: to accurately set the distance between the sensor diaphragm 

and the model surface and to make it easier to remove and replace sensors in case of failure. The 

holder was inserted into a countersink in the model surface. The depth of the countersink was 

machined so all of the sensor diaphragms were set below the model surface (0.030 ±0.002 inch). 

The holder was machined to be flush to the model surface. The transducers were inserted from 

the inside of the model into the holder and then were bonded and sealed to the back side of the 

holder using a soluble glue (DucoTM cement). The locations of the unsteady pressure transducers 

are given in Appendix B. 
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Figure 7.4-3. Kulite® Transducer Holder 

7.4.2 PSP 

PSP is a mature test technique commonly used in wind tunnels [ref. 8]. The main requirement 

for its use is good visibility of the model surfaces of interest for illumination and imaging. 

Multiple cameras were required to cover the model surfaces with acceptably small angles 

between the imaged surface and the camera view axis. The Vinci modeling program [refs. 9, 10] 

is used at ARC to plan camera and light placements and to estimate the resulting imaging 

quality. 

Figure 7.4-4 shows the camera arrangement in the right-hand wall of the test section. The 

cameras and lighting for the various optical measurements need to be included in the Vinci 

modeling to ensure there is no imaging conflict (i.e., viewing or shadowgraph). The left-side of 

Figure 7.4-4 shows the PSP lighting quality. In Figure 7.4-5, the blue indicates good 

illumination and the red areas are approximately two f-stops lower in illumination. The rightside of Figure 7.4-5 shows the surface normal direction relative to the combined views of all the 

cameras. In this case, green indicates that at least one PSP camera has a view that is within 

15 degrees of the surface normal. The darker areas in this image indicate ~60 degrees between 

the surface normal and the best camera view. These areas have lower data quality. The camera 

arrangement restricted these areas to relatively small portions of the struts, which were of lower 

importance to understanding the CM aerodynamics. 

Eight Photometrics CoolSnap K4 2048-by 2048-pixel cameras with 7.4 micrometre (μm) pixel 

pitch were used for the PSP imaging while 15 400 nm light emitting diode lights from Innovative 

Scientific Solutions, Inc., were used for the illumination. The cameras and lamps were located in 

both sidewalls, the floor, and ceiling of the test section. Some of the windows for the cameras 

and lights are shown in Figure 7.2-1. 
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Figure 7.4-4. Layout of Cameras and Lights around Test Section in Vinci Modeling Tool (view of righthand wall; wind is right to left) 

Figure 7.4-5. Illumination quality (left) and Camera View Angle Relative to Surface Normal (right) for 

PSP Cameras 

Note: In left figure, blue indicates good illumination and red is two f-stops lower; in right figure, green indicates 

surface normal within 15 degrees of at least one camera axis. 
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7.4.3 Boundary Layer Measurements 

A boundary layer traverse mechanism was installed in the model for a portion of the test. When 

not installed, the mechanism was replaced by a blank plate machined (and textured) to match the 

heatshield. The traverse insert is shown in Figure 7.4-6. Figure 7.4-7 shows the total pressure 

probe driven by a Zaber Technologies T-NA08A25 stepper motor and TSB28M translation 

stage. The probe tip geometry allowed total pressure measurements to 0.0049 inch above the 

surface and the traverse had approximately 1 inch of travel. The traversing mechanism and 

overall construction of the insert are shown in Figure 7.4-7. 

Figure 7.4-6. Traversing Boundary Layer Total-Pressure Probe, Stanton Gauge, and Static Pressure 

Tap 

Figure 7.4-7. Boundary Layer Insert 

Translation 

stage

Nonconducting, 

traversing post 

Mounting plate

Aluminum - 

polycarbonate 

insert
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The boundary layer traverse insert was made of the same materials as the rest of the heatshields 

(i.e., an aluminum base plate with 0.25 inches of bonded polycarbonate). The insert was 

machined flush with the smooth heatshield to ensure minimal flow disturbance. Blank plates of 

aluminum/polycarbonate laminate were machined to replace the insert for most of the test series. 

Separate blanking plates were made for the smooth and rough heatshields with the appropriate 

surface texture. The outer surface of the traverse insert was nickel-plated to provide a smooth, 

reflective surface for the oil-film interferometry (OFI) measurement of the skin friction. OFI is 

an established measurement technique that is minimally intrusive [refs. 1, 12]. Adjacent to the 

probe were a static pressure tap and a Stanton gauge (Figure 7.4-6) used to measure the skin 

friction [refs. 13, 14, 15]. The Stanton gauge was calibrated against the OFI skin friction 

measurements, which were only obtained at a few test conditions due to the time required for 

accurate OFI measurements. The machining and nickel-plating process left the tap recessed 

below the surrounding surface. This recession was filled with cyanoacrylate glue, sanded 

smooth, and the tap hole re-drilled. 

An image of the insert installed in the rough heatshield is shown in Figure 7.4-8. The black 

fiducial marks around the insert were measured precisely on the model using a portable 

FaroArmTM coordinate measuring machine. These marks were used to calibrate the OFI images. 

OFI skin-friction measurements depend on accurate determination of the spacing of fringes 

generated by a sheared layer of oil. 

Figure 7.4-8. Boundary Layer Insert Mounted in Rough Heatshield 
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7.4.4 Mounting Strut Strain Gauges 

The upper mounting struts were strain gauged for bending to monitor the aerodynamic loads 

acting on the model during the test. This was primarily to ensure safety. Full-bridge gauges 

were located on both upper struts, upper and lower surfaces, near the junction with the lower 

struts, and on the model side of the cable attachment fairing. Figure 7.4-9 shows the strain gauge 

locations on the left upper strut. The flat ribbon cables from the gauges adjacent to the cable 

attachments run along the upper strut trailing edge to the junction with the lower struts. The 

cables run into the main cable bore in the lower struts where the rest of the instrumentation 

cables were located. 

Figure 7.4-9. Strain Gauges on Left Upper Strut 

7.4.5 Additional On-Board Instrumentation 

In addition to the electronically-scanned pressure module, an AoA sensor and a high-rate, 3-axis 

accelerometer were mounted inside the CM model. The AoA was measured using a Honeywell 

QA2000 accelerometer, which was mounted to the back shell in the machined pocket at the 

bottom in the CAD view as shown in Figure 7.3-3. An Omega® model ACC301 3-axis 

accelerometer was used to monitor model vibrations, primarily as a diagnostic tool in case the 

model vibrations were larger than expected. The vibration data was recorded with the unsteady 

pressure signals. 

7.4.6 IR Thermography 

Three IR cameras were used to image the temperature variation over the model from which a 

variety of flow features could be discerned. Two FLIRTM model 6800 cameras and a model 

7800 camera were used for the IR imaging. Germanium windows were mounted in the wind 

tunnel wall, floor, and ceiling to provide the necessary optical access (i.e., two existing window 

locations and one new window location). Vinci [refs. 9, 10] was used to optimize the camera 

and the new window locations and to eliminate potential conflicts with other cameras and 

lighting for PSP, PIV, and shadowgraphy. The IR images were stored in the FLIRTM examiner 

software format as individual tiff or jpeg images, enhanced for specific flow features. 

Strain gauges
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7.4.7 High-Speed Shadowgraph Imaging 

The wind tunnel shadowgraph system was used to image the flow during the test. The frames 

were acquired at a rate of 6,000 per second using a Photron SA1 1024x1024 pixel camera with a 

20 μm pitch. At every test condition, 80 individual images were acquired. The 80 images were 

made into a movie and a single, averaged image showing the mean flow structure. 

7.4.8 PIV Measurements 

The use of PIV for flow measurements has been well-established in laboratories and wind 

tunnels [ref. 16]. PIV system installation is more problematic in larger wind tunnels, particularly 

pressurized wind tunnels where optical access is not straightforward. The measurement plane 

location can increase the installation difficulty. For this experiment, there were two planes of 

interest: a horizontal cross-stream plane with a relatively small imaging area to look at the details 

of the wake shedding near the model, and a stream-wise vertical plane through the center of the 

wake with a large imaging area. The velocity in the smaller plane near the model was measured 

at a rate of 2,000 measurements per second. The velocity in the larger, vertical plane was 

measured at a rate of 2 measurements per second. These two PIV systems are referred to as the 

high-speed and wide-view systems, respectively. For both systems, the goal was to acquire at 

least 2,000 samples for each test condition. The number of samples was chosen to ensure 

convergence of the turbulence statistics on the high-speed system. 

The laser sheet for the wide-view system required installation of several optical elements inside 

the plenum. Figure 7.4-10 shows the laser path and the location of the optics required to form 

the vertical sheet at the center of the test section. The mirrors in the flow downstream of the 

slotted walls, on the sidewall, and on the vertical strut required careful design and installation to 

minimize vibrations. Vibration of those mirrors causes large motions of the laser sheet because 

of the distance from the mirrors to the image area. Custom enclosures were built to house these 

mirrors and to protect them from the airflow and contamination from the PIV seed material 

(i.e., atomized mineral oil). Figure 7.4-11 shows laser sheets for the wide-view and high-speed 

PIV systems. The final mirror in the wide-view laser delivery system was mounted to the 

vertical strut downstream of the test section. 
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Figure 7.4-10. Laser Sheet Paths in 11x11-ft Test Section 

Wide-view 

laser path

Cylindrical lens to 

form laser sheet

Test section 

walls

V∞

High-speed 

laser path

Mirrors in enclosures 

in tunnel air stream

Wind-tunnel 

primary diffuser

Vertical strut
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Figure 7.4-11. PIV Laser Sheets 

The horizontal laser sheet in Figure 7.4-12 was for the high-speed system. It entered the plenum 

through the opposite window from where the wide-view laser entered (i.e., right-side of the 

image in Figure 7.4-11). This sheet traveled through the plenum and across the test section into 

a beam dump to minimize stray reflections. 

