Raymond Castner

Glenn Research Center, Cleveland, Ohio

Troy Lake

Wichita State University, Wichita, Kansas

Exhaust Plume Effects on Sonic Boom for a

Delta Wing and a Swept Wing-Body Model

NASA/TM—2012-217446

April 2012

AIAA–2012–1033
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Raymond Castner

Glenn Research Center, Cleveland, Ohio

Troy Lake

Wichita State University, Wichita, Kansas

Exhaust Plume Effects on Sonic Boom for a

Delta Wing and a Swept Wing-Body Model

NASA/TM—2012-217446

April 2012

AIAA–2012–1033

National Aeronautics and

Space Administration

Glenn Research Center

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NASA/TM—2012-217446 1 

Exhaust Plume Effects on Sonic Boom for a 

Delta Wing and a Swept Wing-Body Model 

Raymond Castner 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135

Troy Lake 

Wichita State University 

Wichita, Kansas 67260 

Abstract 

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

boom. Research has been performed on sonic boom reduction and has included the contribution of the 

exhaust nozzle plume. Plume effect on sonic boom has progressed from the study of isolated nozzles to a 

study with four exhaust plumes integrated with a wing-body vehicle. This report provides a baseline 

analysis of the generic wing-body vehicle to demonstrate the effect of the nozzle exhaust on the near-field 

pressure profile. Reductions occurred in the peak-to-peak magnitude of the pressure profile for a swept 

wing-body vehicle. The exhaust plumes also had a favorable effect as the nozzles were moved outward 

along the wing-span. 

Nomenclature 

AOA Angle of attack, degrees 

Β Nozzle boat-tail angle, degrees 

c Airfoil chord, in. 

D Nozzle diameter (outer), in. 

h Distance below vehicle, in. 

Lv Vehicle length, in. 

M∞ Free-stream Mach number 

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∞

ΔP/P (P – P∞)/ P∞

To Nozzle total temperature, R 

T∞ Free-stream total temperature, R 

t Airfoil thickness, in. 

x Distance along abscissa of pressure signature, in. 

y Distance from nozzle centerline, in. 
NASA/TM—2012-217446 2 

1.0 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 to negative pressure, followed by 

a rise to atmospheric pressure. When propagated to the ground, this pressure profile takes the shape of an Nwave. The two large pressure changes create a “double boom” effect. The impact of the sonic boom is so 

large that the FAA 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 14 CFR Part 36 stage 4. 

To create an aircraft with acceptable noise, programs such as the Quiet Spike (Ref. 3) and the Shaped 

Sonic Boom Demonstrator (SSBD) (Ref. 4) reduced the intensity of the nose shock wave. Most research has 

been focused on reducing the sonic boom contribution caused at the front of the vehicle, and less research has 

occurred in reducing the contribution from aft structures or nozzle exhaust. One exception, Putnam (Ref. 5), 

performed a study of exhaust nozzles and the effects of the exhaust plume. Tests were done in the 4- by 4-

Foot Supersonic Wind Tunnel with pressure measurements taken one diameter away from the nozzle. 

Exhaust nozzle plume effect on sonic boom has progressed from analysis and testing of an isolated 

nozzle (Refs. 6 and 7), to analysis of a slot nozzle embedded in a simplified wing geometry (Ref. 8), to 

analysis of a single engine wing-body model (Ref. 9), to the present analysis of a multi engine wing-body 

model. For the isolated nozzle, it was shown how the lip shock from an under-expanded nozzle plume could 

suppress the nozzle boat-tail expansion and reduce the trailing shock. Similar results were found for the slot 

nozzle embedded in the wing, where the trailing shock caused by the wing expansion was reduced by the 

addition of an under-expanded nozzle plume. Results for a single engine wing-body model also demonstrated 

a reduction in the trailing shock by including an under-expanded nozzle plume. 

The subject of this report is the preliminary study of a simplified wing-body configuration with the 

addition of four engine exhaust nozzle plumes. The purpose is to provide a baseline analysis of a generic 

wing-body configuration, and demonstrate the effect of the nozzle exhaust on the vehicle pressure profiles, 

which could be extrapolated to a sonic boom signature. The Wind-US and Cart3D computational fluid 

dynamic (CFD) codes were used for this analysis. Three types of nozzle plumes were studied: an 

axisymmetric convergent-divergent nozzle, an axisymmetric plug nozzle, and a high aspect ratio nozzle. Two 

vehicle configurations were studied: a 59 delta wing-body model and a 69 swept wing-body model. The 

pressure profiles from these configurations are presented and compared to a baseline vehicle with no 

propulsion to demonstrate the effect of the nozzle exhaust plume. 

