Raymond Castner

Glenn Research Center, Cleveland, Ohio

Exhaust Nozzle Plume Effects on Sonic Boom

Test Results for Vectored Nozzles

NASA/TM—2012-217229

February 2012

AIAA–2011–5974
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Raymond Castner

Glenn Research Center, Cleveland, Ohio

Exhaust Nozzle Plume Effects on Sonic Boom

Test Results for Vectored Nozzles

NASA/TM—2012-217229

February 2012

AIAA–2011–5974

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

47th Joint Propulsion Conference and Exhibit

cosponsored by AIAA, ASME, SAE, and ASEE

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Acknowledgments

John Jones provided design and fabrication of the wind tunnel model and support systems. Facility support was provided by 

Scott Smrdel, Bob Clark and Bob McClusky, at the NASA Glenn Research Center. Test assistance was also provided

by Manan Vyas. Flight data was provided by Ed Haering at the NASA Dryden Research Center. This work was 

funded by the NASA Fundamental Aeronautics Program, Supersonics Project.

Available from

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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.

This work was sponsored by the Fundamental Aeronautics Program 

at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2012-217229 1 

Exhaust Nozzle Plume Effects on Sonic Boom Test Results for 

Vectored Nozzles 

Raymond Castner 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

Reducing or eliminating the operational restrictions of supersonic aircraft over populated areas has 

led to extensive research at NASA. Restrictions were due to the disturbance of the sonic boom, caused by 

the coalescence of shock waves formed off the aircraft. Recent work has been performed to reduce the 

magnitude of the sonic boom N-wave generated by airplane components with a focus on shock waves 

caused by the exhaust nozzle plume. Previous Computational Fluid Dynamics (CFD) analysis showed 

how the shock wave formed at the nozzle lip interacts with the nozzle boat-tail expansion wave. An 

experiment was conducted in the 1- by 1-foot Supersonic Wind Tunnel (SWT) at the NASA Glenn 

Research Center. Results show how the shock generated at the nozzle lip affects the near field pressure 

signature, and thereby the potential sonic boom contribution for a nozzle at vector angles from 3 to 8. 

The experiment was based on the NASA F-15 nozzle used in the Lift and Nozzle Change Effects on Tail 

Shock experiment, which possessed a large external boat-tail angle. In this case, the large boat-tail angle 

caused a dramatic expansion, which dominated the near field pressure signature. The impact of nozzle 

vector angle and nozzle pressure ratio are summarized. 

Nomenclature 

D test nozzle diameter, inches 

NPR nozzle pressure ratio = Pt/P∞

P local static pressure, psia 

Pt total pressure in nozzle, psia 

P∞ free stream static pressure, psia

ΔP/P (P - P∞)/P∞ (also Cp)

t time, seconds 

x axial distance from jet simulator nosecone tip, inches 

y distance from nozzle centerline, inches 

1.0 Introduction 

NASA has been conducting extensive research to reduce the sonic boom signature caused by 

supersonic flight speeds. The Supersonics Project, under NASA’s Fundamental Aeronautics Program, 

studies a number of technology challenges related to supersonic flight. These challenges include sonic 

boom, supersonic cruise efficiency, airport noise, high altitude emissions, lightweight engines/airframes, 

and multidisciplinary design. The present work was relevant to sonic boom reduction, which aims to 

mitigate the disturbance caused by the sonic boom and potentially remove the present aircraft operational 

restriction of supersonic flight over water only. A sonic boom is generated by coalescing shock waves and 

expansion fans formed by aircraft components, which generate an N-wave. The N-wave consists of a rise 

in pressure versus time as the aircraft “bow-wave” passes over an observer, followed by a reduction in 

pressure, and finally a return to atmospheric pressure. Previous work by NASA, such as the Shaped Sonic 
NASA/TM—2012-217229 2 

Boom Demonstrator (Ref. 1) (SSBD) and the Quiet Spike, (Ref. 2) has studied how the sonic boom 

signature generated by the front of the aircraft can be reduced with aircraft shaping. Complementary work 

was desired to reduce the sonic boom signature in the aft portion of the sonic boom N-wave, which in turn 

would decrease the peak-to-peak magnitude and result in a reduced sonic boom. This could be 

accomplished through the study of how aft components, such as the tail, nacelles, and nozzles contributed 

to the right hand portion of the sonic boom N-wave. 