The locations of the PIV measurement areas for both systems and model locations are shown in 

Figure 7.4-12. The greyed-out areas of the wide-view image areas were contaminated by laser 

reflections on the wind tunnel wall, which eliminated the ability to compute the flow velocity in 

that area. 

The wide-view PIV image area was captured by two Redlake EC11000 cameras with 4008x2672 

pixels and 9 μm pixel pitch. The object space resolution was ~4 pixels per millimeter. The laser 

sheet was generated by a Spectra-Physics PIV 400 laser tuned to deliver 400 mJ/pulse. 

 
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PIV data acquisition and processing was performed using La Vision DaVis versions 7.4 and 

8.1.4, respectively. The processing parameters included a 6-pixel high-pass filter, multi-pass 

correlation using 256 pixel windows initially, then finishing with 128x128 pixel, Gaussianweighted windows with 75 percent overlap. This window size was chosen because of the low 

signal-to-noise ratio of the images. The turbulence statistics did not converge when the data was 

processed on smaller windows. The resulting measurement grid was 262 by 154 points in the 

x- and z-directions, respectively. Each vector was measured in a time window of 

~32 milliseconds. The window deformation scheme interpolates and adjusts the vectors to 

capture gradients smaller than an interrogation window. 

Figure 7.4-12. PIV Imaging Areas for Both Model Positions 

The high-speed PIV data were acquired using Phantom 641 cameras with a sensor resolution of 

2560x1600 pixels and a pixel pitch of 10μm. The images were captured using only 2000x600 

pixels in order to achieve the 4,000 frames/second imaging required for 2,000 velocity 

measurements per second. The laser sheet was generated using two Quantronix Darwin Duo 

lasers providing 40 mJ/pulse. 

As with the wide-view measurements, La Vision DaVis software was used for data acquisition 

and processing. The high-speed data were processed using a 6-pixel high-pass filter, correlated 

with the multi-pass scheme starting with 64x64 pixel windows, and reduced to 32x32 pixel 

Gaussian-weighted windows with 75 percent overlap. The resulting measurement grid was 

257x132 points in the x- and y-directions, respectively. 

For both PIV systems, the instantaneous velocity measurements were archived with the velocity 

statistics. A summary of the computed statistical information is given in Table 7.4-1. 

 
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Table 7.4-1. Flow Statistics in Archived PIV Data Files 
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8.0 Test Conduct 

The test was run over a 3-week period from August 20, 2012, to September 6, 2012. An 

additional 3 weeks was spent in the model preparation room prior to installing the test article in 

the wind tunnel. Table 8.0-1 shows the runs accomplished during the test. The highest priority 

was to acquire all data at Mach 0.7 and 15-degree AoA, which represented close to the nominal 

drogue parachute deployment conditions. 

Table 8.0-1. List of Runs Accomplished During Test 

The test was divided into three phases. The first was to acquire PSP, IR, and unsteady pressure 

data with the rough and smooth heatshields. The original plan was to acquire those data types 

for AoA of 0, 15, and 30 degrees. The rough heatshield at 30 degrees AoA was tested first 

because previous testing of the Orion CM showed the flow was the most unsteady at 0 degrees 

AoA with a smooth (un-tripped) heatshield. The load limits of the test (monitored using the 

strain gauges on the upper struts) made it uncertain whether the team could test at 0 degrees 

AoA. Ultimately, time constraints on PIV lighting borrowed from the Air Force’s Arnold 
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Engineering and Development Center, and other aspects of the test (e.g., primarily installation 

and calibrating the strain gauges for aero loading) prevented completion of the planned test 

matrix for this test portion and no data was acquired for 0 degrees AoA with either heatshield. 

Testing of the smooth heatshield was only performed at 30 degrees AoA because of the same 

resource constraints. 

This portion of the test was run mostly at high Reynolds number that varied with Mach number, 

but some runs were done at lower Reynolds numbers, primarily to check out the model at lower 

dynamic pressure. These data proved to be instructive as described in the results section. 

The second test phase was relatively short and examined the heatshield boundary layer. With the 

boundary layer insert installed in the model, the NESC team did not acquire other data types 

except for IR imagery. Only the rough heatshield was tested and only at the high Reynolds 

number conditions. However, some data was acquired at a 0 degree AoA for Mach numbers of 

0.3, 0.5, and 0.7. 

In the third test phase, the PIV measurements took the longest to prepare and run. Throughout 

the first two portions of the test, refinements were made to the PIV system. Camera operation 

was checked out and laser sheet optics were adjusted on the off-shift to minimize the impact on 

the other portions of the test. Keeping the PIV hardware operating and in alignment proved to be 

more challenging than anticipated, which limited the conditions for which PIV data was 

acquired. The NESC team limited this portion of the test to only the rough heatshield and 

15 degrees AoA. For the initial model installation, near wake measurements, data was acquired 

at Mach numbers of 0.3, 0.5, and 0.7. The Mach 0.9 condition was not tested because it 

generated the most unsteady flow and vibration caused PIV system alignment issues. No data 

were obtained at Mach 1.03 since the seeding was poorly distributed across the test section 

because of the plenum evacuation required to reach supersonic conditions. For the second model 

position, far wake measurements, data was acquired at Mach numbers of 0.3, 0.5, and 0.7. 

The first two test phases acquired data for approximately half of the planned test conditions. The 

main omission was the 0 degree AoA conditions. The smooth heatshield data, as will be 

discussed, was contaminated by premature transition caused by surface flaws, so additional data 

would not have been helpful. However, the data acquired were instructive concerning the 

sensitivity of the boundary layer transition to the smoothness of the model surface. 

The PIV portion of the test was less successful in covering the planned test conditions. Only 

about 25 percent of the planned measurements were acquired. Given the lower-than-planned 

productivity, the data acquired was of good quality and covered the flight conditions for drogue 

deployment to a location in the wake close to the drogue parachute location (i.e., 5.5 versus 

7 diameters for the drogue parachutes). 

The following sections describe the results obtained from the various measurement techniques. 

In most cases, representative samples of the data are presented. The archived data is described in 

Appendices C through H, with Appendix C describing the files containing the test conditions and 
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static pressure data. The complete data set is available on the test website on aeroCOMPASS. 

The data can be made available on hard disk. 

8.1 IR Thermography 

The IR thermography worked well during the test providing details of boundary layer transition 

and flow separation. The images showed the suction peak on the heatshield as an area of cooler 

surface temperature, which had not been previously visualized in the AUPWT. The IR data was 

archived as tiff or jpg images of individual runs from the three IR camera views, and as FLIRTM

proprietary files that contain a few seconds of IR video that can be interrogated using FLIRTM 

ExaminIRTM software. Appendix C provides a description of the archived IR images. 

8.1.1 Rough Heatshield IR 

Figure 8.1-1 shows an example of an IR image taken at M 0.7, 30 degrees AoA, and Reynolds 

number 10x106

. This image covers a temperature range that shows the flow details around the 

heatshield shoulder so the stagnation point indicated is not shown. The attached, accelerating 

flow gradually decreases in temperature as the flow accelerates approaching the shoulder. The 

much cooler flow at the maximum velocity region (i.e., lowest pressure or maximum suction) is 

visible as a dark line in the image. The dark line (i.e., ring around the model) is followed by a 

lighter area showing the flow slows after the suction peak. The separation line is apparent as an 

abrupt decrease in temperature, shown as the black area at the beginning of the back shell (but 

still part of the heatshield model part with the plastic overlay) in the image. The aluminum back 

shell was at a different temperature during the testing and the joint between the model materials 

shows up as an abrupt increase in temperature. The plastic fairings over the model strut 

attachment points had a temperature similar to the post-separation heatshield as shown by the 

dark circle around the strut penetrations through the back shell. The transition strips that were 

applied to the support struts (e.g., lower surface of the left-hand strut) are visible. The separated 

shear layer can be seen passing over the right-hand upper strut as a bright band with darker 

(i.e., colder) temperature on the strut further toward the model. 
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Figure 8.1-1. IR Image of the Rough Heatshield at Mach 0.7, AoA 30 degrees, and Reynolds 

Number 107 

An unexpected result was the ability to see the individual Kulite® pressure transducers. The 

transducers generate a small amount of heat when energized which makes them visible as light 

dots. The IR images with the Kulites® turned off and the overall flow patterns were unchanged 

so for most of the test the sensors were left energized. 

Figure 8.1-2 shows a comparison between the pressure distribution and the temperature 

distribution on the model at Mach 0.3, 15 degrees AoA, and Reynolds number 5.3x106

. The two 

images show the same flow patterns. The IR and PSP show the stagnation point and suction 

peak while the IR image shows the separation point more clearly than the PSP. The local flow 

disturbances caused by the Kulite® inserts shows more clearly in the IR image. 

