Raymond S. Castner

Glenn Research Center, Cleveland, Ohio

Exhaust Nozzle Plume Effects on Sonic Boom 

Test Results for Isolated Nozzles

NASA/TM—2011-217137

December 2011

AIAA–2010–4936
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Raymond S. Castner

Glenn Research Center, Cleveland, Ohio

Exhaust Nozzle Plume Effects on Sonic Boom 

Test Results for Isolated Nozzles

NASA/TM—2011-217137

December 2011

AIAA–2010–4936

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

28th Applied Aerodynamics Conference

sponsored by the American Institute of Aeronautics and Astronautics

Chicago, Illinois, June 28–July 1, 2010
Acknowledgments

The author recognizes the contributions of Allison Ronning, from Wichita State University, who performed significant analysis. 

In addition, Martin Silverman, John Jones and Warren Holt 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. 

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

Available from

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7115 Standard Drive

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

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—2011-217137 1

Exhaust Nozzle Plume Effects on Sonic Boom 

Test Results for Isolated Nozzles

Raymond S. 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 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-ft Supersonic Wind Tunnel at the NASA Glenn Research Center 

to validate the computational study. Results demonstrated how the nozzle lip shock moved with 

increasing nozzle pressure ratio (NPR) and reduced the nozzle boat-tail expansion, causing a favorable 

change in the observed pressure signature. Experimental results were presented for comparison to the 

CFD results. The strong nozzle lip shock at high values of NPR intersected the nozzle boat-tail expansion 

and suppressed the expansion wave. Based on these results, it may be feasible to reduce the boat-tail 

expansion for a future supersonic aircraft with under-expanded nozzle exhaust flow by modifying nozzle 

pressure or nozzle divergent section geometry.

Nomenclature

β nozzle boat tail angle, °

D test nozzle diameter, in.

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∞

t time, sec

x axial distance from jet simulator nosecone tip, in.

y distance from nozzle centerline, in.

Introduction

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

by supersonic flight speeds. Current aircraft flight restrictions, allowing supersonic flight over water only, 

are due to the disturbance caused by the sonic boom. The sonic boom is generated by coalescing shock 

waves formed by aircraft components, which generate an N-wave. A sample N-wave sonic boom 

signature is shown in Figure 1. The N-wave consists of a rise in pressure versus time as the aircraft ‘bowwave’ 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 Boom Demonstrator (Ref. 1) (SSBD) and 

the Quiet Spike (Ref. 2) have studied how the sonic boom signature generated by the front of the aircraft 

can be reduced with aircraft shaping.
NASA/TM—2011-217137 2

Figure 1.—Ground-level shaped sonic boom recording in blue, baseline 

F-5E N-Wave in red, SSBD flight 9 (Ref. 1).

Complementary work is 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 can be accomplished by aircraft shaping, and through study of how aft components, such as 

the tail, nacelles, and nozzles contribute to the right hand portion of the sonic boom N-wave.

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

(Ref. 3), and another by Barger and Melson (Ref. 4). 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. 5) revisited this work to use as a baseline for studies on the exhaust nozzle contribution to 

sonic boom. Near-field pressure signatures from the wind tunnel test were examined, where stronger 

pressure signatures were observed for high levels of nozzle pressure ratio (NPR). From the Putnam and 

Capone report, a baseline nozzle was selected for an axisymmetric Computational Fluid Dynamics (CFD) 

study. During the CFD analysis, it was determined that over-expanded and under-expanded operation of 

the 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.

Results from the CFD study were used to develop wind tunnel test capability to both validate and 

predict results obtained for supersonic nozzle concepts. Testing has been performed and comparisons 

made to the CFD predictions for the baseline nozzle. Post-test analysis has been performed to analyze 

differences between test data and CFD predictions. The purpose of this report is (1) to present the wind 

tunnel test results, (2) to make comparisons to the pre-test CFD results, and (3) to make comparisons to 

post-test CFD analysis. The present work studies the effects of the nozzle exhaust plume and the 

contribution to the right hand portion of the sonic boom N-wave. 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.
NASA/TM—2011-217137 3

Nozzle Mj,des (Pt,j/P∞)des α,

deg.