2.0 Computational Modeling 

2.1 Wind-US 

Three types of exhaust nozzles were modeled with Wind-US: (1) a convergent-divergent supersonic 

nozzle replica of “Nozzle 6” (Ref. 5), (2) a plug nozzle, and (3) a high aspect ratio convergent-divergent 

supersonic slot nozzle. The configurations will be discussed in Section 3.0. Wind-US is a general purpose 

fluid flow solver that is used to numerically solve various sets of equations governing physical phenomena 

(Ref. 10). Wind-US was used to take advantage of the established capability to correctly compute nozzle 

plumes with viscous and turbulence effects. The code supports the solution of the Euler and Navier-Stokes 

equations, along with supporting equation sets governing turbulent and chemically-reacting flows. The flow 

solver is parallel and can take advantage of multi-core and multi-cpu hardware. The version used was WindUS 2.0. Wind-US was used with the modified second-order Roe upwind scheme for stretched grids, implicit 

time stepping with a Courant–Friedrichs–Lewy (CFL) number of 1.0, and the Menter shear stress transport 

(SST) turbulence model. 
NASA/TM—2012-217446 3 

2.2 Cart3D 

Two vehicle configurations were studied using Cart3D: (1) a 59 delta wing-body model and (2) a 

69 swept wing-body model. The vehicle configurations were fully three-dimensional models, and will be 

discussed in Section 3.0. Cart3D is a high-fidelity inviscid analysis package for conceptual and 

preliminary aerodynamic design. It allows users to perform automated CFD analysis on complex 

geometry. Geometry for Cart3D is in the form of surface triangulations. These may be generated from 

within a Computer-Aided Design (CAD) package, from legacy surface triangulations or from structured 

surface grids. Cart3D uses adaptively refined Cartesian grids to discretize the space surrounding a 

geometry and cuts the geometry out of the set of “cut-cells” which intersect the surface triangulation. The 

flow solver is parallel and can take advantage of multi-core and multi-cpu hardware. Solutions used the 

Cart3D adjoint adaptation module. This module uses adjoint-weighted residual error-estimates to drive 

mesh adaptation. Once a user specifies outputs of interest (lift, drag, etc.) with a corresponding error 

tolerance, this module automatically meshes the simulation to drive the remaining numerical errors in the 

outputs below the requested tolerance. This module has been validated for sonic boom prediction by 

Wintzer (Ref. 11), and this capability was the reason Cart3D was selected for analysis of this problem. 

The adaptation module allows greatly reduced time spent on mesh generation and analysis. 

3.0 Geometry Modeling 

3.1 Nozzles 

For the nozzle configurations, the computational domain was an axisymmetric structured grid which 

consisted of 8 zones, and 511,299 grid points. The slot nozzle simulation was a similar two-dimensional 

structured grid. The domain extended 18 nozzle diameters downstream of the nozzle exit. Multi-block wallpacked grids were generated for use on parallel processor systems. Viscous wall boundaries were used for 

all nozzle surfaces. Convergence was monitored with nozzle mass flow and the off-body pressure profile at 

150 in. from the nozzle centerline. External flow conditions were run at Mach 2.2 and an angle of attack of 

zero. Cases were run on a local computer cluster and took approximately 6 to 8 hr to converge. 

3.1.1 Convergent-Divergent Nozzle 

Figure 1 displays the Mach 2.0 supersonic convergent-divergent exhaust nozzle, which was a replica 

of the Putnam “Nozzle 6.” This nozzle had a design pressure ratio of 8.12 and simulations were 

performed from NPR = 6 to 18. For the wing-body model propulsion-integration analysis, described in 

Section 3.2, the simulation was performed at an under-expanded pressure ratio of 18, where a highly 

under-expanded plume was selected with the intent to demonstrate a significant plume effect. If 

successful, a NPR closer to the design point would be studied. The CFD grid is shown in Figure 3(a) and 

the nozzle plume shape in Figure 4(a). The shape of the nozzle plume was extracted from the simulation 

and imported into grid generation software, where the nozzle plume was modeled as a solid body. Use of 

solid body models of nozzle plumes was validated by both Putnam (Ref. 5) and Castner (Ref. 9). 