Changes in flow patterns with Mach number are shown in Figure 8.1-3. The images were scaled 

to a temperature range highlighting the flow at the heatshield perimeter. The differences are 

subtle with the largest effect being the diminishing disturbance to the flow by the Kulite® inserts 

with increasing Reynolds number. There is an abrupt change in the extent of the cooler suction 

peak band and an upstream shift in the separation line as Mach number increases from 0.7 to 0.9. 

This is likely due to the formation of a shock at the higher Mach numbers, which tends to force 

the separation location. There will be more discussion of this effect in the sections on the 

unsteady pressure measurements and Schlieren/shadowgraph images. 
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Figure 8.1-2. PSP Image Compared to IR Image with Temperature Range to Show Stagnation Point; 

Rough Heatshield at Mach 0.3, AoA of 15 degrees, and Reynolds Number 5.3x106 

Figure 8.1-3. IR Image Showing Changes in Flow Structure with Increasing Mach Number 

M 1.07, ReD = 7.3x106 M 0.9, ReD = 107

M 0.7, ReD = 107 M 0.5, ReD = 8.7x106 M 0.5, ReD = 5.3x106
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8.1.2 Effect of Reynolds Number on Rough Heatshield 

Given the roughness pattern that was machined into the plastic heatshield surface, it was 

assumed the boundary layer would transition to turbulent near the stagnation point. The IR 

images obtained at the lower Reynolds number (i.e., 1.33x106

) were surprising. Figure 8.1-4 

shows IR images for two runs at the same test conditions (i.e., Mach 0.5 and AoA 30 degrees). 

The dark to light patterns are nearly identical for the two runs performed several hours apart. 

The patterns do not correspond to any physical pattern on the rough heatshield. There were 

variations in the depth of the hex pattern on various parts of the heatshield, particularly nearer 

the stagnation point for this test condition. The note in Figure 8.1-4 points to a seam in the hex 

pattern. To the right of the seam, the depth of the pattern is less than the depth to the left of the 

seam. Figure 8.1-5 shows the difference in roughness depth across the seam at the opposite end 

of the seam. The IR cameras could see the difference in the hex-pattern depth as a difference in 

temperature. It was not clear whether the resulting dark-light (i.e., colder-warmer) pattern in the 

IR images were a result of a bias in transition location due to the roughness difference, 

particularly with the expectation of fully, or nearly so, turbulent flow on the rough heatshield. 

There was a small area of the model where the PSP did not stick to the rough heatshield that is 

visible in Figure 8.1-5. This may be an area where some residual machining lubricant 

(detergent) was left on the surface preventing good adhesion of the base coat of paint. 

Figure 8.1-4. IR Images of Heatshield taken with the Ceiling Mounted Camera at Mach 0.5, AoA of 

30 degrees, and Reynolds Number 1.33x106 

Machining seam 

in hex pattern
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Figure 8.1-5. Difference in Roughness Depth Across Machining Seam 

Note: dimensions in inches. 

8.1.3 Smooth Heatshield 

The smooth heatshield at low Reynolds number showed a complicated pattern of transition that 

was nearly always traceable to a surface defect. Figure 8.1-6 shows a top-view IR image of the 

smooth heatshield at Mach 0.7. Laminar flow is present where the temperature is lower 

(i.e., darker) with turbulent flow in the warmer (i.e., lighter) areas. Figure 8.1-7 shows the same 

test point taken with the side-view camera. Both images show that where the boundary layer 

remains laminar to the shoulder, the flow separates immediately downstream of the suction peak 

with little pressure recovery. 

Figure 8.1-6. Top-view IR Image of Smooth Heatshield; M = 0.7, AoA of 30 degrees, and Reynolds 

Number 10x106
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Figure 8.1-7. Side-view of IR Image; M = 0.7, AoA of 30 degrees, and Reynolds Number 10x106

In addition to identifying specific locations where transition occurs, the images show the effect 

of surface discontinuities around the Kulite® inserts. Where the insert edges were not flush with 

the adjacent model surface, the flow generally separates prematurely. 

The flaws on the smooth heatshield were identified visually and by touch. Some are noted in 

Figure 8.1-7, while the complete map of defects is shown in Figure 8.1-8. The boundary layer 

was sensitive to surface irregularities and in some instances flush 0.020 inch pressure tap holes. 
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Figure 8.1-8. Map of Defects on Smooth Heatshield 

At low Reynolds numbers the boundary layer behavior was significantly different. The flow 

remained laminar for all test conditions to the suction peak around the heatshield shoulder. 

Figure 8.1-9 shows an IR image of a low Reynolds number case at Mach 1.02. The flow 

accelerates to the shock location, where it separates. This is typical of the flow behavior at 

higher Mach numbers, particularly for the smooth heatshield. 

~0.02”

~0.02”
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Figure 8.1-9. IR Image of Smooth Heatshield Mach 1.02, AoA of 30 degrees, and Reynolds Number 

1.33x106

8.2 PSP Measurements 

In general, the PSP data was of high quality. Since it is three-dimensional by its nature, some 

sample results will be shown. The data archive is described in Appendix D. 

8.2.1 Rough Heatshield Results 

Figure 8.2-1 shows the time averaged pressure coefficient plotted on the model and strut surfaces 

for $2$of 30 degrees and Mach numbers of 0.3, 0.7, and 1.07. For the Mach 0.3 case, the 

pressure coefficient over the surface was noticeably noisier than the other two cases. This was a 

general finding with PSP, that the data quality is better at high static and dynamic pressure. The 

noise in the data could have been reduced by sampling longer, but the data was deemed 

sufficiently converged as acquired. The noise diminishes as the flow accelerates toward the 

heatshield shoulder. All of the higher Mach number data was clean and represents the state-ofthe-art for PSP measurements. Figure 8.2-1 shows the flow was separated over the entire back 

shell at all Mach numbers, although the separation location was difficult to pinpoint in these 

color maps. Also apparent in the figure are the pressure data on the upper struts. The PSP was 

applied to the full length of the upper struts and most of the lower strut surfaces to ensure CFD 

validation efforts have good documentation of the aerodynamic surfaces present in the test. 

Run 70
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Figure 8.2-1. Pressure Distributions on CM at AoA of 30 Degrees 

8.2.2 Integrated Forces and Moments 

Since the PSP technique provides pressure over the entire model surface, integration of the 

pressure to obtain aerodynamic forces and moments was possible. This has been done 

successfully, but is less accurate and repeatable than a balance measurement [ref. 8]. Since there 

was no balance in this model, integrating the pressure was the only way to get the aerodynamic 

information. Table 8.2-1 shows the results of these integrations. The moment reference point is 

at the heatshield apex. 

Run 44 Run 46

Cp

Run 48

M 0.3, Re Cp

D = 5.3 x 106 M 0.7, ReD = 10 x 106 M 1.07, ReD = 7.3 x 10 Cp 6
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Table 8.2-1. Integrated Forces and Moments from Surface Pressures 

Mach Alpha 

Re 

(x10-6) 

Smooth

/Rough CA CY CN CPM CLN Run # 

0.70 29.2 1.38 R 0.922 0.011 0.025 -0.075 -0.001 36 

0.30 29.2 1.35 R 0.931 0.011 0.026 -0.076 -0.001 42 

1.11 29.28 1.65 R 1.225 0.011 0.050 -0.073 -0.002 43 

0.30 29.19 5.33 R 0.796 0.016 0.034 -0.059 -0.003 44 

0.50 29.19 8.67 R 0.858 0.003 0.034 -0.067 -0.001 45 

0.70 29.19 10 R 0.932 0.002 0.029 -0.071 -0.002 46 

0.90 29.36 10.01 R 1.064 0.003 0.035 -0.071 -0.003 47 

1.07 29.46 6.64 R 1.223 0.006 0.054 -0.074 -0.003 48 

0.30 14.05 5.33 R 0.858 0.009 0.074 -0.038 0.001 54 

0.50 14.22 8.67 R 0.919 0.004 0.029 -0.036 -0.001 55 

0.70 14.31 10 R 1.005 0.002 0.011 -0.035 -0.002 56 

0.90 14.39 10.01 R 1.105 0.004 0.004 -0.034 -0.003 57 

1.07 14.57 6.62 R 1.265 0.004 0.019 -0.037 -0.002 58 

0.70 14.36 10 R 0.993 0.002 0.009 -0.035 -0.002 60 

0.70 29.15 1.37 S 0.921 -0.002 0.052 -0.074 0.000 69 

1.02 29.21 1.61 S 1.171 0.002 0.066 -0.076 -0.001 70 

0.30 29.05 5.33 S 0.812 -0.016 -0.092 -0.062 0.001 71 

0.50 29.01 8.67 S 0.895 0.004 -0.116 -0.060 -0.006 72 

0.70 28.95 10 S 0.989 0.005 0.076 -0.063 -0.004 73 

0.90 29.19 10.02 S 1.096 0.007 0.012 -0.065 -0.004 74 

1.08 29.41 6.64 S 1.212 0.005 0.043 -0.067 -0.003 75 

0.50 28.9 8.66 S 0.896 0.003 -0.118 -0.058 -0.004 76 

Note: Alpha is AoA and Re is Reynolds number. 