Early work on exhaust nozzle contribution to sonic boom included a report by Putnam and Capone 

(Ref. 4), and another by Barger and Melson (Ref. 5). In the work by Putnam and Capone, nozzles were 

tested from a fully-expanded Mach 1.7 nozzle to a fully-expanded Mach 2.9 nozzle. Their study was 

conducted in a wind tunnel, where near-field pressure measurements were made at one nozzle diameter 

away from the model. Wind tunnel conditions were Mach 2.2 at a simulated altitude of 50,000 ft. In 2009, 

Castner (Ref. 6) performed a CFD analysis, where it was determined that over-expanded and underexpanded operation of a nozzle had an effect on the N-wave boom signature. The work demonstrated the 

feasibility of reducing the magnitude of the sonic boom N-wave by controlling the nozzle plume 

interaction with the nozzle boat-tail flow structure. Under-expanded exhaust nozzle flow had a favorable 

result on near field pressure signatures, suggesting that the nozzle plume shock location could be 

controlled through control of nozzle pressure or nozzle divergent section geometry. 

A complementary activity was performed to study nozzle plume effects in a flight environment. The 

flight test was performed using the NASA F-15 “Active” aircraft used for the Lift and Nozzle Change 

Effects on Tail Shock (LaNCETS) experiment (Ref. 3). The research approach for this experiment was to 

alter the aircraft shock structure by: (1) changing the lift distribution with canard position, (2) changing 

the plume shape with under-expanded and over-expanded nozzle flow, and (3) changing the plume shape 

using thrust vectoring. The test measured resulting shocks with a probing aircraft and used results to 

validate predictive tools. 

The capability to test small-scale models in NASA GRC’s 1- by 1-foot Supersonic Wind Tunnel 

(SWT) was developed to validate results obtained for supersonic nozzle concepts. This capability was 

utilized for a fundamental study of the LaNCETS vectored exhaust nozzle concept. The purpose of this 

report is to present the wind tunnel test results. The present work studied the effects of the vectored 

nozzle exhaust plume on near field pressure signatures, which directly correlate to sonic boom. Wind 

tunnel test results include schlieren images of the nozzle exhaust shock structure and near-field static 

pressure profiles as collected from a static pressure probe. 

2.0 Background 

The effects of under-expanded and over-expanded nozzle flow on the near field pressure signature 

was computed for a baseline nozzle (Ref. 6). Changes in nozzle pressure ratio (NPR) caused changes in 

the magnitude of the observed pressure signature, which would correspond to a change in the observed 

sonic boom signature. Figure 1 shows what happens when far-field pressure signatures were propagated 

to the ground. It was evident that as NPR was increased, the peak-to-peak magnitude for the right hand 

portion of the pressure signature was decreased. 

As a separate activity, a flight test was performed using the NASA F-15 “Active” aircraft (tail number 

837) for the LaNCETS experiment (Ref. 3). The research approach for this experiment was to alter the 

shock structure of NASA’s unique NF-15B airplane by: (1) changing the lift distribution with canard 

positions, (2) changing the plume shape with under-expanded and over-expanded nozzle flow, and (3) 

changing the plume shape using thrust vectoring. The test measured resulting shocks with a probing 

aircraft, NASA’s F-15B with tail number 836. Flight test results were used to validate predictive tools. 

Figure 2 demonstrates results collected from the flight test experiment, and also indicates that a 3

and 6 vector angle did not demonstrate significant differences in any peak values for the aircraft pressure 

signature. Results for the 8 vector angle are not shown, but were similar. For the flight test, a positive 

angle was an upward vector angle. 
NASA/TM—2012-217229 3 

Figure 1.—Estimated sonic boom signature at ground through an ideal 

atmosphere, NPR 6 to 10. 

