Raymond Castner 

Glenn Research Center, Cleveland, Ohio

Stephen Shaw and Eric Adamson

Boeing Commercial Aircraft, Seattle, Washington

Stephanie Simerly

Glenn Research Center, Cleveland, Ohio

Background Pressure Profi les for Sonic Boom

Vehicle Testing in the NASA Glenn 8- by 6-Foot

Supersonic Wind Tunnel

NASA/TM—2013-217903

July 2013
NASA STI Program . . . in Profi le

Since its founding, NASA has been dedicated to the 

advancement of aeronautics and space science. The 

NASA Scientifi c and Technical Information (STI) 

program plays a key part in helping NASA maintain 

this important role.

The NASA STI Program operates under the auspices 

of the Agency Chief Information Offi cer. It collects, 

organizes, provides for archiving, and disseminates 

NASA’s STI. The NASA STI program provides access 

to the NASA Aeronautics and Space Database and 

its public interface, the NASA Technical Reports 

Server, thus providing one of the largest collections 

of aeronautical and space science STI in the world. 

Results are published in both non-NASA channels 

and by NASA in the NASA STI Report Series, which 

includes the following report types:

• TECHNICAL PUBLICATION. Reports of 

completed research or a major signifi cant phase 

of research that present the results of NASA 

programs and include extensive data or theoretical 

analysis. Includes compilations of signifi cant 

scientifi c and technical data and information 

deemed to be of continuing reference value. 

NASA counterpart of peer-reviewed formal 

professional papers but has less stringent 

limitations on manuscript length and extent of 

graphic presentations.

• TECHNICAL MEMORANDUM. Scientifi c 

and technical fi ndings that are preliminary or 

of specialized interest, e.g., quick release 

reports, working papers, and bibliographies that 

contain minimal annotation. Does not contain 

extensive analysis.

• CONTRACTOR REPORT. Scientifi c and 

technical fi ndings by NASA-sponsored 

contractors and grantees.

• CONFERENCE PUBLICATION. Collected 

papers from scientifi c and technical 

conferences, symposia, seminars, or other 

meetings sponsored or cosponsored by NASA.

• SPECIAL PUBLICATION. Scientifi c, 

technical, or historical information from 

NASA programs, projects, and missions, often 

concerned with subjects having substantial 

public interest.

• TECHNICAL TRANSLATION. Englishlanguage translations of foreign scientifi c and 

technical material pertinent to NASA’s mission.

Specialized services also include creating custom 

thesauri, building customized databases, organizing 

and publishing research results.

For more information about the NASA STI 

program, see the following:

• Access the NASA STI program home page at 

http://www.sti.nasa.gov

• E-mail your question to help@sti.nasa.gov

• Fax your question to the NASA STI 

Information Desk at 443–757–5803

• Phone the NASA STI Information Desk at

 443–757–5802

• Write to:

 STI Information Desk

 NASA Center for AeroSpace Information

 7115 Standard Drive

 Hanover, MD 21076–1320
Raymond Castner 

Glenn Research Center, Cleveland, Ohio

Stephen Shaw and Eric Adamson

Boeing Commercial Aircraft, Seattle, Washington

Stephanie Simerly

Glenn Research Center, Cleveland, Ohio

Background Pressure Profi les for Sonic Boom

Vehicle Testing in the NASA Glenn 8- by 6-Foot

Supersonic Wind Tunnel

NASA/TM—2013-217903

July 2013

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

43rd Fluid Dynamics Conference and Exhibit

sponsored by the American Institute of Aeronautics and Astronautics

San Diego, California, June 24–27, 2013
Acknowledgments

This work was funded by the NASA Fundamental Aeronautics Program, the High Speed Project.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

Available electronically at http://www.sti.nasa.gov

Trade names and trademarks are used in this report for identifi cation 

only. Their usage does not constitute an offi cial endorsement, 

either expressed or implied, by the National Aeronautics and 

Space Administration.