The integrated axial force coefficient is shown in Figure 8.2-2(a) for the rough heatshield. The 

results are shown for AoAs of roughly 15 and 30 degrees. Figure 8.2-2(b) shows the axial force 

coefficient for the rough and smooth heatshields as a function of Mach number at 30 degrees 

AoA. The rough heatshield produces less axial force than the smooth heatshield because of the 

pressure recovery after the suction peak that does not occur with the smooth heatshield. 
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a) Rough heatshield at AoA of 15 and 30 degrees (b) Rough and smooth heatshields at AoA of 30 

degrees 

Figure 8.2-2. Axial Force Coefficient Variation with Mach Number 

8.3 PIV Measurements 

The PIV measurements of the wake velocity distributions were the primary goal of this test and 

provide insight into the mean flow and turbulence. In addition, the high-speed PIV data in the 

near wake gives time-accurate velocity data for the lower Mach numbers from which frequency 

information was computed. The PIV data archive is described in Appendix E. 

Figure 8.3-1 shows the mean velocity distribution in the wake from the wide-view PIV system. 

Velocity fields are shown with the model in the upstream and downstream locations. The cutout 

in the lower left of the figure shows the model location relative to the measurements. The 

velocity measurements were continuous between the two measurement areas, in spite of the gap 

between them. The gap was caused by an area of reflected laser light on the wind tunnel wall, 

which was visible in one of the cameras. The image contamination from the reflected light 

meant the velocities could not be computed. A smaller reflection is noted in Figure 8.3-1, which 

caused a bias in the computed velocity in that region. 

 

α = 30°

α = 15°

Rough Heat Shield

Rough

Smooth

α = 30°

Rough Heat Shield

Smooth Heat Shield
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Figure 8.3-1. Average Velocity in CM Wake; Mach 0.7, AoA of 14.0 degrees, and Reynolds Number 

10x106 Color Contours of Vx, Every 10th Vector Shown 

The root mean square (RMS) of Vx is shown in Figure 8.3-2 for Mach 0.7, AoA of 14.0 degrees, 

and Reynolds number 10x106

. At this test condition, a total of 4,000 image pairs were acquired 

and the turbulence statistics (e.g., RMS, turbulent kinetic energy, etc.) were nearly identical 

when computed with either 2,000 or 4,000 velocity measurements. The distribution of Vx RMS 

was consistent between the two measurement stations in spite of the measurements being taken 

on different days at the two locations. 

One frame of instantaneous velocity data from the high-speed PIV system is shown on the left 

side of Figure 8.3-3. Vectors show the in-plane velocity (i.e., Vx and Vy) while the color 

contours show the velocity normal to the plane (i.e., Vz). V∞ for the measurements in 

Figure 8.3-3 is 102 meters per second (m/s), so the instantaneous out-of-plane velocity can be a 

significant fraction of free-stream. 
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Figure 8.3-2. RMS of Vx in CM Wake; Mach 0.7, AoA of 14.0 degrees, and Reynolds Number 10x106

The right-side of Figure 8.3-3 shows the time-averaged velocity vectors for the same test 

conditions as on the left side, with color contours of Vx. The location of the high-rate PIV 

measurement plane was selected to capture the shear layer at Mach 0.7 at 15 degrees AoA in the 

near wake. The location was based on a Reynolds-averaged Navier-Stokes computation of the 

flow around an isolated capsule. The measurements in Figure 8.3-3 show the measurement area 

may have missed getting all the way outside the shear layer, but most of the shear layer is 

resolved at Mach 0.7. 

The Vx RMS is shown in Figure 8.3-4 for the same test conditions. Figure 8.3-4 confirms the 

measurement area did not reach the inviscid outer flow since the Vx RMS reached a minimum of 

8 percent of the free-stream velocity at the edge of the measurement area. 
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Figure 8.3-3. Velocity Measurements from High-speed PIV; Mach 0.3, AoA of 14.0 degrees, Reynolds 

Number 5.3x106

, and V∞ = 103 m/s 

Note: Instantaneous velocity on left, average velocity on right. 

Figure 8.3-4. RMS X-velocity Measurements from High-speed PIV; Mach 0.3, AoA of 14.0 degrees, and 

Reynolds Number 5.3x106
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The information needed for the parachute load estimates is the dynamic pressure at the 

appropriate distance downstream of the capsule. Figure 8.3-5 shows the measured velocity along 

the center of the wake from the PIV measurements at Mach 0.7 and AoA of 30 degrees. At 

~5.1 diameters downstream, the wake velocity was relatively low (i.e., ~145 meters per second 

(m/s)) compared to the free-stream value of 235 m/s. The velocity fluctuations were large in the 

wake with an RMS level near 50 m/s (Figure 8.3-2). 

Figure 8.3-5. Distribution of Vx Along Center of Wake at Mach 0.7, AoA of 13.9 degrees, and Reynolds 

Number 107 

Note: Red line connects data collected with model in aft position to that collected in the forward position. 

Figure 8.3-6 shows time-averaged velocity profiles through the wake at several downstream 

locations. The profile nearest the model was beyond where the minimum mean axial velocity 

was no longer negative. At that location the PIV measurement plane extended just outside of the 

shear layer and the flow acceleration due to wind tunnel blockage and streamline curvature is 

evident (i.e., velocity greater than free-stream). The y/D value in the figure is referenced to the 

minimum velocity point at the given axial station rather than from a physical location in the wind 

tunnel. 

-0.50

-0.25

0.00

0.25

0.50

0.75

1.00

0.0 1.0 2.0 3.0 4.0 5.0 6.0

Vx/V∞

x/D
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Figure 8.3-6. Variation in Total Velocity at Various Downstream Locations; Mach 0.7, AoA of 13.9 

degrees, and Reynolds Number 107 

Note: Velocity normalized by the free-stream velocity, yCL/D is referenced to the minimum velocity point in the 

wake at a given x/D location. 

The effect of the wake unsteadiness on the load history of a parachute in the wake is difficult to 

establish from the model data. Figures 8.3-7 and 8.3-8 show instantaneous velocity fields far 

downstream of the model with a line approximating the parachute diameter and position centered 

on the wake. The inflated Orion drogue parachute is approximately the same diameter as the 

CM. Figure 8.3-7 shows a velocity field that is relatively low at the parachute location while 

Figure 8.3-8 is at an instant showing high velocity at the parachute location. 
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Figure 8.3-7. Contours of Vx in Far-field Wake (Mach 0.7, AoA of 15 degree) at an Instant Showing 

Low Velocity Across Approximate Drogue Parachute Diameter, V∞ = 234 m/s 

Figure 8.3-8. Contours of Vx in Far-field Wake (Mach 0.7, AoA of 15 degrees) at an Instant Showing 

High Velocity Across Approximate Drogue Parachute Diameter, V∞ = 234 m/s 
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Knacke defines the drag loss coefficient (DLC) to estimate the loads acting on a parachute in the 

wake of another body (Figure 21 in reference 1). It is determined by measuring the parachute 

drag in a uniform free-stream and in the wake of the other body. The ratio of the value in the 

wake to that in the uniform free-stream is the DLC: 

D freeair

Dwake

C

C

DLC 

If the parachute shape does not change significantly when in the wake, then the DLC is an 

indirect measure of the relative integrated effective dynamic pressure acting on the parachute in 

the wake to that experienced by the parachute in free air. An approximation of the effective 

dynamic pressure acting on a parachute can be made by integrating the dynamic pressure in the 

flowfield over the parachute plan area. Assuming the flow on the parachute upper and lower 

halves are axi-symmetric around the center of the parachute, the DLC can be approximated as: 

ydy V

V y DLC

y r

y r

3

 

  f

| 2

2

( )

The DLC can be calculated instantaneously or using time-averaged velocities. The DLC as 

reported in reference 1 is a time-averaged quantity. It was computed at a variety of downstream 

locations from the current velocity data at Mach 0.7. Figure 8.3-9 shows the results of these 

integrations for available locations in the current data set. It is clear from the comparison with 

the Dp/DB ~1 curve that integrating the current results in lower values of DLC than in the 

literature. One reason is the bodies for most of the data referenced in the figure were obtained 

with a parachute behind and body with a higher fineness ratio than the capsule. The Apollo 

drogue chute behind the CM was an exception and it lies on the curve-fit to the rest of the data 

plotted by Knacke. It is not clear how to account for this discrepancy, but the approximation of 

axi-symmetric flow from the planar data set and the asymmetry apparent in the current data 

could be part of the issue. Regardless, it appears using the curves in Knacke would be 

reasonable. 

Using the instantaneous velocity fields shown in Figures 8.3-6 and 8.3-7 to compute DLC gives 

a range of fluctuating loads that the parachute might experience. For the low-velocity instance, 

the DLC is 0.35 while at the high-velocity instance it is 0.79. The two velocity measurements do 

not necessarily represent the highest and lowest integrated loads. These measurements were 

selected as representative high- and low-velocity from the first 100 of the 4,000 velocity 

instances measured at this test condition. 
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Figure 8.3-9. DLC from Current Study Compared to Results in Reference 1 

Note: The red symbols should be compared with the lower curve, Dp/DB ~1.0. Should be most comparable to the 

Apollo drogue chute behind CM (symbol II, Figure scanned from reference 1). 