β,

deg.

x1 L k s dth de db Ath Ae Ae/Ath x2

1 2.920 32.58 11.50 0 15.240 119.380 15.240 18.519 7.577 15.011 15.240 45.09 176.98 3.925 2.908

2 2.272 11.97 7.28 0 15.240 119.380 15.240 18.747 10.264 15.011 15.240 82.75 176.98 2.139 4.224

3 2.740 24.77 11.50 0 12.192 116.332 12.192 16.878 8.255 15.011 15.240 53.52 176.98 3.307 2.118

4 2.523 17.72 11.50 0 9.144 113.284 9.144 14.883 9.144 15.011 15.240 65.66 176.98 2.695 2.118

5 2.267 11.88 9.06 0 9.144 113.284 9.144 15.011 10.288 15.011 15.240 83.12 176.98 2.129 1.953

6 2.024 8.12 6.04 5 9.109 113.249 9.144 15.105 10.223 13.417 13.646 82.08 141.39 1.722 1.778

7 1.700 4.94 3.04 10 8.999 113.139 9.144 15.166 10.231 11.836 12.065 82.21 110.03 1.338 1.580

Figure 2.—Baseline Nozzle, from Putnam (Ref. 3).

Background

A baseline nozzle was developed for testing in the NASA Glenn Research Center (GRC) 1-ft by 1-ft

Supersonic Wind Tunnel (Ref. 6). Figure 2 depicts the design of the Mach 2.0 baseline supersonic 

exhaust nozzle configuration, which was a replica of the Putnam ‘Nozzle 6.’ Previous computational and 

wind tunnel tests were performed on this nozzle at flight speeds of Mach 2.2 and an altitude of 50,000 ft.

Figure 3 shows the larger scale wind tunnel model used by Putnam and Capone, where an x/D of zero is 

the tip of the nose cone. Near-field pressure measurements were made at one nozzle diameter above the 

test nozzle centerline. Data were taken from ‘Nozzle 6’, with a 5° boat tail angle, which was digitized 

from the report and displayed in Figure 4. Previous CFD data were obtained to match the experimental 

test conditions (Castner (Ref. 5).

Pressure signatures for ‘Nozzle 6’, plotted as a ratio of ∆P/P∞, were compared to Putnam’s

experimental data. The results from the comparison are plotted in Figure 4 for a NPR of 8, which was the 

design pressure ratio. Agreement between the experiment and the axisymmetric CFD was fair, and 

displays a reduction in the near-field static pressure caused by the flow expansion off the nozzle boat-tail.

The expansion was followed by a shock at the nozzle lip, and then an expansion back to flight conditions.

(Note that the near-field pressure distribution is different from an ‘N-wave’. With propagation, this 

disturbance takes the shape of an N-wave, as demonstrated in Figure 5(c)).
NASA/TM—2011-217137 4

Figure 3.—Exhaust nozzle simulator from Putnam (Ref. 3).

Figure 4.—Near-field pressure profile at 1 nozzle diameter from baseline 

'Nozzle 6', NPR of 8.

While NPR of 8 produces fully expanded flow, simulations were also conducted at NPR of 6 and 7, 

resulting in over-expanded nozzle flow; and then simulations were conducted at NPR of 9 and 10, 

resulting in under-expanded nozzle flow. The effects of under-expanded and over-expanded nozzle flow 

were computed at a distance far from the nozzle, 150 in. away, or 10 nozzle diameters away. Changes in 

NPR caused changes in the magnitude of the observed pressure signature, which would correspond to a 

change in the observed sonic boom signature and are shown in Figure 5(a). Locations of the cut planes for 

the near field and far-field CFD data are shown in Figure 5(b). Finally, Figure 5(c) shows what happens to 

the far-field pressure signatures when they are propagated to the ground. In Figure 5(c) it is evident that 

as NPR was increased, the peak-to-peak magnitude for the right hand portion of the pressure signature 

was decreased. These were results intended for validation in the current experiment, conducted in the 

GRC 1- by 1-ft Supersonic Wind Tunnel.
NASA/TM—2011-217137 5

Figure 5.—(a) Far-field pressure profiles ∆P/P∞ at y = 150 in from nozzle centerline. 