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

This work was sponsored by the Fundamental Aeronautics Program 

at the NASA Glenn Research Center.
NASA/TM—2013-217903 1

Background Pressure Profiles for Sonic Boom Vehicle Testing in the 

NASA Glenn 8- by 6-Foot Supersonic Wind Tunnel

Raymond Castner

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Stephen Shaw and Eric Adamson

Boeing Commercial Aircraft

Seattle, Washington 98124

Stephanie Simerly

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Abstract

In an effort to identify test facilities that offer sonic boom measurement capabilities, an exploratory 

test program was initiated using wind tunnels at NASA research centers. The subject of this report is the 

sonic boom pressure rail data collected in the Glenn Research Center 8- by 6-Foot Supersonic Wind 

Tunnel. The purpose is to summarize the lessons learned based on the test activity, specifically relating to 

collecting sonic boom data which has a large amount of spatial pressure variation. The wind tunnel 

background pressure profiles are presented as well as data which demonstrated how both wind tunnel 

Mach number and model support-strut position affected the wind tunnel background pressure profile. 

Techniques were developed to mitigate these effects and are presented.

Nomenclature

h Distance below vehicle, in.

l Vehicle length, in.

M∞ Free-stream Mach number

P Measured static pressure, psia

P∞ Test section static pressure, psia

ΔP P - P∞.

ΔP/P (P - P∞)/ P∞.

T∞ Test section total temperature, R.

t Time, sec.

xts tunnel station, in.

xa Travel distance for axial linear actuator, in.

I. Introduction

Supersonic travel is not allowed over populated areas due to the disturbance caused by the sonic

boom. At the vehicle nose there is a rise in pressure, followed by a steady decrease in pressure, followed 

by a rise to atmospheric pressure. When propagated to the ground, this pressure profile takes the shape of 

an N-wave. The two large pressure changes create a “double boom” effect. The impact of the sonic boom 
NASA/TM—2013-217903 2

is so large that the Federal Aviation Administration has issued a noise policy for supersonic aircraft 

stating: “Since March 1973, supersonic flight over land by civil aircraft has been prohibited by regulation 

in the United States. The Concorde (Ref. 1) was the only civil supersonic airplane that offered service to 

the United States, and it is no longer in service” (Ref. 2). The same policy also states that “noise operating 

rules would propose that any future supersonic airplane produce no greater noise impact on a community 

than a subsonic airplane.” Subsonic noise limits are prescribed in Code of Federal Regulations 14-Part 36

stage 4 (Ref. 3).

An experimental validation program was sponsored by NASA in which Boeing designed a supersonic 

aircraft configuration. The configuration was to generate a low sonic boom signature intended to enable

flight over populated areas (Ref. 4). To assess facilities that offered sonic boom measurement capability, 

an exploratory test program was initiated using wind tunnels at NASA research centers.

The first exploratory test was conducted in April 2012 at the NASA Ames Research Center (ARC)

9- by 7-Foot Supersonic Wind Tunnel, where Mach numbers range from 1.55 to 2.55. The second test 

was conducted in June 2012 at the ARC 11-Foot Transonic Wind Tunnel, where Mach numbers range 

from 0.2 to 1.45. The third exploratory test was performed in September 2012 at the Glenn Research 

Center (GRC) 8- by 6-Foot Supersonic Wind Tunnel (Ref. 5) (8×6 SWT), where Mach numbers range 

from 0.36 to 2.0.

The subject of this report is the sonic boom pressure measurement rail (referred to as the pressure rail)

data collected in the 8×6 SWT. The purpose is to summarize the lessons learned based on the test activity, 

specifically relating to collecting sonic boom pressure profiles in the 8×6 SWT which has a large amount 

of spatial pressure variation. The wind tunnel background pressure profiles are presented. Results

demonstrate how both wind tunnel Mach number and model support strut position affect the measured

pressure profiles. Techniques were developed to mitigate these effects and are presented.

II. Experimental Setup

A. Test Setup

The 8×6 SWT transonic test section was configured with the 14-ft test section, setup for Schlieren 

imaging, and the 5.8 percent porosity setting, shown in Figure 1. Porous blanks, where the bleed holes 

were filled with plugs, were installed in the upstream windows. The Schlieren windows were rotated to 

the lower 45° position. Halfway through the test there were concerns regarding the porous blanks

influencing data quality on the pressure rail. The Schlieren windows were re-located to the upstream 

position and the porous blanks downstream. Test section configuration after this switch was the 14-ft test 

section and 5.8 percent modified porosity setting. The Schlieren windows were rotated so that they were 

positioned upstream and on the tunnel centerline. The window position did not affect the results.

Three models were tested at supersonic speeds ranging from Mach 1.2 to Mach 1.8. A pressure rail was 

installed in ceiling plates along the centerline of the test section to capture model pressure signature data. 