3.1.2 Plug Nozzle 

The plug nozzle was simplified to a single stream nozzle, and was a replica of the Chenoweth 

(Ref. 12) nozzle core stream only, shown in Figure 2. The design pressure ratio was 26.3 and simulations 

were performed at NPR = 8, 14, and 18. Again, the wing-body simulation was performed at a NPR = 18. 

The CFD grid is shown in Figure 3(b) and the resulting nozzle plume shape in Figure 4(b). For the wingbody model analysis, the plume was extracted from the simulation and imported into grid generation 

software and the nozzle plume was modeled as a solid body. 

 
NASA/TM—2012-217446 4 
NASA/TM—2012-217446 5 

3.1.3 Slot Nozzle 

A fully-expanded Mach 2.2 slot nozzle was designed, operating at a NPR = 12, for highly underexpanded flow. The NPR = 12 plume was selected due to the excessive size of the plume at NPR = 18. 

The nozzle was a simple convergent-divergent geometry with a throat height of 2.05 in. and an exit height 

of 3.88 in. There was a short external boat-tail of 5, and an external height of 4.57 in. The CFD grid is 

shown in Figure 3(c). In the wing-body analysis, the slot nozzle had an effect on a larger portion of the 

wing area, so a smaller plume shape at the reduced pressure ratio was selected. The nozzle plume shape 

(Fig. 4(c)) was extracted from the simulation and imported into grid generation software and the nozzle 

plume was modeled as a solid body. In the wing-body analysis the nozzle had an aspect ratio of 11.2:1. 

3.2 Vehicles 

Two vehicle configurations, from Hunton (Ref. 13), were studied (Fig. 5): a 59 delta wing-body 

model and a 69 swept wing-body model. Vehicles were studied with a notional 4-engine configuration. 

Figure 6 shows a comparison between the Concorde (Ref. 1) and the 59 delta wing-body model, with the 

scale factor estimated at 9:1. Table 1 shows comparisons between some key aircraft dimensions. 

For the Cart3D CFD analysis, the original wing-body vehicle scale was increased by a factor of 

14.11, where the vehicle body diameter was modeled as 15.1 in. This diameter allowed use of previous 

modeling work that was performed with the 59 delta wing-body vehicle and Putnam’s “Nozzle 6”. At 

this scale, the notional propulsion was 3.38 in. in diameter (instead of a very small 0.23 in. diameter). 

Dimensions of the nozzles are given in Table 2. Exit area and throat area were not held constant during 

this study, and the plug nozzles were nominally 24 percent larger in diameter than the convergentdivergent nozzle. A size for “Nozzle 6” was estimated from the Concorde propulsion system, assuming 

10,000 lb of thrust per engine at cruise conditions, and then geometrically reduced to the nominally 1/9 

scale difference between the Concorde and the model. When the same procedure was used for the plug 

nozzle, there was a difference in exit area. 
NASA/TM—2012-217446 6 

 
NASA/TM—2012-217446 7 

TABLE 1.—COMPARISONS BETWEEN THE CONCORDE AND THE 59 DELTA WING-BODY MODEL 

Concorde 59 delta wing-body Scale factor 

(Concorde: Delta wing-body dimensions) 

Length 2428 in. Length 247.5 in. 9.8 : 1 

Wing Span 446 in. Wing span 52.3 in. 8.5 : 1 

TABLE 2.—KEY NOZZLE DIMENSIONS 

Nozzle NPR T0 Exit area sq-in. Exit diameter, in. Outer diameter, in. 

Convergent-divergent 8, 18 530 R 2.66 3.38 3.75 

Plug 8, 18 530 R 4.9 3.52 4.64 

 

NPR T0 Exit area sq-in. Exit width, in. Exit height, in. 

Slot 12 530 R 2.8 5.6 0.5 

The model was run on the Pleiades computer cluster at the NASA Ames Research Center. The adjoint 

adaptation module uses Open MultiProcessing (OpenMP) for code parallelization within a node, and at 

the time of writing, could not utilize message passing interface (MPI) to work between nodes, required on 

Peliades architecture. As a result, a single node with 12 processors was used until the node ran out of 

memory, at approximately 4.5 million cells. 