Figure 2.—LaNCETS aircraft ∆P/P flight test results for Mach 1.2 at 40,000 ft altitude, nozzle vector 

angle of 3 and 6. 

In addition to the flight test, small-scale tests were conducted in the NASA GRC 1- by 1-foot SWT to 

collect wind tunnel data for isolated nozzles. Results were intended to provide fundamental understanding 

of the effect of a vectored nozzle on vehicle pressure signature. 
NASA/TM—2012-217229 4 

3.0 Test Setup 

The small-scale wind tunnel model, tested in the GRC 1- by 1-foot SWT was designed to operate 

across a variety of wind tunnel Mach numbers, nozzle pressure ratios, and simulated altitudes. This 

section is to review the wind tunnel model, instrumentation, and location in the wind tunnel. 

3.1 Model 

The wind tunnel model consisted of three major parts: the upper plenum, the strut, and the lower 

plenum (as displayed in Fig. 3). A photo of the rig installation is also provided in Figure 4. The upper 

plenum design was 1-in. outer diameter, and had an overall length from the nosecone to the nozzle exit of 

8.21 in. Both nosecone and nozzle were attached with fine pitched threaded connections. The support 

strut was 0.35 in. in thickness and contained nine passages to supply airflow. Air supply lines were 

0.25 in. diameter, and the passages supplied 90 psia of pressure to achieve 20 psia at the nozzle inlet for 

test conditions at 50,000 ft altitude. To accommodate larger nozzles and testing at lower altitudes, the 

Figure 3.—Wind tunnel model cut-away, dimensions in inches. 

Figure 4.—Wind tunnel model photograph, downward 

view through the top schlieren window. 

 
NASA/TM—2012-217229 5 

maximum design pressure at the strut was 233 psia, to achieve 52 psia at the nozzle inlet. Choked orifice 

holes were needed on top of the strut supply lines to obtain required mass flow. Orifice holes were 

0.210 in. diameter. Air was supplied to the bottom of the strut through a lower plenum by the centralized 

450 psig pressurized air system at the NASA Glenn Research Center. 

The location of the strut and model was designed to move forward and aft in the wind tunnel test 

section. This was accomplished with a slot, allowing 6 in. of overall travel. In this manner the rig location 

could be modified to avoid unforeseen wind tunnel flow effects or shock wave reflections. Manual 

changes to rig location could only be made between tests with the wind tunnel turned off. 

3.2 Instrumentation 

NPR was determined based on mass flow measurement and nozzle throat diameter. Sonic flow 

equations for a venturi were used from Blevins (Ref. 7) to perform this calculation. An external static 

pressure probe was designed to survey near-field static pressure profiles at one nozzle diameter above the 

rig centerline. The static pressure probe was mounted on an airfoil probe support, and a sliding ball screw 

actuator, from the opposite side of the tunnel. The probe passed through the tunnel sidewall and the 

actuator was installed in an enclosure to match the tunnel pressure. The overall assembly, showing the 

probe location, is provided in Figure 5. 

The static pressure probe had a capability to travel axially from x/D = 6.9 to 13.5 to capture pressure 

profiles, starting 0.65 in. upstream of the nozzle exit, and extending downstream to where the plume 

interacts with the tunnel/wall shock reflection. The probe design, based on Pickney (Ref. 8) and presented 

in Figure 6, had less sensitivity than a traditional static pressure probe design to angle of attack, while 

achieving relaxation of static pressure at the location of the measurement ports. These probes were 

mounted on a 3/8 in. thick airfoil strut (Fig. 5), allowing 5.625 in. of probe length between the static 

pressure holes and the airfoil support. 

The model was placed between two transparent sidewalls in the 1- by 1-foot SWT for schlieren 

photography. 

Figure 5.—Wind tunnel model cross section as installed in the 1- by 1-foot SWT test section, dimensions 

in inches. 
NASA/TM—2012-217229 6 

Figure 6.—Short static pressure probe design (Ref. 7). 