The model was installed on the sting in an upside down orientation, so the pressure signature, or boom 

signature, from the bottom of the model was measured by the pressure rail. A force balance was installed 

between the model and the strut, and was used to measure both forces and moments on the model.

During the test, each model’s spatial position was varied horizontally (x-axis, in line with the sting) 

up to 24 in. by a linear actuator and vertically (z-axis) up to 64 in. by raising and lowering the supersonic 

strut. Pressure measurements from the translations were averaged to generate the vehicle pressure 

profiles. The supersonic strut position was controlled remotely, and was raised or lowered while the 

tunnel was operating.
NASA/TM—2013-217903 3

B. Models

The Boom model was a 0.65 percent scale replica of the Boeing N+2 Quiet Experimental Vehicle 

Concept (Ref. 4) (QEVC) and is displayed in Figure 2. It was 14.994 in. long, 0.858 in. thick and had a 

6.732 in. wingspan. This low-boom wing/body configuration was manufactured in five pieces: nose 

assembly, wing/body, upper-swept strut and attachment, cone-section, and balance shield. All leading and 

trailing edges on the boom model were 0.004 in. thick.

The Performance model was a 1.79 percent scale replica of the Boeing N+2 QEVC. It was 41.158 in. 

long, 2.343 in. thick and had an 18.524 in. wingspan. It had a removable nose assembly, outboard wing, 

nacelle, aft deck, vertical tail, horizontal airfoil, and canard. All leading and trailing edges on the 

performance model were also 0.004 in. thick.

The axisymmetric model, number 2, (AS2) was used to obtain calibration data (Fig. 3). The AS2

calibration model was the same scale and had approximately the same initial overpressure as the

1.79 percent scale research model. Data were obtained from the AS2 calibration model at the same test 

conditions indicated for the research model testing. Angle of attack for the AS2 calibration model was set 

to zero.

C. Instrumentation

The 14-in. pressure rail was 90 in. long, 14 in. tall, and 1 in. thick at its base. The rail tapered down to 

0.050 in. thickness at the tip. Instrumentation included 420 pressure taps distributed 0.1575 in. apart over a 

66 in. length, Figure 4. The leading edge of the pressure rail was positioned at Tunnel Station 158.2.

D. Spatial Averaging Measurement Method

Spatial averaging (Ref. 6) was conducted by averaging data at several positions in the fore and aft 

direction for an “X-sweep”, or in the vertical direction for a “Z-sweep”. The model was moved in 

incremental distances of approximately 4 rail pressure ports, for a total of 13 model positions (Ref. 4).

Test conditions were nominally between Mach 1.2 and Mach 1.8, where Mach number was initially 

held to ±0.01, and later improved to ±0.001. In the 8×6 SWT, Mach number is controlled by three 

variables: (1) flex wall nozzle throat position, (2) compressor speed, and (3) balance chamber exhaust 

valve position. The exhaust valve in the balance chamber controls the flow of air exiting the porous walls, 

which has an effect on the static pressure in the test section.

Measurements of vehicle pressure profile (ΔP/P) were collected as defined by the following equation:

ΔP/P = ((Prail- P∞) / P∞ )model - ((Prail- P∞) / P∞ )reference run 

Reference runs were originally performed at h = 68 in., and later changed to an in-line reference at

h = 56 in. (Fig. 5).

III. Experimental Results

The initial sonic boom pressure profiles collected in the 8×6 SWT did not match predictions, as 

demonstrated with data collected from the AS2 calibration model. The AS2 model pressure profile should 

be characterized by a nose shock followed by a “flat-top” region. The signature collected displayed the 

nose shock, but the “flat-top” feature was missing. Additionally, profile values were lower than zero 

leading up to the nose shock, which was an undesireable characteristic. The profile is shown in Figure 6, 

where the 13 individual measurements are plotted, along with the spatial average. Profiles are shown for 

Mach 1.8, but were consistent for all Mach numbers examined in this experiment.
NASA/TM—2013-217903 4

Resolution of this problem was accomplished through an examination of four factors: (1) the 

reference run, (2) variation of rail pressures with Mach number, (3) variation of the reference run profile 

with Mach number, and (4) the effect of strut position on the reference run. After each factor was 

understood and controlled, a significant improvement to the AS2 model sonic boom pressure profile was 

demonstrated.