Baseline 'Nozzle 6' NPR 6 to 10. (b) Location of cut planes for CFD predictions of 

∆P/P∞. (c) Estimated sonic boom signature at ground through an ideal atmosphere, 

NPR 6 to 10.
NASA/TM—2011-217137 6

Test Setup

The wind tunnel model for the experiment was designed to operate across a variety of wind tunnel 

Mach numbers, nozzle pressure ratios, and simulated altitudes. Design Mach numbers ranged from 1.4 to 

3.0. The range of NPRs were from 6 to 12, and altitudes were from 30,000 ft (4.36 psia) to 50,000 ft 

(1.68 psia). The actual test nozzle was a replica of Putnam’s (Ref. 3) ‘Nozzle 6’. This section is to review 

the wind tunnel model, instrumentation, and model location in the wind tunnel.

Model

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

plenum, as displayed in Figure 6. A photo of the rig installation is also provided in Figure 7. The upper 

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

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

was 3/8 in. in thickness and contained nine passages to supply airflow. Air supply lines were 1/4 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 

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

centralized 450 psig pressurized air system at GRC.

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.

Figure 6.—Wind tunnel model cut-away.
NASA/TM—2011-217137 7

Figure 7.—Wind tunnel model—side view.

Figure 8.—Wind tunnel model cross section as installed in the 1- by 1-ft SWT test section.

Instrumentation

Instrumentation was located internally to measure NPR, and was limited to one total pressure probe 

upstream within the test nozzle, located on centerline, and one static pressure tap. There were two 

problems with the internal total pressure probe. First, the location was too close to the air supply passages 

in the strut, and the measurement was unsteady. Second, the internal total pressure probe was fragile and 

failed, thus the nozzle inlet instrumentation was replaced with a calculation of the NPR, based on mass 

flow measurement and nozzle throat diameter. Sonic flow equations for a venturi were used from Blevins

(Ref. 9) 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 8.

The static pressure probe had a capability to travel axially 8.4 in. to capture pressure profiles, starting 

1.6 in. upstream of the nozzle exit, and extending downstream to where the plume interacts with the 

tunnel/wall shock reflection. Two static pressure probes were built, a conventional and a short design,

based on Pinckney (Ref. 7). The conventional probe design, in Figure 9, had difficulty capturing shock 

waves. This difficulty was caused by the long distance between the probe tip and the static pressure 
NASA/TM—2011-217137 8

measurement ports, where the measurement was not in the same flow environment as the probe tip. The 

short probe design, Figure 10, had less sensitivity to angle of attack, while achieving similar relaxation of 

static pressure at the location of the measurement ports. These probes were mounted on a 3/8 in. thick 

airfoil strut, allowing 5.625 in. of probe length between the static pressure holes and the airfoil support.

For all experimental results presented, the shorter probe was used.

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

photography. Schlieren images were compared to Mach number contours generated by CFD analysis.

Model Location

The location of the model in the wind tunnel was a concern. Correct model location provided 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 11, 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).

Figure 9.—Conventional static pressure probe (Ref. 7).

Figure 10.—Short static pressure probe design (Ref. 7).
NASA/TM—2011-217137 9

Figure 11.—Schlieren image of tunnel shock structure.

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 and 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 fixed by 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 really on the wind tunnel wall, 6 in. away from the nozzle plume.

This was validated by pre-test calculations and flow visualization, presented in Figures 12(a) and (b).

Finally, the nozzle plume shock structure is visible, which was the objective of the experiment. 

To validate 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 on the 

sidewall, except where a half-span shock from the support strut impinged on the glass sidewall. The oil 

and paint mixture, located on the wind tunnel walls, showed axial streaking up to the shock reflection, 

visible on the left side of Figure 12(a). At this point, the paint was removed from the sidewall by the 

shock interaction with the thick boundary layer on the tunnel wall, which caused flow separation and 

resulted in the oil pattern. Figure 12(b). also shows the shock wave pattern on the model strut and 

nosecone, and validated the presence of the nosecone and strut interaction shock wave.

As a final validation, the static pressure survey probe was replaced with a pitot probe. This probe was 

placed 1 in. away from the test nozzle centerline, and 1.63 in. upstream of the nozzle exit lip. Local pitot

pressure and local static pressure were measured with the probes during two separate wind tunnel runs

and used to compute a local Mach number. The local Mach number at this location was 1.956, which was 

very close to the tunnel set point of Mach 1.96.
NASA/TM—2011-217137 10

Figure 12.—(a) Oil flow visualization photo; view from upstream showing shock footprint on side 

wall. (b) Oil flow visualization photo; side view showing shock footprint on the strut.