Figure 7.—Schlieren image of tunnel shock structure. 

3.3 Model Location 

The location of the model in the wind tunnel was a concern. Correct model location was necessary to 

provide a clean nozzle flow-field at the correct Mach number. The nozzle flow-field also had to be free of 

shock reflections and shock interactions. A schlieren image, Figure 7, was collected with the nozzle 

operating at an NPR of 8. Six types of shock waves were evident. It was a challenge to place the nozzle 

inside these shocks and provide a clean nozzle flow-field. Sources for shock waves in the tunnel were: 

1. Wind tunnel nozzle block 

2. Nosecone tip 

3. Nosecone and strut interaction 

4. Tunnel wall reflections 

5. Strut leading edge 

6. Nozzle plume (shock of interest) 

 
NASA/TM—2012-217229 7 

The shock wave from the wind tunnel nozzle block was generated by a small step in the tunnel wall. 

This step was located at the end of the interchangeable wind tunnel throat block, upstream of the model. 

Through careful model placement, the shock wave from the nosecone tip was co-located with the wind 

tunnel nozzle block shock. Shock waves from the nosecone and strut interaction were set by the strut 

geometry. This model design was kept as short as possible to avoid the wind tunnel wall reflections. The 

strut leading edge shock wave was a lateral shock wave that impacted the wind tunnel walls in a plane 

that appeared to interfere with the nozzle exhaust. However, the placement of this shock interaction was 

on the wind tunnel wall, 6 in. away from the nozzle plume. This interaction location was validated by pretest calculations and flow visualization. Similarly, the strut trailing edge shock and expansion did not 

interfere with the nozzle plume. Finally, the nozzle plume shock structure is visible, which was the 

objective of the experiment. 

To determine the interaction of the model strut shock wave with the wind tunnel glass sidewall, flow 

visualization was performed with a fluorescent paint and oil mixture. Results indicated clean flow at the 

nozzle location (Ref. 9). 

3.4 Nozzle 

The test nozzle was a 1:51.5 scale model of the nozzle installed on the NASA F-15 “Active” aircraft. 

The nozzles, (see samples for 0 vector angle in Fig. 8 and for 8 vector angle in Fig. 9) had convergentdivergent geometry with an area ratio of 1.5 and a fully expanded NPR of 6.2 at Mach 2.0. Tests were run 

at Mach 2.0 to reduce the impact of shock reflections on the test model. However, LaNCETS flight data 

for vectored nozzle configurations were collected at Mach 1.2 and 1.4. The difference in Mach number 

between the flight experiment and the wind tunnel test reflects the intent to provide fundamental 

understanding of the effect of a vectored nozzle on vehicle pressure signature, and not an intent to 

explicitly match the flight test data. 

Figure 8.—The 0 vectored nozzle, dimensions in inches. 
NASA/TM—2012-217229 8 

Figure 9.—8 vectored nozzle, dimensions in inches. 

4.0 Results 

4.1 Near-Field Experimental Results 

In the 1- by 1-foot SWT, the experimental pressure signatures were measured at 1-diameter away 

from the nozzle centerline, collected on the side opposite the model support strut. Data were collected at 

this location due to space restrictions in a small wind tunnel, and is consistent with previous experiments 

(Ref. 9) and comparisons to far-field pressure data. Experimental pressure signatures (ΔP/P∞) were 

collected over a 6 in. axial distance, from x/D = 6.9 to 13.5, at 0.1 in. increments. 

The near-field pressure dropped as flow expanded around the nozzle boat-tail. This expansion was 

followed by a rapid pressure rise due to the shock at the nozzle lip, and was followed by a secondary 

expansion and shock around the nozzle plume. As NPR was increased, the pressure rise due to the lip 

shock increased. 