A. Reference Run

The uncorrected pressure profile (ΔP/P) was measured by the pressure rail. To represent the pressure 

variation in an empty wind tunnel, the wind tunnel model was moved to a position 12-in. off the floor

(h = 68 in.). In this position the Mach waves from the model did not intersect the pressure rail. The tunnel 

background pressure profile was characterized by a number of shock and expansion waves, starting at

xts = 175 to xts = 210 in. along the pressure rail. The pressure profile between the peaks exhibited steep 

pressure gradients. The region from xts = 210 to xts = 278 exhibited shock and expansion waves of smaller 

magnitude, as shown in Figure 7. At values greater than xts = 280, the profile returned to a series of strong 

shock and expansion waves, some caused by the model support system.

During testing, this reference pressure pattern was subtracted from the measurement to determine 

vehicle pressure profile, or sonic boom signature. As this reference profile shifted in position or 

magnitude, it had a dramatic effect on the vehicle measurement. The effect depended on location. In an 

area of steep pressure gradient, small changes in the reference run had a large effect on the measurement.

Conversely, the effect was smaller in areas of low pressure gradients, as will be shown in Section III.C.

B. Variation of Individual Rail Pressures

Individual measurement ports on the pressure rail were examined for the data that was collected from 

September 4 to 10, 2012. During this time, the tunnel Mach number was controlled to ±0.01. An example 

is provided where the value of ΔP/P at the first pressure rail measurement port was examined. The first 

port exhibited a linear correlation with Mach number, indicating that variation in Mach number caused 

the value of ΔP/P to change by 0.01, see Figure 8. This 0.01 variation for the individual pressure rail port

was greater than the peak-to-peak value of the AS2 signature to be measured.

The wind tunnel Mach number exhibited variations from Mach 1.786 to Mach 1.792, a difference of

0.006. If variation in wind tunnel Mach number could be controlled to ±0.001, then the variation in ΔP/P 

would also be reduced.

C. Variation of the Reference Run Pressures

The effect of Mach number on all of the pressure rail pressures was examined. For this data, the wind 

tunnel Mach number was controlled to ±0.001, and was held at constant values of Mach 1.776, 1.777, 

1.778, and 1.779. Pressure rail pressures were normalized to the Mach 1.776 condition. For some 

locations on the pressure rail, the ΔP/P pressure variation was 0.007, as seen in Figure 9. This data

demonstrates how regions of high pressure gradient affected the reference run as Mach number was 

increased. The variations in background pressure profile had the potential to cause significant offset in the 

measured sonic boom pressure profile.

D. Effect of Strut Position

The supersonic strut in the 8×6 SWT was not symmetric, as there was only blockage below the 

support sting and not above. During tunnel operation, the remote strut movement allowed reference runs 

at two model heights, h = 68 in. (near the floor) and h = 56 in. Variations in the reference values of ΔP/P, 

some greater than 0.018, were noted when the reference run height was changed. It was found that as the 

model was moved to different heights, the resulting change in test section blockage produced variations in 
NASA/TM—2013-217903 5

Mach number. To collect adequate pressure profiles, a reference run was performed at h = 56 in., and then 

the sonic boom data were collected at the same height. This was achieved by moving the model to the aft 

portion of the pressure rail (x = 0, linear actuator fully retracted) for a reference run, and then to the 

forward portion of the pressure rail for the measurement. Data sets from x = 12 to 24 in. were averaged to 

obtain the final vehicle pressure profile. Figure 10 shows the difference in the reference run pressure 

profile due to model height.

Near the end of the test, the Mach variations due to strut movement were controlled to within tolerance 

by the tunnel control system, but took a large amount of test time. Minimizing strut movement reduced the 

amount of time needed for the tunnel to maintain the desired conditions during data collection. 

E. AS2 Model Signatures

The AS2 calibration model signature was collected and the reference run was conducted at the same 

Mach number as the measurement, with a tolerance of ±0.001. The reference run was conducted at

h = 56 in., which was the same height used in the measurement. In this case, the AS2 model pressure 

profile was captured properly; with a value of zero leading up to the nose shock, and a “flat-top” region 

after the nose shock (Fig. 11). These results for the AS2 model were the same as those reported by Magee

(Ref. 3), who also reported the near field pressure profiles for the Boeing QEVC vehicle concept as 

collected in the 8×6 SWT. Even with these improvements, boom model profiles sometimes exhibited a 

small zero offset. To mitigate the zero offset, improvements to the process need to be investigated. These 

improvements include the effect of strut position on wind tunnel Mach number, and options for tighter 

control of Mach number.