Results

This section compares results from pre-test CFD predictions with the experimental results from the 

GRC 1-ft by 1-ft Supersonic Wind Tunnel. Comparisons include schlieren images, CFD contour plots 

(Mach number), and static pressure surveys at one nozzle-diameter above the nozzle

Near-Field CFD Results

Pre-test CFD results were obtained with a two-dimensional model of ‘Nozzle 6’ from Putnam and 

Capone (Ref. 3). Detailed CFD results are presented in Castner(Ref. 5), and included the ∆P/P at 150 in.

away from the nozzle centerline. The previous Figure 5(a) shows the pressure signature results at 150 in.

away, and were considered far-field results for this discussion. In the far-field, pressure signatures were

provided for NPRs from 6 to 10. All the signatures have the same characteristic. The first reduction in 

pressure is the expansion off the nozzle boat-tail. This is followed by a shock at the nozzle lip, which is 

then followed by a secondary expansion and shock off the nozzle exhaust plume. The first reduction in 

pressure, caused by nozzle boat-tail expansion, is cut short by the shock at the nozzle lip as NPR is 

increased from 6 to 10.

Near-field CFD results are shown in Figure 13. Comparisons with wind tunnel data were not possible 

at a location 150 in. away from the nozzle, given the 1- by 1-ft wind tunnel cross section. Comparisons 

were made with near-field static pressure surveys, which were performed at one nozzle diameter above 

the nozzle centerline, on the side opposite the model support strut. From the CFD, the near-field pressure 

pattern consisted of an expansion from the nozzle boat-tail, followed by a strong shock at the nozzle lip.

In the near-field, the nozzle lip shock appeared to dominate the pressure signature as NPR was increased 

from 6 to 10. As NPR was increased the location of the nozzle lip shock moved a small amount and the 

shock strength increased. Also, the stronger lip shock caused a weaker secondary expansion and shock 

around the nozzle plume.

Figure 14(a) shows a color contour plot of Mach number for the CFD solution. In this plot, the boat-tail 

expansion, nozzle lip shock, plume expansion and plume shock are visible. If the nozzle lip shock was 

observed far from the nozzle, the location of the shock changed with NPR. This can be seen from 

Figure 14(b) and (c). The location of the nozzle lip shock at NPR 6 is highlighted, and then superimposed on 

the contour plot for NPR 10. In this manner the change in nozzle lip shock location can be demonstrated.
NASA/TM—2011-217137 11

Figure 13.—Predicted Near-filed pressure profiles ∆P/P∞ at y = 1.0 in from nozzle 

centerline. Baseline 'Nozzle 6' NPR 6 to 10, Mach 2.2.

Figure 14.—(a) 'Nozzle 6' CFD Mach Number Contour NPR 8. (b) 'Nozzle 6' CFD Mach Number Contour NPR 6. 

(c) 'Nozzle 6' CFD Mach Number Contour NPR 10.
NASA/TM—2011-217137 12

The angle of the nozzle lip shock increased with respect to the nozzle centerline. When observed at the 

measurement location, the nozzle lip shock moved forward and suppressed the nozzle boat-tail expansion, 

and caused a corresponding reduction in the boat-tail expansion, and a favorable change in the observed 

pressure signature. A reduction in boat-tail expansion could be exploited to reduce the right hand portion 

of the sonic boom N-wave, thereby reducing sonic boom signature.

Near-Field Experimental Results

In the 1- by 1-ft Supersonic Wind Tunnel, the experimental pressure signatures were measured at 

one-diameter away from the ‘Nozzle 6’ centerline, collected on the side opposite the model support strut.

Data were collected at this location due to space restrictions, as far-field data at 150 in. away was not 

possible in a small wind tunnel. Experimental pressure signatures (ΔP/P∞) were collected over a 6 in.

axial distance at 0.1 in. increments, and are in shown in Figure 16. The same pressure pattern was present 

and consisted of an expansion from the nozzle boat-tail, followed by a strong shock at the nozzle lip. In 

the near-field, the nozzle lip shock appeared to dominate the pressure signature as NPR was increased 

from 6 to 10. The location of the nozzle lip shock moved a small amount as NPR was increased and the 

shock strength increased as NPR was increased. Also, in cases with a stronger lip shock, a weaker 

secondary expansion and shock around the nozzle plume was observed.