8.4 Unsteady Pressure Measurements 

The flow over bluff bodies is generally unsteady and the unsteady pressure measurements 

provide insight into the character of the unsteadiness. The unsteady pressure data archive is 

described in Appendix F. Observing the model motion and unsteady strain gauge readings gave 

a good indication of the unsteadiness. In general, the model shook more at lower AoAs at a 

given Mach number than at higher AoAs. The unsteadiness increased with Mach number to 

above Mach 0.7. Above Mach 0.7, the flow separation seemed to be fixed by the shocks at the 

heatshield shoulder, which greatly reduced the model dynamics. When the flow was supersonic, 

the unsteadiness was small. The upper strut strain gauge readings became steadier and the model 

motion was almost completely eliminated. These observations were consistent with the unsteady 

pressure measurements. At supersonic conditions the pressure fluctuations were low and 

increased at Mach numbers below 0.9. 

Integrated over drogue chute area 
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The unsteady pressure transducers provided quantitative information that corroborated the visual 

observations. Figure 8.4-1 is a map showing the 1/3rd octave sound pressure levels (SPLs) for 

the heatshield transducers on the rough heatshield at Mach 0.7, AoA of 29.25 degrees, and 

Reynolds number 10x106

. The color of each box in the figure represents the 1/3-octave SPL for 

a given sensor with the sensor numbers on the left-side of the SPL map. There are four groups of 

10 Kulites® and the transducers in each group of 10 are located in the inserts as shown in the 

heatshield sketch on the left-side of the figure. The level of unsteadiness is an indicator of flow 

separation as noted in Figure 8.4-1. The SPL for sensors upstream of the flow separation is 

relatively low (i.e., green or blue) while at, or slightly downstream of separation, the levels are 

significantly higher (i.e., orange or red). The separation locations determined from the unsteady 

pressure levels match well with the locations visualized in the IR images. 

Figure 8.4-1. 1/3-octave SPL Levels for Unsteady Pressures; Mach 0.7, AoA of 29.25 degrees, and 

Reynolds Number 10x106

The dominant shedding frequency of the model is apparent in the SPL map in Figure 8.4-1. The 

levels in the 100 to 125 Hertz (Hz) frequency range are higher on almost all the sensors relative 

to the adjacent frequency bins. Figure 8.4-2 shows the narrow-band spectra of two of the sensors 

(i.e., K02 and K16) for sensors at the same free-stream conditions as in Figure 8.4-1. Two sets 
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of curves are shown for the smooth and rough heatshields. The Strouhal number (St) for the 

rough heatshield is ~15 percent lower than for the smooth heatshield, which is defined as: 

where f is the shedding frequency, D is the characteristic width of the body (diameter in this 

case), and V is the free-stream velocity. 

The roughness increases the effective model diameter in regard to the dominant shedding 

frequency. A decrease in St for a circular cylinder with roughness is documented in reference 

17, which reported up to a 12 percent decrease in St using wire mesh attached to the model 

surface for roughness. While the boundary layer turbulence generated by a wire mesh may differ 

from that due to a hexagonal dimple pattern, the current findings are consistent with those in 

reference 17. 

Figure 8.4-2. Narrow Band SPL Levels Showing Effect of Roughness on Shedding Frequency; Mach 

0.7, AoA of 29.25 degrees, and Reynolds Number 10x106

The Mach number has an effect on the St. Figure 8.4-3 shows the Mach number effect for the 

CM at 30 degrees AoA. The Reynolds number is not the same at all Mach numbers, but the 

change in shedding is striking at Mach 1.07 where the shedding is weaker than the subsonic 

St  fD

V
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Mach numbers. The frequency was higher for the supersonic case where St = 0.33 versus 

~0.2 for the subsonic/transonic conditions. 

Figure 8.4-3. Narrow Band SPL Levels Showing Effect of Mach Number on Shedding Frequency; 

Mach 0.7, AoA 29.25 degrees, and Reynolds Number 10x106

Time correlations between various sensors were examined to ascertain whether the pattern of 

flow separation dynamics could be discerned. Figure 8.4-4 shows the results of time correlations 

on the rough heatshield for the Mach 0.7, 30 degrees AoA, and high Reynolds number condition. 

In Figure 8.4-4 the reference signal comes from K8, at the top of the model and is located in the 

separated flow. K19, K29, and K39 are 180, 90, and 30 degrees from the reference transducer, 

respectively. Following the logic in the figure with K39, K29, and K19 lagging K8 in that order, 

the case can be made for a spiral separation pattern moving clockwise looking at the heatshield. 

Additional analysis and/or testing is required to verify or refute this conjecture. 
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Figure 8.4-4. Correlation Coefficient versus Time Referenced to Sensor K8; Mach 0.7, AoA of 29.25 

degrees, and Reynolds Number 10x106

8.5 Boundary Layer Measurements 

Boundary layer measurements were made on the heatshield for a limited set of conditions and 

outlined in Table 5.2-1. All of the measurements were made on the rough heatshield because of 

time constraints during the test. The boundary layer data archive is described in Appendix G. 

8.5.1 Boundary Layer Surveys 

Boundary layer profiles were documented at the traversing probe location (x = 0.205 inch, 

y = 0 inch, and z = 2.800 inch) from the heatshield apex. The insert is smooth and the length of 

smooth surface from the edge of the insert to the probe tip is 0.5 inch. Figure 8.5-1 shows 

boundary layer profiles plotted as local Mach number versus distance from the model surface. 

The plot shows the effect of AoA at free-stream Mach 0.7 with an additional profile for Mach 

1.05 at 29.4 degree AoA. In general, increasing the AoA increases the boundary layer thickness 

due to the longer run from the stagnation point to the measurement location. The Mach number 

at the edge of the boundary layer increased with AoA at the measurement station. Increasing the 

Mach number from 0.7 to 1.05 at 29.4 degrees AoA did not change the boundary layer thickness. 

This change increased the local Mach number at the boundary layer edge. 
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Figure 8.5-1. Boundary Layer Profiles for Rough Heatshield 

The skin friction was measured using OFI for three flow conditions at 15 degrees AoA: Mach 

0.3 at Reynolds number 5.3x106, Mach 0.7 at Reynolds number 10x106, and M 1.05 at Reynolds 

number 6.6x106

. Figure 8.5-2 shows the oil fringes with the resulting skin friction values for the 

Mach 0.7 condition. These three OFI skin friction values were used to calibrate the Stanton 

gauge, located adjacent to the probe. The Stanton gauge provided the skin friction measurements 

for the rest of the test conditions for which boundary layer profiles were acquired. Limiting the 

number of OFI skin friction measurements was done to minimize wind tunnel test time as each 

oil film measurement required approximately an hour of testing. The location of the probe 

relative to the rough heatshield surface is shown in Figure 8.5-2 as are the locations of the static 

pressure tap and the Stanton gauge adjacent to the probe. The dots surrounding the insert are 

fiducial marks allowing precise registration of the image relative to the model surface. The dots 

provide accurate scaling in the image from which the fringe spacing can be determined. 
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Figure 8.5-2. Example of Oil Film Images for Measuring Skin Friction; Mach 0.7, AoA of 15 degrees, 

and Reynolds Number 10x106

, Skin Friction Coefficient = 0.0026 directed 4.6 degrees 

from Vertical 

9.0 Findings, Observations, and NESC Recommendations 

9.1 Findings 

The following findings were identified: 

F-1. The heatshield roughness has a significant effect on the capsule aerodynamics 

(e.g., smooth heatshield drag is up to 8 percent higher than for the rough heatshield). 

F-2. The fundamental unsteadiness of the flow around a capsule precludes the use of nontime-accurate computational tools to determine the wake deficit. 

F-3. The unsteadiness increases as the AoA is decreased from 30 to 0 degrees. 

F-4. Handbook values for the wake deficit behind capsules at subsonic/transonic Mach 

numbers are similar to the results of the current test. The handbook values appear to be 

conservative for trailing distances less than ~6 capsule diameters with no evidence the 

trend changes at larger downstream distances. 

F-5. Wake motion and velocity fluctuation are not well represented in the handbook 

information used for parachute load estimates, but these qualities could be important 

when designing for maximum loads. 
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F-6. Large unsteadiness in the near wake could lead to unexpectedly large instantaneous loads 

on inflating parachutes. 

9.2 Observations 

The following observations were identified: 

O-1. The St for the capsule shape is close to 0.2 for Mach numbers below 0.9. For supersonic 

free-stream, the shedding is less periodic with the lowest dominant frequency at a St of 

0.4. 

O-2. Large velocity fluctuations are seen in the CM model wake. In the near wake, the RMS 

of Vx exceeds 20 percent of the free stream velocity. The eddies shed from the capsule 

are large, particularly at Mach numbers of 0.3, 0.5, and 0.7. The shadowgraph videos 

show at a distance of 2 capsule diameters downstream of the model, the eddies can travel 

laterally ±0.25D from the mean shear layer location. 

9.3 NESC Recommendations 

The following NESC recommendations were identified and directed towards the MPCV CPAS 

project unless otherwise identified: 

R-1. Include the effect of physically scaled heatshield roughness in wind tunnel tests. (F-1)

R-2. Perform computational modeling of bluff-body aerodynamics in a time-accurate manner. 

(F-4, F-5)

10.0 Alternate Viewpoint 

There were no alternate viewpoints identified during the course of this assessment by the NESC 

team or the NESC Review Board (NRB) quorum. 