IV. Conclusions

A test was conducted in the NASA Glenn Research Center 8- by 6-Foot Supersonic Wind Tunnel to 

assess the capability to make sonic boom measurements on a vehicle design with a low sonic boom 

signature. The test was conducted in September 2012. Three models were tested at supersonic speeds 

from Mach 1.2 to Mach 1.8, where pressure measurements were collected with a pressure rail mounted in 

the tunnel ceiling. This report summarizes the lessons learned to collect quality sonic boom pressure 

profiles; specifically relating to the background pressure profile that contains a large amount of spatial 

variation. Results were demonstrated for Mach 1.8 and were applicable to all Mach numbers.

Spatial variation of the background pressure profile in the wind tunnel was caused by a series of 

shock and expansion waves which were generated and reflected throughout the wind tunnel. These waves 

were likely created by the wind tunnel nozzle, throat, and/or imperfections in the tunnel walls. However, 

the origins of these shock and expansion waves were not well understood. These waves needed to be 

controlled when collecting sonic boom data on models with low sonic boom signatures. It was shown that 

the shock and expansion waves caused a background pressure profile with a large amount of spatial 

variation and steep pressure gradients.

As the tunnel Mach number changed, the shock and expansion waves generated and reflected through 

the tunnel changed position, related to the Mach angle. As these waves changed position, the background 

pressure profile moved upstream or downstream in very small increments. This movement was 

demonstrated by the correlation of the individual rail measurements with changes in Mach number. 

During the measurement of sonic boom signatures, the shock and expansion waves must not move. If the 

shocks are not stationary, background pressures that were subtracted from the measurement, move 

dramatically up and down steep gradients. The movement of the background pressure profile made for 

incorrect measured vehicle pressure profiles.
NASA/TM—2013-217903 6

For this test, the wind tunnel Mach number had to be maintained to ±0.001 and the model support 

strut had to remain in one position. Mach number was controlled by three variables: (1) flex wall nozzle 

throat position, (2) compressor speed, and (3) balance chamber exhaust valve position. The 8×6 SWT has 

several unique features that created variability of the Mach number in the test section. These features 

included the asymmetric strut, the flow through the porous walls, and the balance chamber pressure. When 

these key variables were controlled, successful measurement of sonic boom pressure profiles was achieved. 

Figure 1.—Wind tunnel layout for the sonic boom model in the 8×6 SWT.

Model Linear 

actuator

Supersonic 

strut

14-in. pressure rail

Pressure rail leading edge at tunnel 

 

Schlieren 

window
NASA/TM—2013-217903 7

Figure 2.—Boeing N+2 validation model and 14-in. sonic boom pressure rail 

installed in the 8×6 SWT.

Pressure

rail
NASA/TM—2013-217903 8

Figure 3.—AS2 calibration model.

Figure 4.—14-in. pressure rail.
NASA/TM—2013-217903 9

Figure 5.—Reference run positions.

Figure 6.—Mach 1.8, AS2 model signature at h = 56 in., reference run h = 68 in.

Floor reference run, h=68:

In-line reference run, h=56: Model Strut

Pressure rail

Non-zero leading Nose shock

signature

ΔP/P

xts, in
NASA/TM—2013-217903 10

Figure 7.—Tunnel reference run ΔP/P, Mach 1.8.

Figure 8.—Rail pressure variation with Mach number, nominal Mach 1.8.

-0.1000

-0.0800

-0.0600

-0.0400

-0.0200

0.0000

0.0200

0.0400

0.0600

0.0800

0.1000

170 180 190 200 210 220 230 240 250 ΔP/P

xts, in

Steep pressure gradients, 

Series of shocks and expansions

0.0300

0.0350

0.0400

0.0450

0.0500

0.0550

0.0600

1.785 1.786 1.787 1.788 1.789 1.79 1.791 1.792 1.793

ΔP/P, port number 1

Mach
NASA/TM—2013-217903 11

Figure 9.—Change in pressure profiles for reference run with Mach number, nominal Mach 1.8.

Figure 10.—Effect of strut position on reference run signature, h = 68 in. compared to h = 56 in. height.

