Rodrick V. Chima

Glenn Research Center, Cleveland, Ohio

Computational Analysis of a Low-Boom

Supersonic Inlet

NASA/TM—2011-217223

November 2011

AIAA–2011–3801
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Rodrick V. Chima

Glenn Research Center, Cleveland, Ohio

Computational Analysis of a Low-Boom

Supersonic Inlet

NASA/TM—2011-217223

November 2011

AIAA–2011–3801

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

29th Applied Aerodynamics Conference

sponsored by the American Institute of Aeronautics and Astronautics

Honolulu, Hawaii, June 27–30, 2011
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Available electronically at http://www.sti.nasa.gov

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

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

either expressed or implied, by the National Aeronautics and 

Space Administration.

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

This work was sponsored by the Fundamental Aeronautics Program 

at the NASA Glenn Research Center.
A low-boom supersonic inlet was designed for use on a conceptual small supersonic 

aircraft that would cruise with an over-wing Mach number of 1.7. The inlet was designed to 

minimize external overpressures, and used a novel bypass duct to divert the highest shock 

losses around the engine. The Wind-US CFD code was used to predict the effects of capture 

ratio, struts, bypass design, and angles of attack on inlet performance. The inlet was tested in 

the 8-ft by 6-ft Supersonic Wind Tunnel at NASA Glenn Research Center. Test results 

showed that the inlet had excellent performance, with capture ratios near one, a peak core 

total pressure recovery of 96 percent, and a stable operating range much larger than that of 

an engine. Predictions generally compared very well with the experimental data, and were 

used to help interpret some of the experimental results. 

I. Introduction 

The low-boom supersonic inlet (LBSI) was designed by engineers at Gulfstream Aerospace Corporation (GAC) 

for use on a small aircraft that would cruise at a Mach number of 1.6 at 45,000 feet1

. Two features of the inlet were 

designed to minimize external overpressures that would contribute to the sonic boom signature of the aircraft. First, 

the external cowl angle was kept small to minimize external shocks. Second, the inlet was designed to capture nearly 

100 percent of the oncoming supersonic flow, thereby minimizing external flow spillage. 

The inlet used an axisymmetric, relaxed isentropic compression centerbody design to decelerate the flow into a 

non-uniform terminal shock. The shock was weak at the centerbody to minimize shock/boundary-layer interaction, 

but it was stronger at the cowl. 

The inlet had two concentric flow paths. The inner, or core, flow path led to the aerodynamic interface plane 

(AIP) where the engine would be attached. The outer, or bypass, flow path was designed to route the low 

momentum air from the strongest part of the normal shock around an engine and gearbox, and back to the nozzle 

stream. This dual-stream design improved core recovery by diverting the high shock losses near the cowl through 

the bypass duct, and also minimized external shocks by keeping the gearbox out of the freestream flow2

. 

The LBSI was tested in the 8-ft by 6-ft Supersonic Wind Tunnel (SWT) at NASA Glenn Research Center (GRC) 

in the fall of 20103

. The tests were conducted by a team of researchers from GRC, GAC, the University of Illinois at 

Urbana-Champaign, and the University of Virginia. Two inlet designs were tested – the dual-stream design intended 

for an aircraft, and a simpler single-stream design intended for more detailed flow physics measurements and CFD 

validation. The single-stream inlet will be reported in other publications and is not discussed here. Both inlets were 

tested with and without flow control devices, including microramps and vane vortex generators on the centerbody. 

The flow control devices will also be reported elsewhere. 

Research team members performed many CFD analyses of the two inlets before the test was started. The 

analyses were done with several different CFD codes and at different levels of fidelity, including: 

• Design of the compression spike using a method of characteristics code1

, 

• Axisymmetric analyses to refine the design and predict general performance trends4

, 

• An analysis of the dual-stream inlet coupled to a Rolls-Royce fan to predict engine stability characteristics5

, 

• 3-D analyses of the inlets including struts and bypass geometry (present work), and 

• Analyses of the inlets with microramps and vane vortex generators 6, 7, 8. 

This paper describes a 3-D CFD analysis of the dual-stream LBSI without flow control devices, and compares 

the results of the analysis with data from a wind tunnel test. The inlet and test are described briefly, and the CFD 

analysis is described in more detail. The analysis was performed before the test to estimate the performance of the 

inlet, and to investigate the effects of struts, bypass vanes, and angle of attack on performance. Here the results of 

Rodrick V. Chima 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

NASA/TM—2011-217223 1

Computational Analysis of a Low-Boom Supersonic Inlet 
the analysis are compared to the experimental data, primarily to validate the predictions, and also to explain some of 

the experimental results that are not obvious from the data. 

II. Dual-Stream Low-Boom 

Supersonic Inlet 

Engineers at Gulfstream Aerospace 

Corporation have been investigating 

technologies that would enable the use of a 

small, low-boom supersonic aircraft. Their 

reference aircraft, (Ref. 1) was designed to 

cruise at 45,000 feet at a Mach number of 

1.6, with an over-wing Mach number of 1.7. 