11.0 Other Deliverables 

No unique hardware, software, or data packages, outside those contained in this report, were 

disseminated to other parties outside this assessment. 

12.0 Lessons Learned 

No applicable lessons learned were identified for entry into the NASA Lessons Learned 

Information System (LLIS) as a result of this assessment. 

13.0 Recommendations for NASA Standards and Specifications 

No recommendations for NASA standards and specifications were identified as a result of this 

assessment. 
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14.0 Definition of Terms 

Corrective Actions Changes to design processes, work instructions, workmanship practices, 

training, inspections, tests, procedures, specifications, drawings, tools, 

equipment, facilities, resources, or material that result in preventing, 

minimizing, or limiting the potential for recurrence of a problem. 

Finding A relevant factual conclusion and/or issue that is within the assessment 

scope and that the team has rigorously based on data from their 

independent analyses, tests, inspections, and/or reviews of technical 

documentation. 

Lessons Learned Knowledge, understanding, or conclusive insight gained by experience 

that may benefit other current or future NASA programs and projects. The 

experience may be positive, as in a successful test or mission, or negative, 

as in a mishap or failure. 

Observation A noteworthy fact, issue, and/or risk, which may not be directly within the 

assessment scope, but could generate a separate issue or concern if not 

addressed. Alternatively, an observation can be a positive 

acknowledgement of a Center/Program/Project/Organization’s operational 

structure, tools, and/or support provided. 

Problem The subject of the independent technical assessment. 

Proximate Cause The event(s) that occurred, including any condition(s) that existed 

immediately before the undesired outcome, directly resulted in its 

occurrence and, if eliminated or modified, would have prevented the 

undesired outcome. 

Recommendation A proposed measurable stakeholder action directly supported by specific 

Finding(s) and/or Observation(s) that will correct or mitigate an identified 

issue or risk. 

Root Cause One of multiple factors (events, conditions, or organizational factors) that 

contributed to or created the proximate cause and subsequent undesired 

outcome and, if eliminated or modified, would have prevented the 

undesired outcome. Typically, multiple root causes contribute to an 

undesired outcome. 

Supporting Narrative A paragraph, or section, in an NESC final report that provides the detailed 

explanation of a succinctly worded finding or observation. For example, 

the logical deduction that led to a finding or observation; descriptions of 

assumptions, exceptions, clarifications, and boundary conditions. Avoid 

squeezing all of this information into a finding or observation 
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15.0 Acronyms List 

μm Micrometre 

A&M Agricultural and Mechanical 

AMA Analytical Mechanics Associates, Inc. 

AoA Angle of Attack 

ARC Ames Research Center 

AUPWT Ames Unitary Plan Wind Tunnel 

CAD Computer-aided Design 

CEV Crew Exploration Vehicle 

CFD Computational Fluid Dynamics 

CM Crew Module 

CPAS Capsule Parachute Assembly System 

DLC Drag Loss Coefficient 

FML Fluid Mechanics Laboratory 

ft Foot 

Hz Hertz 

IR Infrared 

LaRC Langley Research Center 

m/s Meters per Second 

MPCV Multi-Purpose Crew Vehicle 

NESC NASA Engineering and Safety Center 

NRB NESC Review Board 

OFI Oil-Film Interferometry 

PIV Particle Image Velocimetry 

psid Pounds per Square Inch 

PSP Pressure Sensitive Paint 

RANS Reynolds-averaged Navier-Stokes 

Re Reynolds Number 

RMS Root Mean Square 

SDS Standard Data System 

SLS Selective Laser Sintering 

SPL Sounds Pressure Level 

St Strouhal Number 

15.1 Nomenclature 

AF Axial Force, acting in the x-direction in the model-fixed coordinate system 

CA Axial Force Coefficient, CA  AF

q
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Re Reynolds Number, Re  VD



M Mach number 

Ps(t) Static Pressure at time = t, Ps  t  Ps  Ps

(t)

Ps Time-averaged Static Pressure 

Ps

(t) Fluctuating component of Static Pressure 

P∞ Time-averaged free-stream Static Pressure 

q Free-Stream Dynamic Pressure 

t Time, seconds 

V Free-stream Velocity 

Vx x-component of local velocity 

Vy y-component of local velocity 

Vz z-component of local velocity 

x Stream-wise coordinate, wind-tunnel axes 

y Cross-stream, lateral coordinate, wind-tunnel axes 

z Cross-stream, vertical coordinate, wind-tunnel axes 

 Free-stream air density 

 Free-stream kinematic viscosity 

16.0 References 

1. Knacke, T. W., Parachute Recovery Systems Design Manual, Para Publishing, January 

1992. 

2. Braslow, A. L., and Knox, E. C., “Simplified Method for Determination of Critical Height 

of Distributed Roughness Particles for Boundary Layer Transition at Mach Numbers from 

0 to 5,” NACA TN 4363, 1958. 

3. Braslow, A. L., Hicks, R. M., and Harris, R. V., Jr., “Use of Grit-Type Boundary-LayerTransition Trips on Wind-Tunnel Models,” NASA TN D-3579, 1966. 

4. Amaya, M. A., and Boone, A. R., “Calibration of the 11x11-Foot Transonic Wind Tunnel 

at the NASA Ames Research Center,” AIAA 41st AIAA/ASME/SAE/ASEE Joint 

Propulsion Conference and Exhibit, AIAA paper 2005-4277, Tucson, AZ, July 10-13, 

2005. 

5. Sengupta, A., et al, “Performance of a Conical Ribbon Drogue Parachute in the Wake of a 

Subscale Orion Command Module,” 2012 IEEE Aerospace Conference, Big Sky, MT, 

March 3-10, 2012. 

6. Brauckmann, G. J., “Final Report for Test 89-CA: High-Reynolds Number Test of the 

Orion Crew Module in the NASA LaRC National Transonic Facility,” MPCV 

Aerosciences Report EG-CAP-12-65, September 2012. 
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7. Alunni, A. I., Olson, M. W., Gökçen, T., and Skokova, K. A., “Comparisons of Surface 

Roughness in Laminar and Turbulent Environments for Orion Thermal Protection System,” 

AIAA paper 2011-3776, 42nd AIAA Thermophysics Conference, Honolulu, HI, 

June 27-30, 2011. 

8. Sellers, M. E., “Demonstration of a Temperature-Compensated Pressure Sensitive Paint on 

the Orion Launch Abort Vehicle,” AIAA Paper 2011-3166, AIAA Applied Aerodynamics 

Conference, Honolulu, HI, June 2011. 

9. Driver, D. M., “Application of Oil-Film Interferometry Skin-Friction Measurements to 

Large Wind Tunnels,” Experiments in Fluids, vol. 34, pp. 717-725, 2003. 

10. Naughton, J. W., and Liu, T., “Photogrammetry in Oil-Film Interferometry,” AIAA 

Journal, vol. 45, no. 7, pp. 1620-1629, July 2007. 

11. Bradshaw, P. and Gregory, N., “The Determination of Local Turbulent Skin Friction from 

Observations in the Viscous Sub-Layer,” Aeronautical Research Council Report R&M No. 

3202, Her Majesty’s Stationery Office, London, March 1959.

12. East, L. F., “Measurement of Skin Friction at Low Subsonic Speeds by the Razor-Blade 

Technique,” A.R.C.R. & M, 1968. 

13. Smith K. G, Gaudet L., and Winter K. G., “The Use of Surface Pitot Tubes as Skin-Friction 

Meters at Supersonic Speeds,” A.R.C.R. & M, 1964. 

14. Schairer, E. T., Heineck, J. T., Walker, S. M., and Yaste, D. M., “Predicting Camera Views 

for Image-Based Measurements in Wind Tunnels,” AIAA Paper 2005-1349, 43rd AIAA 

Aerospace Sciences Meeting and Exhibit, Reno, NV, Jan. 10-13, 2005. 

15. Kushner, L. K., “Planning Image-Based Measurements in Wind Tunnels,” AIAA Paper 

2011-930, 49th AIAA Aerospace Sciences Meeting and Exhibit, Orlando, FL, Jan. 4-7, 

2011. 

16. Heineck, J. T., Schairer, E. T., and Walker, S. M., “PIV in NASA Ames Unitary Wind 

Tunnels,” AIAA 49th Aerospace Sciences Meeting, Orlando, FL, January 4-7, 2011. 