Comparisons to the previous CFD results show that the trends in the experimental data were the same 

as computational results; however, the values were different. For example, the following comparisons 

were taken from data at NPR 8: for the boat-tail expansion the minimum ΔP/P∞ value from CFD was 

–0.209 while the minimum experimental value was –0.172. For the nozzle lip shock the CFD value was 

0.118 while the experimental value was –0.018. The trend in the experimental data was the same as the 

CFD, where increasing levels of NPR provided a stronger nozzle lip shock and would provide the same 

results if the shock waves could propagate to the far-field. The stronger nozzle lip shock at high NPR 

would again intersect the boat-tail expansion earlier and suppress the expansion wave. Evidence of the 

movement of the nozzle lip shock can be seen in the pressure signatures in Figure 15, and also in the 

schlieren images. The data for the NPR of 13 was skipped to save time during the experiment.

Figure 15.—Experimental results, 'Nozzle 6' ∆P/P, Mach 1.96.
NASA/TM—2011-217137 13

Figure 16.—(a) Schlieren images of plume for NPR 6. (b) Sample schlieren images of plume for NPR 8, 10, 

11 and 12. (c) Schlieren images of plume for NPR 14
NASA/TM—2011-217137 14

Schlieren images are shown in Figure 16, and also demonstrate the same flow features as presented 

in CFD contour plots of Mach number. A description of the extra shock structures seen in the images

was provided in a previous section. In the schlieren images, the following flow features can be observed: 

(1) the expansion wave from the nozzle boat-tail, (2) the shock at the nozzle lip, (3) the secondary expansion 

from the nozzle plume, and (4) the secondary shock from the nozzle plume. Figure 16 illustrates how the 

shock location moves with NPR. For an NPR of 6, the location of the nozzle lip shock was highlighted in 

Figure 16(a). The shock location from the NPR 6 plume was superimposed on the schlieren image for an

NPR of 14 (Fig. 16(c)) and demonstrates the location of the lip shock moved upstream. In the near-field, the 

change is subtle, but becomes noticeable at greater distances from the nozzle.

If these shock waves for the under-expanded nozzle cases, greater than NPR=8, were allowed to 

propagate to the far-field the effect would be a reduction in peak-to-peak pressure differential. In practice, 

creating under-expanded flow can be achieved with increasing nozzle pressure or changing divergent 

section geometry (Stitt (Ref. 8)). The shock from the nozzle lip suppresses the nozzle boat-tail expansion

and a reduction in expansion could improve the right hand portion of the sonic boom signature. From 

previous studies (Castner (Ref. 5)), the thrust coefficient of ‘Nozzle 6’ ranged from 0.98 at NPR of 6 to 

0.991 at NPR of 10, resulting in reasonable nozzle performance.

Differences Between CFD and Experimental Results

Differences between the CFD results and the experimental results were evident in the schlieren 

images. Close examination shows that the shock wave from the nozzle lip was anchored at the nozzle lip, 

but curved away from the nozzle, see Figure 17. When observed at the measurement plane, this curvature 

caused the location of the nozzle lip shock to change, and intersect the nozzle boat-tail expansion earlier 

than in the CFD results. In computational predictions, the shock wave formed off of the nozzle lip with no

curvature. It was also evident in the schlieren images that the shock waves from the top of the nozzle and 

the bottom of the nozzle were not the same. The cause for the change in shape for the nozzle lip shock 

appears to be deflection of the jet plume, due to influence of the wind tunnel model support strut. This can 

also be seen in Figure 18 as the shock diamond in the nozzle plume is not on nozzle centerline.

The deflection of the jet plume was viewed in the schlieren images, and was also validated with posttest CFD analysis, provided in Figure 18 and 19. Post-test CFD included a three-dimensional model of the 

entire wind tunnel model at the proper model scale, while pre-test CFD was performed on an axisymmetric 

model which did not include the nosecone or strut. Levels of pressure profiles from the three-dimensional 

CFD analysis did not match the experimental results, see Figure 19. The boat tail expansion was captured 

well, but the strength of the nozzle lip shock was stronger in three-dimensional CFD than in the experiment.

Figure 17.—Schlieren image showing curvature of 

nozzle lip shock, NPR of 8.
NASA/TM—2011-217137 15

Figure 18.—Three-dimensional CFD analysis results: Mach number contour.

Figure 19.—'Nozzle 6' Comparison of three-dimensional CFD analysis to 

experimental data at NPR of 8.