3.2.4 59 Delta Wing-Body Model 

The 59 delta wing-body model was a simplified platform for a propulsion integration study on sonic 

boom. The wing-body model was modified with a long sting, tapered at the trailing end. The tapered sting 

was installed to eliminate the effects of a tail cone and tail, so propulsion effects could be studied with a 

focus on the nozzle plume. The overall length of the body and sting was 398.9 in., and the trailing portion 

of the sting had the same geometry as the nosecone. Engines were installed in an embedded configuration 

at the back of the wing as shown in Figure 7. Embedded nozzles were used to isolate plume effects, 

without the effects of an inlet flow field. The engine plumes were constructed of the actual exhaust plume 

shape, followed by a cylindrical solid body (rectangular solid body for the slot nozzle). The cylindrical 

portion was extended aft to a point where the termination of the plume would not affect the pressure 

signature from the nose, wing, and nozzles. The configurations tested with the 59 delta wing-body model 

are shown in Table 3, and included three different levels of engine placement (spacing and stagger). 

3.2.5 69 Swept Wing-Body Model 

The 69 swept wing-body model was used as an alternate platform for propulsion integration studies. 

The swept wing-body model was also modified with a long sting, tapered at the trailing end, so 

propulsion effects could be studied with a focus on the nozzle plume. Engines were installed at the back 

of the wing as shown in Figure 8. Again, embedded nozzles were used to isolate plume effects. For this 

configuration the engine locations were studied at three positions, locations are shown in Figure 9. The 

engine plumes were again constructed of the actual exhaust plume shape, followed by a cylindrical solid 

body. The cylindrical portion was extended aft to a point where the termination of the plume would not 

affect the pressure signature from the nose, wing, and nozzles. The configurations tested with the 69

swept wing-body model are also shown in Table 3. 
NASA/TM—2012-217446 8 
NASA/TM—2012-217446 9 

TABLE 3.—59 DELTA WING-BODY MODEL AND 69 SWEPT WING-BODY MODEL CONFIGURATIONS 

Configuration Vehicle 

type 

Nozzle 

type 

Inboard nozzle 

location, in. 

Nozzle 

spacing, 

nozzle 

diameters 

Nozzle 

stagger, 

nozzle 

diameters 

NPR Mach / T∞ ,R 

C430 Delta-wing C-D 13.17 3 0 18 2.2 / 530 

C413 Delta-wing C-D 13.17 1 3 18 2.2 / 530 

C410 Delta-wing C-D 13.17 1 0 18 2.2 / 530 

P430 Delta-wing Plug 13.17 3 0 18 2.2 / 530 

P413 Delta-wing Plug 13.17 1 3 18 2.2 / 530 

S451 Delta-wing Slot 12.64 1 0 12 2.2 / 530 

C411i Swept-wing C-D 9.5 1 1 8, 18 2.2 / 530 

C411m Swept-wing C-D 13.5 1 1 8, 18 2.2 / 530 

C411o Swept-wing C-D 17.5 1 1 8, 18 2.2 / 530 

P411i Swept-wing Plug 9.9 1 1 8, 18 2.2 / 530 

P411m Swept-wing Plug 13.9 1 1 8, 18 2.2 / 530 

P411o Swept-wing Plug 17.9 1 1 8, 18 2.2 / 530 

S451 Swept-wing Slot 12.64 1 0 12 2.2 / 530 

Fig. 1. Convergent-divergent "Nozzle 6 5". 

Fig. 2. Plug nozzle 12. 

4.0 Results 

Pressure profiles (ΔP/P) were computed at three body lengths (h = 3 Lv) below the vehicle. Three sets of 

results will be presented: (1) pressure profiles for the 59 delta wing-body model at a NPR = 18, (2) pressure 

profiles for the 69 swept wing-body model at a NPR = 18, and (3) pressure profiles for the 69 swept wingbody model at a NPR = 8. Propulsion was analyzed with different levels of engine placement, including 

spacing and stagger (i.e., one nozzle exit plane located behind the other). 

4.1 ∆P/P Pressure Profile for 59 Delta Wing-Body Model 

Pressure profiles are displayed in Figure 10 for the 59 delta wing-body model at a distance of three 

vehicle lengths (h = 3 Lv) below the vehicle. Major features in these pressure profiles include the nose shock, 

wing shock, wing expansion, and trailing shock back to ambient pressure. The baseline vehicle pressure 

signature is the 59 delta wing-body model with no propulsion installed. Configurations with propulsion 

installed all had a small reduction in the peak-to-peak magnitude of the pressure profile. It was difficult to 

differentiate been the pressure profiles for most nozzle installations. Only two configurations demonstrated a 

small improvement over the others: (a) the slot nozzle configuration and (b) the plug nozzle with a spacing of 3 

and no stagger. These results did not demonstrate a significant impact on the overall peak-to-peak magnitude 

of the pressure profile, and were only useful for a small improvement in the vehicle pressure profile. 