17. Shih, W. C. L., Wang, C., Coles, D., and Roshko, A., “Experiments on Flow Past Rough 

Circular Cylinders at Large Reynolds Numbers,” Journal of Wind Engineering and 

Industrial Aerodynamics, vol. 49, pp. 351-368, 1993. 
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17.0 Appendices 

Appendix A. Static Pressure Tap Locations 

Appendix B. Unsteady Pressure Tap Locations 

Appendix C. IR Thermography Image Archive Description 

Appendix D. PSP Data Archive Description 

Appendix E. PIV Data Archive Description 

Appendix F. Unsteady Pressured Data Archive Description 

Appendix G. Boundary Layer Data Archive Description 

Appendix H. Test Condition and Static Pressure Tap Data Archive Description 

 
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Appendix A. Static Pressure Tap Locations 

TAP ID COMPONENT NOMINAL 

ANGLE

X, inch Y, inch Z, inch RADIUS, 

inch 

1 HEATSHIELD 0 0.000 0.000 0.000 0.000 

2 HEATSHIELD 30 0.109 1.022 1.770 2.043 

3 HEATSHIELD 30 0.587 2.356 4.081 4.712 

4 HEATSHIELD 30 1.313 3.489 6.043 6.978 

5 HEATSHIELD 30 1.492 3.710 6.425 7.419 

6 HEATSHIELD 30 1.657 3.859 6.683 7.717 

7 HEATSHIELD 30 1.928 3.962 6.862 7.924 

8 HEATSHIELD 30 2.259 4.000 6.927 7.999 

9 HEATSHIELD 30 2.593 3.965 6.868 7.930 

10 HEATSHIELD 30 2.891 3.881 6.722 7.762 

11 HEATSHIELD 210 0.109 -1.022 -1.770 2.043 

12 HEATSHIELD 210 0.587 -2.356 -4.081 4.712 

13 HEATSHIELD 210 1.313 -3.489 -6.043 6.978 

14 HEATSHIELD 210 1.492 -3.710 -6.425 7.419 

15 HEATSHIELD 210 1.657 -3.859 -6.683 7.717 

16 HEATSHIELD 210 1.928 -3.962 -6.862 7.924 

17 HEATSHIELD 210 2.259 -4.000 -6.927 7.999 

18 HEATSHIELD 210 2.593 -3.965 -6.868 7.930 

19 HEATSHIELD 210 2.891 -3.881 -6.722 7.762 

20 HEATSHIELD 150 0.109 1.022 -1.770 2.043 

21 HEATSHIELD 150 0.587 2.356 -4.081 4.712 

22 HEATSHIELD 150 1.313 3.489 -6.043 6.978 

23 HEATSHIELD 150 1.492 3.710 -6.425 7.419 

24 HEATSHIELD 150 1.657 3.859 -6.683 7.717 

25 HEATSHIELD 150 1.928 3.962 -6.862 7.924 

26 HEATSHIELD 150 2.259 4.000 -6.927 7.999 

27 HEATSHIELD 150 2.593 3.965 -6.868 7.930 

28 HEATSHIELD 150 2.891 3.881 -6.722 7.762 

29 HEATSHIELD 330 0.109 -1.022 1.770 2.043 

30 HEATSHIELD 330 0.587 -2.356 4.081 4.712 

31 HEATSHIELD 270 0.109 -2.043 0.000 2.043 

32 HEATSHIELD 90 0.109 2.043 0.000 2.043 

33 BACK SHELL 9.500 -3.947 0.000 3.947 

34 STRUT FAIRING LHS 8.272 -4.655 0.089 4.656 
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35 STRUT FAIRING LHS 7.954 -4.738 -0.989 4.840 

36 STRUT FAIRING LHS 7.083 -5.093 -1.613 5.342 

37 STRUT FAIRING LHS 6.108 -5.679 -1.618 5.905 

38 STRUT FAIRING LHS 5.298 -6.279 -1.090 6.373 

39 STRUT FAIRING LHS 4.901 -6.601 -0.089 6.602 

40 STRUT FAIRING LHS 5.200 -6.357 0.962 6.429 

41 STRUT FAIRING LHS 5.967 -5.777 1.570 5.987 

42 STRUT FAIRING LHS 6.929 -5.174 1.652 5.431 

43 STRUT FAIRING LHS 7.844 -4.772 1.125 4.903 

44 BACK SHELL 330 4.901 -3.300 5.718 6.602 

45 BACK SHELL 210 4.901 -3.300 -5.718 6.602 

46 BACK SHELL 9.500 3.947 0.000 3.947 

47 STRUT FAIRING RHS 8.272 4.655 0.089 4.656 

48 STRUT FAIRING RHS 7.954 4.738 -0.989 4.840 

49 STRUT FAIRING RHS 7.083 5.093 -1.613 5.342 

50 STRUT FAIRING RHS 6.108 5.679 -1.618 5.905 

51 STRUT FAIRING RHS 5.298 6.279 -1.090 6.373 

52 STRUT FAIRING RHS 4.901 6.601 -0.089 6.602 

53 STRUT FAIRING RHS 5.200 6.357 0.962 6.429 

54 STRUT FAIRING RHS 5.967 5.777 1.570 5.987 

55 STRUT FAIRING RHS 6.929 5.174 1.652 5.431 

56 STRUT FAIRING RHS 7.844 4.772 1.125 4.903 

57 BACK SHELL 150 4.901 3.300 -5.718 6.602 

58 BACK SHELL 30 4.901 3.300 5.718 6.602 

59 BACK SHELL 0 4.901 0.000 6.602 6.602 

60 BACK SHELL 180 4.901 0.000 -6.602 6.602 

61 BACK FACE 0 10.504 0.000 2.500 2.500 

62 BACK FACE 180 10.504 0.000 -2.500 2.500 

63 PROBE 0 0.205 0.000 2.800 2.800 

64 TRAVERSE 0 0.207 0.250 2.800 2.811 

65 TRAVERSE 0 0.212 0.500 2.800 2.844 

 
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Appendix B. Unsteady Pressure Tap Locations 

TAP 

ID 

NOMINAL 

ANGLE X, IN Y, IN Z, IN RADIUS, IN 

ACTUAL 

ANGLE 

1 0 1.490 0.775 7.374 7.415 6.00 

2 0 1.584 0.686 7.584 7.615 5.17 

3 0 1.712 0.564 7.755 7.776 4.16 

4 0 1.866 0.418 7.880 7.892 3.03 

5 0 2.035 0.257 7.961 7.965 1.85 

6 0 2.212 0.089 7.997 7.997 0.64 

7 0 2.391 -0.081 7.989 7.989 -0.58 

8 0 2.566 -0.247 7.937 7.941 -1.78 

9 0 2.732 -0.404 7.844 7.854 -2.95 

10 0 2.895 -0.558 7.741 7.761 -4.12 

11 180 1.490 -0.775 -7.374 7.415 186.00 

12 180 1.584 -0.686 -7.584 7.615 185.17 

13 180 1.712 -0.564 -7.755 7.776 184.16 

14 180 1.866 -0.418 -7.881 7.892 183.03 

15 180 2.035 -0.257 -7.961 7.965 181.85 

16 180 2.212 -0.089 -7.997 7.997 180.64 

17 180 2.391 -0.081 -7.989 7.990 180.58 

18 180 2.566 0.247 -7.937 7.941 178.22 

19 180 2.732 0.404 -7.844 7.854 177.05 

20 180 2.895 0.558 -7.740 7.761 175.87 

21 270 1.490 -7.374 0.775 7.415 276.00 

22 270 1.584 -7.584 0.686 7.615 275.17 

23 270 1.712 -7.755 0.564 7.776 274.16 

24 270 1.866 -7.881 0.418 7.892 273.03 

25 270 2.035 -7.961 0.257 7.965 271.85 

26 270 2.212 -7.997 0.089 7.997 270.64 

27 270 2.391 -7.989 -0.081 7.989 269.42 

28 270 2.566 -7.937 -0.247 7.941 268.22 

29 270 2.732 -7.844 -0.404 7.854 267.05 

30 270 2.895 -7.740 -0.558 7.761 265.87 

31 330 1.490 -3.016 6.774 7.415 336.00 

32 330 1.584 -3.198 6.911 7.615 335.17 

33 330 1.712 -3.389 6.998 7.776 334.16 

34 330 1.866 -3.578 7.034 7.892 333.03 
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TAP 

ID 

NOMINAL 

ANGLE X, IN Y, IN Z, IN RADIUS, IN 

ACTUAL 

ANGLE 

35 330 2.035 -3.758 7.023 7.965 331.85 

36 330 2.212 -3.921 6.970 7.997 330.64 

37 330 2.391 -4.064 6.878 7.989 329.42 

38 330 2.566 -4.182 6.751 7.941 328.22 

39 330 2.732 -4.272 6.591 7.854 327.05 

40 330 2.895 -4.354 6.424 7.761 325.87 

41 0 4.061 0.000 7.087 7.087 0.00 

42 0 10.061 0.000 3.623 3.623 0.00 

43 180 4.061 0.000 -7.087 7.087 180.00 

44 180 10.061 0.000 -3.623 3.623 180.00 

 
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Appendix C. Test Condition and Static Pressure Tap Data Archive 

Description 

Whenever data was acquired during the wind-tunnel test, the Standard Data System (SDS) 

acquired its own set of data to document the test section conditions and the time-averaged static 

pressures measured on the model. In addition, the wall static pressure measurements made on 

the test-section walls are in the files in the form of pressure coefficients. The first line of this 

space-delimited file defines the data in each column. The files are located in the folder titled 

SDS. The files are named by run number (R0066.dat, for example). The two files named 

WallPortCoordinate11ftTWT contain the locations of the wall static pressure taps in wind-tunnel 

coordinates (in feet). 

 
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Appendix D. IR Thermography Image Archive Description 

All of the acquired IR images are located in the folders named 120CA_ceiling_IR, 

120CA_floor_IR, and 120CA_Nsidewall. The folder names refer to the camera locations. The 

files in these folders are in a proprietary FLIRTM format that is read by the ExaminIRTM software. 