Figure 10 shows results for the 0 vectored nozzle case. The plot demonstrates trends in the near field 

pressure as NPR was increased from 6.2 to 10; the strength of the nozzle lip shock increased, and the 

shock angle increased with respect to the nozzle exit centerline. The increase in shock strength caused a 

rise in the pressure, as measured at the first peak. The reduction in the first expansion was caused by the 

change in the shock angle, where the lip shock cut off the boat-tail expansion. As seen in results from Bui 

(Ref. 10), this nozzle had a dramatic boat-tail angle, and the external flow around the nozzle separated 

and formed an aerodynamic boat-tail, which created a low pressure expansion that dominated the pressure 

signature. 

Nozzles were tested at 3, 6, and 8 of vector angle. Results are presented in Figure 11 for the 

–8 vector angle. When compared to the 0 vectored nozzle results, the boat-tail expansion did not 

dramatically change with vector angle at –8. At NPR = 6.2 the lowest value of ∆P/P was –3.74, and can 

be compared to a similar value of –3.84 for the 0 vectored nozzle case. This is evidence that the high 

boat-tail angle caused an expansion that dominated the near field pressure signature. As the nozzle was 

vectored down, the shock from the nozzle lip was weaker than in the 0 vector case. As shown at NPR of 

6.2, the values for the first peak in the pressure signature decreased to a ∆P/P of –0.150 for –8 versus the 

∆P/P of –0.108 for 0. Values for the secondary expansion and shock around the nozzle plume were not 

greatly affected. Increasing NPR from 6.2 to 12 demonstrated a similar trend as seen in the 0 vector 

angle. The overall difference in peak-to-peak value of the pressure signature did not change significantly. 
NASA/TM—2012-217229 9 

Figure 10.—0 vectored nozzle ∆P/P, where x = 0 at the nosecone tip. 

Figure 11.—Minus 8 vectored nozzle ∆P/P. 

 
NASA/TM—2012-217229 10 

Figure 12.—Plus 8 vectored nozzle ∆P/P. 

Results for the positive 8 vector angle are displayed in Figure 12. Again the minimum value of 

–0.377 for ∆P/P did not change significantly, and demonstrated that the nozzle boat-tail angle again 

dominated the pressure signature. However, for the 8 vector angle, the nozzle lip shock strength 

increased over both the –8 and 0 vectored nozzle case, and measured 0.026 instead of –0.108 for the 0

vectored nozzle case. As NPR was increased, the nozzle lip shock became stronger, and moved forward. 

This was the only case where NPR greater than 12 were tested to clearly demonstrate the trends for the 

nozzle lip shock, and to demonstrate that the strong lip shock could reduce the expansion around the 

nozzle boat-tail. This reduction is shown in Figure 12. These trends agree with previous supersonic 0

vectored nozzle test results collected by Castner (Ref. 9). Based on the previous conclusions, if these 

shocks were allowed to propagate to the far-field, the effect would be a reduction in the peak-to-peak 

pressure differential. 

In Figure 13, trends are shown for both increasing and decreasing nozzle vector angle for a constant 

NPR of 12. The result shows that the vector angle had an impact on the near field pressure signature. The 

effect was most noticeable in the increasing strength of the nozzle lip shock, but not on the level of boattail expansion. In the case of this nozzle design, the boat-tail expansion was the dominant flow feature. 

Past work on nozzles with lower boat-tail angle showed that increasing NPR had an impact on the peakto-peak level of the near field pressure signature. In this work, the vector angle was expected to have a 

similar effect and change the angle of the lip shock, but it did not. This result suggests that care should be 

taken to reduce the nozzle boat-tail angle as much as possible for low boom supersonic aircraft design. 

 
NASA/TM—2012-217229 11 

Figure 13.—∆P/P for vector angles from –8 to 8 at NPR 12. 

4.2 Schlieren Results 

The 0 vector angle nozzle is displayed in Figure 14, where the location of the lip shock is 

highlighted for NPR of 8. The lip shock is difficult to trace back to the nozzle lip, as it was formed off the 

interaction between the nozzle boat-tail and the lip shock. This could also be seen in CFD results for the 

nozzle geometry with exposed convergent and divergent flaps from Bui (Ref. 10), and is illustrated in 

Figure 15. 