4.2 ∆P/P Pressure Profile for 69 Swept Wing-Body Model, NPR = 18 

Figure 11 shows the pressure profiles at a location three vehicle lengths below the vehicle for the swept 

wing-body model with convergent-divergent (C-D) nozzles operating at a NPR = 18. In this case, the 

pressure profile is characterized by a nose shock, a wing shock, the nozzle lip shock, and the trailing shock. 

For the C-D nozzles, all configurations showed a favorable reduction in the peak-to-peak value of the 

pressure profile, when compared to the baseline vehicle with no propulsion. As the engine nacelles were 

moved from their inboard, to middle, to outboard location along the wing-span, there was a reduction in the 

magnitude of the trailing shock, which would be favorable for a reduction in sonic boom. 

Configuration code = C411i 

 i = inner, m = middle, o = outer position 

 engine spacing: 1 or 3 engine diameters (or slot height 0.5 in.) 

 engine stagger: 0,1 or 3 engine diameters 

 4 engines 

 C = C-D nozzle, P = Plug nozzle, S = Slot nozzle 
NASA/TM—2012-217446 10 
NASA/TM—2012-217446 11 

Figure 12 shows the pressure profiles for the swept wing-body model with plug nozzles operating at 

NPR = 18, also at three vehicle lengths below the vehicle. When compared to the baseline vehicle with no 

propulsion, the peak-to-peak pressure profile was increased with the addition of plug nozzles located at 

the inboard location along the wing-span. As the engines were moved to the middle and outboard 

locations, the magnitude of the trailing shock decreased. For this study, the plug nozzles were 25 percent 

larger than the C-D nozzles and were potentially too large to have a beneficial effect on the pressure 

profile. This type of detrimental effect on sonic boom was previously demonstrated in studies of isolated 

nozzles where an underexpanded plume reduced the near-field pressure profile for moderate increases in 

NPR, but a large increase in NPR caused an increase in the peak-to-peak pressure profile (Ref. 6). Results 

for the slot nozzle case are not shown as it had no improvement over the C-D or the plug nozzle results. 

4.3 ∆P/P Pressure Profile for 69 Swept Wing-Body Model, NPR = 8 

Simulations from the previous section resulted in under-expanded nozzle flow. For the C-D nozzle, 

the design point was a NPR = 8. At this point, the thrust coefficient was computed to be 0.99, but at 

NPR = 18 the thrust coefficient was reduced to 0.91. To study the effect at the design point, simulations 

for the swept wing-body geometry were re-calculated using a NPR = 8. 

Figure 13 contains the pressure profiles for the C-D nozzles at NPR = 8. All configurations showed a 

favorable reduction in the peak-to-peak value of the pressure profile, when compared to the baseline 

vehicle with no propulsion. Again, as the engine nacelles were moved from their inboard, to middle, to 

outboard location along the wing-span, there was a reduction in the magnitude of the trailing shock. For 

the C-D nozzles operating at their design point, these profiles show the benefit that this embedded 

prolusion configuration has on a generic swept wing vehicle. 

Figure 14, shows results for the plug nozzles at NPR = 8 (on the swept wing-body vehicle). As 

before, when compared to the baseline vehicle with no propulsion, the peak-to-peak pressure profile was 

increased with the addition of plug nozzles located at the inboard location along the wing-span. As the 

engines were moved to the middle and outboard locations, the magnitude of the trailing shock again was 
NASA/TM—2012-217446 12 

decreased. These results suggest that the underexpanded plume shape generated by the plug nozzle at 

NPR = 18 was too large to be of benefit for this configuration, but operation at NPR = 8 showed 

reductions in the peak-to-peak value of the pressure profile as the engine locations were moved outward. 

This is again similar to the effect found in isolated nozzle research, where an underexpanded plume 

reduced the near-field pressure profile for moderate increases in NPR, but a large increase in NPR caused 

an increase in the peak-to-peak pressure profile. 
NASA/TM—2012-217446 13 

5.0 Conclusions 

Three types of nozzles and two types of wing-body vehicles were used to study propulsion integration 

for nozzle plume effects on sonic boom. The study was not intended to be an aircraft design, but an 

attempt to isolate nozzle plume effects on a simple integrated configuration. The vehicle configuration did 

not include a tail cone, vertical tail, or horizontal stabilizer. 

Results for the delta wing-body vehicle indicated that the three different types of nozzles had a small 

effect on the overall pressure profile, showing only 10 percent improvement in the trailing shock strength. 