Non-proprietary images taken from the FLIRTM software are presented in another set of folders 

named 120CA_ceiling_images, 120CA_floor_images, and 120CA_Nsidewall_images. At least 

one image (in tiff format) from each run is presented in this folder. In some cases multiple 

images for particular runs are presented, usually processed over different temperature ranges to 

bring out different flow features. 
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Appendix E. PSP Data Archive Description 

The PSP data is stored by run and sequence number according to the model geometry. The files 

are stored in the folders named PSP_AoA_15_Rough, PSP_AoA_30_Rough, and 

PSP_AoA_30_Smooth. Inside each of these folders is the grid file onto which the pressures are 

mapped for that particular model configuration (configx.grid) and a folder (named “grid”)

containing the corresponding pressure data as pressure coefficients. The grids used are multigrids that are stored in Plot3D format. The corresponding pressure coefficient data is stored in 

Plot3D function format. Multiple CFD plotting packages are available that can read this data 

format. 

Within the “grid” folders are the individual Cp data files. They are named using the run and 

sequence numbers (for example, “005503.p3d” which contains data from Run 55 sequence 03).
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Appendix F. PIV Data Archive Description 

PIV measures the velocity on a grid of points in the measurement area. In this experiment, two 

stereoscopic PIV systems were used to acquire three components of velocity measured in the 

PIV coordinate system. In this coordinate system the x-axis is the same as the wind tunnel 

x-axis. The y-axis for the PIV measurements is the other axis in the plane of the laser sheet. For 

the high-speed system, the y-axis is the same as the wind-tunnel y-axis. The PIV y-axis for the 

wide-view system is equivalent to the wind tunnel z-axis. The PIV data is presented in metric 

units with the x-, y-, and z-dimensions in millimeters (in the PIV coordinate systems) and the 

velocity data in meters/second. 

The PIV data are located in several compressed files (currently located on aeroCOMPASS) listed 

in the lower left of Table F-1. The wide-view files contain the reduced data from the wide-view 

PIV system that captured the vertical centerline plane of the wake. The high-speed files contain 

the data acquired with the high-speed PIV system in a small area of a horizontal plane in the 

wake. The folders inside these compressed files are structured similarly and are described 

below. Table F-1 shows the contents of the all the data folders. 
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Table F-1. Contents of High-speed and Wide-view PIV Data Folders

 

Using the high-speed data as an example, the folder names contain information about the 

particular run during which the data was acquired. For example, the “M.3_3.5D_r150_s4” folder 

contains all 2000 of the measured vector fields, in Tecplot format, for run 150, sequence 4 with a 

free-stream Mach number of 0.3 and the model in its upstream location. For cases with the 

model in the downstream location the designation is “near” instead of “3.5D”. Each file in this 

folder contains the measurement grid and the velocity vector at each grid point. Detailed 

information about the test conditions for all of the runs is located in the SDS Data directory on 

aeroCOMPASS. 

There are two other folders that contain information obtained from the 2000 velocity 

measurements. The folder “M.3_3.5D_r150_s4_scalar” contains statistical information about the 
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velocity field for run 150, also in Tecplot format. Each of the 15 files contains the measurement 

grid with a single scalar value associated at each grid point. The values contained in each of the 

files are shown in Table F-2. 

The third folder for each of the PIV runs contains the averaged velocity vector. For the case 

described above the folder name is M.3_3.5D_r150_s4_vector. This particular file in the folder 

has the average velocity vector at each of the grid points of the same measurement grid as in the 

other folders/files for this data point. 

All of the high-speed PIV data was acquired at a rate of 2,000 Hz. Each image was processed on 

32x32 pixel interrogation areas with 75 percent overlap between adjacent interrogation areas. 

The entire image area did not have adequate seeding and/or image quality so a smaller area was 

chosen for the final processing to ensure that the data is of high quality at every grid point. With 

that processing and image area, the measurement grid provides a velocity measurement 

approximately every 1.4 millimeters in the x- and y-directions. For the measurements made with 

the model in the upstream location the measurement area extends approximately 240mm stream 

wise and 167mm span wise. When the model was located in its downstream position the 

measurement area was somewhat smaller, extending approximately 152 millimeters stream wise 

and 170 millimeters span wise. The measurement areas of the high-speed PIV system are shown 

in Figure F-1. The x and y locations are given in millimeters. 
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Table F-2. Scalar Data File Contents

 
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Figure F-1. Measurement Area for High-speed PIV System; Measurements in Millimeters

The wide-view files are arranged in a similar manner, but there are two additional parameters in 

the file name because results processed on two different interrogation areas are given and 

additional samples are available for some conditions. A basic folder name for the wide-view 

data is “M0.3_far_r150_2000samps_64_IA_vector.” “M0.3” designates Mach 0.3, “far” 

designates model in the upstream location (near is for the downstream model location), 

(4148, -319)

(4150, -152) (4384, -150)

(4394, -315)

Measurement area with 

model in upstream location

(4180, -150) (4329, -147)

(4179, -319) (4334, -318)

Measurement area with model 

in downstream location
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“2000samps” says there are 2000 samples for this particular data set (“4000samps” or “_s1 and 

s2_4000samps” would indicate 4000 samples), and “64_IA” indicates a 64x64 pixel 

interrogation area (alternate is “128_IA” for a 128x128 interrogation area).

The imaging area was much larger for the wide-view PIV system, but once again not all of the 

area provided images that could be successfully processed. To summarize some of the issues 

limiting the useful image area, the laser sheet moved more than desired (probably about 1 sheet 

thickness or about 5mm), the seed was not uniformly dispersed in the image area, and there were 

several reflections of stray laser illumination on the test section walls that were visible in the 

camera images. These imperfections resulted in limiting the size of the area of the flow that 

could be accurately measured and increased the size of the interrogation area. This effectively 

reduced the measurement density from that seen from the high-speed system. 

A 64x64 interrogation area was the smallest that gave accurate (interrogation-area-independent) 

average velocity measurements. With 75 percent overlap between interrogation areas, the 

spacing between measurement points was approximately 5.3mm in x- and y-directions. The cut 

out in the data at the lower left is a mask applied to the processing surrounding the CM model to 

prevent glare from the laser reflecting off the model from affecting the PIV measurements. The 

spacing between measurement points increases to 10.6mm for the 128x128 pixel interrogation 

area processing. The coarser grid processing gives more reliable results for the statistical 

information about the flowfield while the time-averaged velocities given in the 64x64 processing 

are very accurate and the resolution is better than in the coarser processing grid. In all cases, the 

interrogation area for the wide-view measurements has corner points roughly as shown in 

Figure F-.2 (wind tunnel coordinates in mm). 

 
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Figure F-2. Measurement Area for Wide-view PIV System; Measurements in Millimeters

 

(3948, 278)

(3948, 27)

(4158, -172)

(4164, -225)

(4264, -276) (5067, -305)

(5059, 416)
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Appendix G. Unsteady Pressured Data Archive Description 

The unsteady pressure measurements are located on aeroCOMPASS in the final data folder of 

the 120-CA test location. The unsteady pressure data folder contains a file for each run that 

includes a summary of the test conditions and the time history of the 44 Kulite® pressure 

transducers. All data was acquired at a rate of 6400 Hz. The files are comma-separated variable 

type with names structured as “t11-0247Txxpytime.cvs” where t11-0247T represents the facility 

test number, xx is the run number, and y is the sequence or point number. When there were 

multiple data points acquired for a given run, there will be a file for each data point. The data is 

presented as CP’ at each time step where CP’ is the unsteady component of the pressure 

coefficient. It is defined as where is the instantaneous static 

pressure at time t and is the time-averaged static pressure. 

 

Ps  t

Ps
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Appendix H. Boundary Layer Data Archive Description 

The boundary layer data is archived in a folder named “BL data files,” currently on the 

aeroCOMPASS site. Each of the files in this folder contains the boundary layer survey acquired 

during the run in the file name (BLrun122.txt, for example). The files are tab-delimited text with 

the first line defining what values are in each of the columns. 
REPORT DOCUMENTATION PAGE Form Approved

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Technical Memorandum

 4. TITLE AND SUBTITLE

Orion Multi-Purpose Crew Vehicle (MPCV) Capsule Parachute Assembly 

System (CPAS) Wake Deficit Wind Tunnel Testing

5a. CONTRACT NUMBER

 6. AUTHOR(S)

Ross, James C.; Schuster, David M.

 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

NASA Langley Research Center

Hampton, VA 23681-2199

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L-20384 NESC-RP-11-00697

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NASA

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Subject Category 16-Space Transportation and Safety

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19a. NAME OF RESPONSIBLE PERSON

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14. ABSTRACT

During descent after re-entry into the Earth's atmosphere, the Orion CM deploys its drogue parachutes at approximately Mach 

0.7. Accurately predicting the dynamic pressure experienced by the drogue parachutes at deployment is critical to properly 

designing the parachutes. This NASA Engineering and Safety Center assessment was designed to provide a complete set of 

flowfield measurements on and around an idealized Orion Crew Module shape with the most appropriate wind tunnel 

simulation of the Orion flight conditions prior to parachute deployment. This document contains the details of testing and the 

outcome of the assessment.

15. SUBJECT TERMS

Capsule Parachute Assembly System; Wind Tunnel Testing; Multi-Purpose Crew Vehicle; NASA Engineering and Safety 

Center; Wake Deficit

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