Despite these issues, use of this model in the 1- by 1-ft SWT did demonstrate the main effects of the 

nozzle plume structure, and validated how the nozzle boat-tail expansion is reduced by the shock at the 

nozzle lip.

Conclusions

Background was provided on a baseline nozzle used to study nozzle plume effects on sonic boom.

Previous CFD analysis showed how the shock wave was formed at the nozzle lip and interacted with the 

nozzle boat-tail expansion wave. 

An experiment was designed and conducted in the 1-ft by 1-ft Supersonic Wind Tunnel to validate 

the computational study. The wind tunnel model consisted of a test nozzle supported by an upper plenum 

and strut. Instrumentation allowed measurement of NPR and near-field static pressure, presented as
NASA/TM—2011-217137 16

∆P/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.

Pre-test axisymmetric CFD results were presented at one-diameter above the test nozzle centerline, as 

it was not possible to get far-field pressure data in a small wind tunnel. Results show how the nozzle lip 

shock moved with increasing NPR and reduced the nozzle boat-tail expansion, causing a favorable 

change in the observed pressure signature.

Experimental results were presented for comparison to the CFD results. Experimental pressure 

signatures at one-diameter above the nozzle centerline were consistent with the pressure signatures from 

analysis, however, the values were different. The strong nozzle lip shock at high values of NPR intersect 

the nozzle boat-tail expansion and suppress the expansion wave.

Differences between the pre-test CFD and the experiment were studied with a post-test threedimensional CFD analysis. The main cause of the differences was attributed to the curvature of the shock 

wave from the nozzle lip observed at the top of the nozzle in experimental results. This curvature resulted 

in a change in the location of the nozzle lip shock when compared to analysis. The change in the shape of 

the lip shock was caused by upward deflection of the nozzle plume, which was visible in the threedimensional CFD and the schlieren photos from the experiment. The curved shock wave intersected the 

nozzle boat-tail expansion earlier and caused additional suppression of the boat-tail expansion.

In this experiment, under-expanded nozzle flow was achieved by increasing NPR. Under-expanded 

nozzle flow can also be achieved through changes in divergent section geometry. The shock from the

nozzle lip suppresses the boat-tail expansion, and increasing under-expanded flow causes stronger shock 

waves at an increased angle with respect to the nozzle centerline. The increased shock angle causes 

increased suppression of the flow expansion around the external boat-tail geometry. A reduction in the 

expansion can decrease the magnitude of the right hand portion of the sonic boom signature. The results 

also suggest that a future study on exhaust nozzle vector angle may be of interest to the sonic boom 

community.

The present analysis and experiment developed a set of consistent trends that demonstrate the 

feasibility of modifying the right hand portion of the sonic boom N-wave caused by the boat-tail 

expansion, through a change in nozzle pressure or divergent section geometry.

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. 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.

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

Sonic Boom Signature, NASA TP-3172.

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

Configurations, NASA/TM—2008-215414.

6. Soeder, R., Stark. D., Leone, J., and Henry, M.: NASA Glenn 1- by 1- Foot Supersonic Wind Tunnel 

User Manual, NASA/TM—1999-208478.

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

8. Stitt, L. E.: Exhaust Nozzles for Propulsion Systems with Emphasis on Supersonic Cruise Aircraft, 

NASA RP-1235.

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

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4. TITLE AND SUBTITLE

Exhaust Nozzle Plume Effects on Sonic Boom Test Results for Isolated Nozzles 

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6. AUTHOR(S)

Castner, Raymond, S. 

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WBS 984754.02.07.03.13.05 

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National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field 

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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 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-ft Supersonic Wind Tunnel at the 

NASA Glenn Research Center to validate the computational study. Results demonstrated how the nozzle lip shock moved with increasing 

nozzle pressure ratio (NPR) and reduced the nozzle boat-tail expansion, causing a favorable change in the observed pressure signature. 

Experimental results were presented for comparison to the CFD results. The strong nozzle lip shock at high values of NPR intersected the 

nozzle boat-tail expansion and suppressed the expansion wave. Based on these results, it may be feasible to reduce the boat-tail expansion 

for a future supersonic aircraft with under-expanded nozzle exhaust flow by modifying nozzle pressure or nozzle divergent section 

geometry. 

15. SUBJECT TERMS

Exhaust; Nozzle; Plume; Wing; Sonic boom 

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