The aircraft would use radical aerodynamic 

shaping and a Quiet Spike9

 to minimize 

sonic boom. The inlets make up much of the 

frontal area of the aircraft, so they were 

designed carefully to minimize their 

contribution to the overall sonic boom 

characteristics. Axisymmetric, externalcompression inlets were chosen for 

simplicity. 

Reference 1 also describes an innovative 

approach for designing the external 

compression surface of the inlet that gives 

improved recovery and reduced sonic boom overpressures over conventional designs. Inlets designed with this 

approach have been tested experimentally10 at small scale and modeled computationally5, 11, and shown to have good 

recovery and stability characteristics. 

In the present study a dual-stream inlet was designed using the approach described in Ref. 1. It was designed to 

reduce the Mach number from 1.7 over the wing to about 0.65 at the fan face, with high total pressure recovery, 

minimal distortion, and with minimal external overpressures. The inlet, shown in Figure 1, consists of a centerbody, 

a core flow stream, and a concentric bypass flow stream. Five thin struts connect the centerbody, bypass splitter, and 

cowl. 

The centerbody consists of an isentropic compression spike, a curved throat region, and a subsonic diffuser (not 

visible in the figure). The compression spike was designed to produce a variable-strength normal shock at the throat. 

The Mach number ahead of the shock is about 1.3 near the centerbody, which is low enough to avoid boundary-layer 

separation. The Mach number at the cowl is close to the design over-wing value of 1.7, which generates a strong 

normal shock with high total pressure loss. After the shock the Mach number is reduced through the subsonic 

diffuser to 0.65 at the AIP. 

The bypass duct was used to capture the high-loss flow near the cowl and divert it around the engine and back to 

the nozzle stream. This removes the high-loss flow from the core stream, and maximizes the total pressure recovery 

at the AIP. One of the goals of the present study was to demonstrate the performance and stability of an inlet with 

this type of a bypass duct. 

In the test the bypass flow was throttled using five choked exit plates with annular slots that returned the bypass 

flow to the freestream. In an aircraft application the bypass duct would extend to the nozzle, and convergingdiverging vanes would be used to expand the flow back to supersonic speeds. 

The bypass duct can be used to keep the engine gearbox out of the external flow, where it would contribute to 

boom. The inlet considered here was sized for a Rolls-Royce Tay engine, which has a gearbox that extends almost 

160 degrees around the perimeter. The bypass duct used 10 curved vanes to route the flow around the gearbox area. 

To maintain subsonic flow in the bypass duct it was necessary to increase the flow area downstream of the lip, 

which led to an eight-degree external cowl angle. 

The bypass duct can also be used to divert separated flow from the sharp lip away from the core flow. Separation 

is a common problem with sharp-lipped supersonic inlets at low subsonic speeds and at angles of attack or yaw. In 

Ref. 5 it was shown that lip separation at low takeoff speeds could be contained in the bypass duct by blocking the 

duct exit. The present CFD work has shown that separation caused by high pitch or yaw angles is also diverted away 

from the core. 

Figure 1. Dual stream inlet cutaway. 

NASA/TM—2011-217223 2
III. 8-ft by 6-ft Supersonic Wind Tunnel Test 

Figure 2. Dual-stream inlet model and support hardware. 

The inlets were tested in the 8-ft by 6-ft SWT at NASA Glenn Research Center12. The test section of the tunnel 

is 8 feet high by 6 feet wide, and 23.5 feet long. The tunnel walls are perforated and surrounded by an evacuated 

balance chamber to remove the wall boundary layers. Large 26.5-inch diameter windows provided optical access to 

the test section, and a schlieren system was used to visualize the flow around the inlet. A high-speed Phantom V310 

camera operating at 20004000 frames per second recorded steady or unsteady schlieren images. 

For this test the tunnel was run in a closed-loop cycle, with the air passing through a dryer and a cooler to permit 

continuous operation. Data was taken at Mach numbers of 0.5 and 1.4 – 1.8. The SWT is an atmospheric tunnel that 

operates at high dynamic pressure and temperature. At M = 1.8 the tunnel total pressure is 23.2 psi and the total 

temperature is 631 R. The models and instrumentation were designed to handle these conditions. 

Gulfstream engineers performed the initial aerodynamic and mechanical design of the model, and TriModels, 

Inc. in Huntington Beach, CA did the detailed design and fabrication. The model was instrumented with a total of 

241 static or total pressure taps on the centerbody, cowl, two boundary-layer rakes, and eight AIP rakes. Standard 

SAE 1420 total pressure rakes13 were located at the AIP, with eight five-probe rakes. A sixth probe was added to 

each rake near the hub to better resolve the hub boundary layers, which were expected to be large. 