The shock and expansion wave from the delta wing was stronger than the shock and expansion from the 

swept wing, and the nozzle lip shock was not able to affect the wing expansion as seen in past studies of 

isolated nozzles. 

Changes in the pressure profile were favorable for the swept wing-body configuration, where 

convergent-divergent nozzles produced a reduction in the trailing shock. The plug nozzles, when 

operating at NPR = 8, also produced a reduction in the trailing shock. The exhaust plumes reduced the 

peak-to-peak magnitude of the pressure profile as the nozzles were moved outward along the wing-span. 

Movement of the nozzles resulted in a 28 to a 44 percent reduction in the strength of the trailing shock, 

depending on nozzle pressure ratio. 

This preliminary study utilized a stepping stone approach to focus on how nozzle plumes interact with 

the wing-body vehicle. Nozzle plume effects for a four engine vehicle have a visible effect on the near 

field pressure profile. Based on the improvements in near field profiles, the following future analysis are 

feasible: (1) a design of experiments study to determine optimum nozzle placement and pressure ratio, 

and (2) a study which includes the addition of the vehicle tail-cone and tail. 

References 

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

[retrieved 24 February 2011]. 

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

apl/noise_emissions/ supersonic_aircraft_noise/ [retrieved 5 December 2011]. 

3. Freund, D., Howe, D., Simmons, F., and Schuester, L., “Quiet Spike Prototype Aerodynamic 

Characteristics From Flight Test,” AIAA−2008−125, Jan. 2005.

4. Graham, D., Dahlin, J., Meredith, K., and Vadnais, J., “Aerodynamic Design of Shaped Sonic 

Boom Demonstration Aircraft,” AIAA−2005−8, Jan. 2005. 

5. 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, 1969. 

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

Configurations,” NASA/TM—2008-215414 (AIAA−2008−3729), June 2008. 

7. Bui, T., “CFD Analysis of the Nozzle Jet Plume Effects on Sonic Boom Signature,” 

AIAA−2009−1054, Jan. 2009. 

8. Castner, R.S., “Slot Nozzle Effects for Reduced Sonic Boom on a Generic Supersonic Wing 

Section,” AIAA−2010−1386, Jan. 2010. 

9. Castner, R.S., “Analysis of Exhaust Plume Effects on Sonic Boom for a 59-Degree Wing Body 

Model,” AIAA−2011−917, Jan. 2011. 

10. Towne, C.E., “Wind-US Users Guide, Version 2.0,” NASA/TM—2009-215804, Oct. 2009. 

11. Wintzer, M., Nemec, M., and Aftosmis, M., “Adjoint-Based Adaptive Mesh Refinement for Sonic 

Boom Prediction,” AIAA−2008−6593, Aug. 2008. 

12. Chenoweth, F.C., and Lieberman, A., “Experimental Investigation of Heat-Transfer Characteristics 

of a Film-Cooled Plug Nozzle With Translating Shroud,” NASA TN D-6160, Feb. 1971. 

13. Hunton, L., Hicks, R., and Mendoza, J., “Some Effects of Wing Planform on Sonic Boom,” NASA 

TN-D 7160, Jan. 1973. 
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1. REPORT DATE (DD-MM-YYYY)

01-04-2012 

2. REPORT TYPE

Technical Memorandum 

3. DATES COVERED (From - To)

4. TITLE AND SUBTITLE

Exhaust Plume Effects on Sonic Boom for a Delta Wing and a Swept Wing-Body Model 

5a. CONTRACT NUMBER

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Castner, Raymond; Lake, Troy 

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

WBS 984754.02.07.03.13.02 

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-18176 

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-2012-217446 

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited 

Subject Categories: 02 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

Supersonic travel is not allowed over populated areas due to the disturbance caused by the sonic boom. Research has been performed on 

sonic boom reduction and has included the contribution of the exhaust nozzle plume. Plume effect on sonic boom has progressed from the 

study of isolated nozzles to a study with four exhaust plumes integrated with a wing-body vehicle. This report provides a baseline analysis 

of the generic wing-body vehicle to demonstrate the effect of the nozzle exhaust on the near-field pressure profile. Reductions occurred in 

the peak-to-peak magnitude of the pressure profile for a swept wing-body vehicle. The exhaust plumes also had a favorable effect as the 

nozzles were moved outward along the wing-span. 

15. SUBJECT TERMS

Sonic booms; Exhaust nozzle; Computational fluid dynamic (CFD); Exhaust gases; Plumes; Delta wings; Body-wing configurations 

16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF

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19 

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

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443-757-5802 

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