Further evidence of the dominant lip shock location can be observed in Figure 16 which shows a 

schlieren image from –8 of vector angle. In this case, the shock location was drawn from the 0 case in 

Figure 14, and overlaid on the shock location for –8. There was no noticeable change in the shock 

location, which was also observed in the pressure traces where levels of ∆P/P for the boat-tail expansion 

were similar in magnitude. 

Schlieren results are also presented from the 8 vector angle case in Figure 17. This image 

demonstrates that the shock location moves and changes angle to a very small degree, due to the dominant 

effects of the boat-tail angle on the nozzle lip shock location. 

The effects of increasing NPR are presented in Figure 18 at an NPR of 12. The lip shock location is 

again highlighed and then superimposed on Figure 19, which is an image at a very high NPR value of 25. 

NPRs above 12 were needed to get the nozzle lip shock to have a noticeable movement, as was 

demonstrated in previous experiments and analyses. As NPR was increased from 12 to 25, the lip shock 

moved forward and changed angle with respect to the nozzle centerline. This result was confirmed in 

measurements, see Figure 12, where the nozzle boat-tail expansion was not as strong and the minimum 

∆P/P value increased from –0.35 to –0.29. 
NASA/TM—2012-217229 12 

Figure 14.—0 vector angle nozzle schlieren image, NPR 8. 

Figure 15.—0 vector angle nozzle CFD from Bui (Ref. 10). 
NASA/TM—2012-217229 13 

Figure 16.—Minus 8 vectored nozzle schlieren image, NPR 8. 

Figure 17.—Plus 8 vectored nozzle schlieren image, NPR 8. 
NASA/TM—2012-217229 14 

Figure 18.—Sample schlieren images of 8 vectored nozzle plume for NPR 12. 

Figure 19.—Sample schlieren images of 8 vectored nozzle plume for NPR 25. 
NASA/TM—2012-217229 15 

5.0 Conclusions 

Background was provided on vectored nozzles used to study nozzle plume effects on sonic boom. 

Previous analysis and experiments showed how a shock wave formed at the nozzle lip and interacted with 

the nozzle boat-tail expansion. 

An experiment was designed and conducted in the 1- by 1-foot Supersonic Wind Tunnel. The wind 

tunnel model consisted of a test nozzle supported by an upper plenum and strut. Instrumentation allowed 

calculation of NPR and measurement of near-field static pressure, presented as ΔP/P∞. Schlieren images 

and flow visualization were used to validate proper placement of the model in the wind tunnel to avoid 

shock reflections and provide for a clean nozzle flow-field. 

Results were similar to previous experiments where the near field pressure dropped as flow expanded 

around the nozzle boat-tail. This expansion was followed by a rapid pressure rise due to the shock at the 

nozzle lip, and was followed by a secondary expansion and shock around the nozzle plume. As NPR was 

increased, the pressure rise due to the lip shock increased. As the nozzle was vectored down, the 

magnitude of the lip shock decreased, while trends with increasing NPR remained the same. As the nozzle 

was vectored up, or towards the measurement plane, the magnitude of the lip shock strengthened. Trends 

again remained the same as NPR was increased. 

For small changes in NPR, the changes in the nozzle shock location did not affect the boat-tail 

expansion, as they did in previous results. This was due to the severe boat-tail angle for this geometry, 

which caused a large separation and expansion which dominated the pressure signature. A change in this 

expansion pressure, which would reduce sonic boom, was not realized until very high NPR was set. In the 

case of this nozzle design, the boat-tail expansion was the dominant flow feature. Past work on nozzles 

with lower boat-tail angle showed that increasing NPR had an impact on the peak-to-peak level of the 

near field pressure signature. In this work, the vector angle was expected to have a similar effect and 

change the angle of the lip shock, but did not. Vectored nozzles with reduced boat-tail angle were not 

tested for comparison, but would make a valid follow on study. If vector angle were to be beneficial for 

sonic boom, the shock location for the lip shock would need to interrupt the expansion wave from the 

nozzle boat-tail, as demonstrated in previous studies. 