-0.035

-0.025

-0.015

-0.005

0.005

0.015

175 185 195 205 215 225 235 245

ΔP/P

Xts, in

Mach 1.776

Mach 1.777

Mach 1.778

Mach 1.7792

Mach 1.7798

0.007 variation

Xts, in

0.018 variation

h = 68 in

h = 56 in

ΔP/P 0.02

0.04

0.06

0.08

0.1

0.12

-0.02

-0.04

-0.06

-0.08

-0.1
NASA/TM—2013-217903 12

Figure 11.—Mach 1.8, AS2 model signature at h = 56 in., reference run h = 56 in.

References

1. “Concorde SST: History” Concorde SST, http://www.concordesst.com/history/historyindex.html

[retrieved 24 Feb. 2013].

2. “Supersonic Aircraft Noise,” http://www.faa.gov/about/office_org/headquarters_offices/apl/

noise_emissions/supersonic_aircfaft_noise/ [retrieved 24 Feb 2013].

3. “Part 36—Noise Standards: Aircraft Type and Airworthiness Certification,” http://www.ecfr.gov/cgibin/text idx?c=ecfr&SID=f159fee0853e9ba1e2753a39f483f653&rgn=div5&view=text&node=

14:1.0.1.3.19&idno=14 [retrieved 10 May 2013].

4. Magee, T., Shaw, S., and Fugal, S., “Experimental Validations of a Low-Boom Aircraft Design,” 

AIAA-2013-646, Jan. 2013.

5. Soeder, R., “NASA Lewis 8- by 6-Foot Supersonic Wind Tunnel User Manual,” NASA TM-105771, 

Feb. 1993.

6. Morgenstern, J., “How to Accurately Measure Low Sonic Boom or Model Surface Pressures in 

Supersonic Wind Tunnels,” AIAA–2012–3215, Jan. 2013.

Nose shock

Flat top region

Zero leading signature

xts, in

ΔP/P

REPORT DOCUMENTATION PAGE Form Approved 

OMB No. 0704-0188 

The public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the 

data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this 

burden, to Department of Defense, Washington Headquarters Services, Directorate for Information Operations and Reports (0704-0188), 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302. 

Respondents should be aware that notwithstanding any other provision of law, no person shall be subject to any penalty for failing to comply with a collection of information if it does not display a currently valid OMB 

control number. 

PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS. 

1. REPORT DATE (DD-MM-YYYY)

01-07-2013 

2. REPORT TYPE

Technical Memorandum 

3. DATES COVERED (From - To)

4. TITLE AND SUBTITLE

Background Pressure Profiles for Sonic Boom Vehicle Testing in the NASA Glenn 8- by 6-

Foot Supersonic Wind Tunnel 

5a. CONTRACT NUMBER

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Castner, Raymond; Shaw, Stephen; Adamson, Eric; Simerly, Stephanie 

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

WBS 475122.02.03.03.02 

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field 

Cleveland, Ohio 44135-3191 

8. PERFORMING ORGANIZATION

 REPORT NUMBER

E-18715 

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

Washington, DC 20546-0001 

10. SPONSORING/MONITOR'S

 ACRONYM(S)

NASA 

11. SPONSORING/MONITORING

 REPORT NUMBER

NASA/TM-2013-217903 

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited 

Subject Categories: 01, 02, 05, and 09 

Available electronically at http://www.sti.nasa.gov 

This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

In an effort to identify test facilities that offer sonic boom measurement capabilities, an exploratory test program was initiated using wind 

tunnels at NASA research centers. The subject of this report is the sonic boom pressure rail data collected in the Glenn Research Center 8- 

by 6-Foot Supersonic Wind Tunnel. The purpose is to summarize the lessons learned based on the test activity, specifically relating to 

collecting sonic boom data which has a large amount of spatial pressure variation. The wind tunnel background pressure profiles are 

presented as well as data which demonstrated how both wind tunnel Mach number and model support-strut position affected the wind tunnel 

background pressure profile. Techniques were developed to mitigate these effects and are presented.

15. SUBJECT TERMS

Pressure distribution; Sonic booms; Supersonic wind tunnels 

16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF

 ABSTRACT

UU 

18. NUMBER

 OF 

 PAGES

20 

19a. NAME OF RESPONSIBLE PERSON

STI Help Desk (email:help@sti.nasa.gov) 

a. REPORT

U 

b. ABSTRACT

U 

c. THIS 

PAGE

U 

19b. TELEPHONE NUMBER (include area code)

443-757-5802 

Standard Form 298 (Rev. 8-98)

Prescribed by ANSI Std. Z39-18