The wind tunnel test results are consistent with the LaNCETS flight test data, where significant 

effects were not present due to vectoring the nozzle up and down at 3, 6, and 8 of vector angle. While 

this experiment was not a direct comparison due to differences in aircraft geometry and flight Mach 

number, it demonstrates consistency between the fundamental research and the flight test demonstration. 

References 

1. Graham, D., et al.: Aerodynamic Design of Shaped Sonic Boom Demonstration Aircraft. 

AIAA−2005−0009, 2005. 

2. Freund, D., et al.: Quiet Spike Prototype Aerodynamic Characteristics From Flight Test. 

AIAA−2008−125, 2005. 

3. Moes, T.: Sonic Boom Research at NASA Dryden: Objectives and Flight Results from the Lift and 

Nozzle Change Effects on tail Shock (LaNCETS) Project, International Test & Evaluation 

Association, Feb. 2009. 

4. Putnam, L. and Capone, F.: Experimental Determination of Equivalent Solid Bodies to Represent Jets 

Exhausting into a Mach 2.20 External Stream, NASA TN-D-5553. 

5. Barger, R.L. and Melson, N.D.: Comparison of Jet Plume Shape Predictions and Plume Influence on 

Sonic Boom Signature, NASA TP-3172. 

6. Castner, R.S.: Analysis of Plume Effects on Sonic Boom Signature for Isolated Nozzle 

Configurations, NASA/TM—2008-215414. 

7. Blevins, R.D.: Applied Fluid Dynamics Handbook, Krieger, Florida, 1992 pp. 126-135. 

 
NASA/TM—2012-217229 16 

8. Pinckney, S.Z.: A Short Static Pressure Probe Design for Supersonic Flow, NASA TN D-7978. 

9. Castner, R.S.: Exhaust Nozzle Plume Effects on Sonic Boom Test Results for Isolated Nozzles, 

AIAA–2010–4936, 2010. 

10. Bui, Trong T.: “Computational Fluid Dynamics Analysis of Nozzle Plume Effects on Sonic Boom 

Signature,” AIAA Journal of Aircraft, Vol. 48, No. 2, March 2011, pp. 368–380. 
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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-17790 

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-2012-217229 

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited 

Subject Categories: 01, 05, and 07 

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

Reducing or eliminating the operational restrictions of supersonic aircraft over populated areas has led to extensive research at NASA. 

Restrictions were due to the disturbance of the sonic boom, caused by the coalescence of shock waves formed off the aircraft. Recent work 

has been performed to reduce the magnitude of the sonic boom N-wave generated by airplane components with a focus on shock waves 

caused by the exhaust nozzle plume. Previous Computational Fluid Dynamics (CFD) analysis showed how the shock wave formed at the 

nozzle lip interacts with the nozzle boat-tail expansion wave. An experiment was conducted in the 1- by 1-foot Supersonic Wind Tunnel 

(SWT) at the NASA Glenn Research Center. Results show how the shock generated at the nozzle lip affects the near field pressure

signature, and thereby the potential sonic boom contribution for a nozzle at vector angles from 3° to 8°. The experiment was based on the 

NASA F-15 nozzle used in the Lift and Nozzle Change Effects on Tail Shock experiment, which possessed a large external boat-tail angle. 

In this case, the large boat-tail angle caused a dramatic expansion, which dominated the near field pressure signature. The impact of nozzle 

vector angle and nozzle pressure ratio are summarized.

15. SUBJECT TERMS

Exhaust nozzles; Plumes; Sonic booms 

16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF

 ABSTRACT

UU 

18. NUMBER

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 PAGES

22 

19a. NAME OF RESPONSIBLE PERSON

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

a. REPORT

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b. ABSTRACT

U 

c. THIS 

PAGE

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19b. TELEPHONE NUMBER (include area code)

443-757-5802 

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