The dual-stream inlet model and the wind tunnel support hardware are shown in Figure 2. The model was 

mounted on a hydraulic strut that could be rotated to vary the angle of attack from -2 to +5 degrees. Yaw angles 

could not be varied. 

The inlet core flow expanded through a subsonic diffuser into a 16-inch diameter cold pipe. The core flow was 

throttled using a hydraulically actuated conical mass flow plug (MFP) at the exit of the cold pipe. The MFP was 

calibrated to measure the core-stream mass flow rate to within a few tenths of a percent. The bypass channels were 

throttled using interchangeable choke plates at the channel exit. CFD was used before the test4

 to estimate the choke 

plate area needed to give the desired mass flow ratio of m bypass / m core = 0.7 at full capture. The choke plate area was 

verified by exchanging plates early in the test. 

The dual-stream inlet had total pressure rakes at the exit of each bypass channel (Figure 1), with five probes in 

each rake. The mass flow through each channel was estimated using the average total pressure, the tunnel total 

temperature, and by assuming that M = 1.0 at the exit plate. 

Inlet capture ratio compares the actual mass flow through the bypass and core to the maximum mass flow that 

the inlet could capture. Capture ratio is defined by 

capture ratio = (m core + m bypass )/(V rcowl

2 )

The capture ratio was expected to be 1.0 at M = 1.8 when the MFP was wide open, but the measurements 

indicated a capture ratio of 1.145. The core mass flow was assumed to be accurate, so the bypass mass flow was 

reduced using a discharge coefficient CD = 1 / 1.145 = 0.873 . This CD corrected the overall capture ratio to 1.0 at 

this condition, and the same CD was used for all other operating conditions. Since CD might vary with operating 

conditions, the accuracy of bypass flow rate measurement is uncertain. 

NASA/TM—2011-217223 3
IV. Computational Model 

A. Differences Between Wind Tunnel Model and Flight (CFD) Inlet Designs 

The computations shown later were made for flight hardware at altitude conditions, while the inlet that was 

tested in the SWT had differences in scale, flow conditions, strut chord, and bypass geometry. These differences, 

summarized in Table 1 and discussed below, were expected to have only minor effects on the comparison between 

the measured and computed results. 

Table 1. Differences between wind tunnel model and flight (CFD) designs. 

Parameter SWT test Flight (CFD) Effects 

Diameter 12 in. (1/4.86 scale) 58.4 in. None 

Mach number 1.664 nominal 1.7 Slight differences in shock angles 

ReD 5.2  106 12  106 Minimal 

Strut size 4.75 in. chord at hub 

(scale size) 

7.45 in. chord at hub Differences in centerbody pressures and 

recovery behind struts 

Bypass exit 

geometry 

5 channels, diverging 

radius 

1 channel, constant 

radius 

Possible differences in secondary flow and 

losses 

The CFD was performed for a 58.4-inch diameter inlet at 45,000 feet and M = 1.7 . These conditions give a 

Reynolds number based on diameter of ReD = 5.2  106 . The test was performed on a 12-inch diameter (1 / 4.86

scale) model, in an atmospheric wind tunnel at M = 1.664 and ReD = 12  106 . Differences in freestream Mach 

number were accounted for by comparing pressure ratios or coefficients. Differences in Reynolds number were 

expected to be insignificant, since a sublimation technique used in the experiment showed that the centerbody 

boundary layer became fully turbulent very close to the leading edge. 

Five swept struts were used to hold the centerbody in both the wind tunnel model and flight designs. The struts 

were tapered from hub to tip. The scaled leading edge locations of the two designs were identical, but the chord of 

the wind tunnel model was increased by 2.7 inches (13 inches flight scale) uniformly along the span. Thus, the flight 

strut had a scale hub chord of 4.75 inches and the wind tunnel model strut had a hub chord of 7.45 inches. This 

design change was unintended and was not discovered until after the test. The increase in chord length decreases the 

recovery measured downstream of the struts in the wind tunnel model. However, only one AIP rake was located 

directly behind a strut, so the effects on overall recovery are expected to be small. The struts had NACA profiles 

with t / c = 0.04 , so the wind tunnel model had relatively more blockage between the struts. This increase in 

blockage causes centerbody pressures between the struts measured in the wind tunnel to be lower than pressures 

predicted by the CFD. 

The flight inlet and wind tunnel model had identical bypass vanes near the leading edge. In the wind tunnel 

model (Figure 1) the vanes were thickened away from the leading edge, and every other vane was extended to the 

end of the bypass duct for structural reasons. Additionally, the hub and tip radii were increased in the plenum region 

to provide a favorable pressure gradient before the choke plates. In the flight inlet the bypass vanes stop at the end of 

the gearbox fairing, and the flows from the individual channels merge in a common, constant-area plenum region 

(compare Figures 2 and 4). These differences in bypass duct geometry probably create differences in secondary 

flows and pressure distributions in the plenum region, but they are not expected to affect the mass flow or recovery 

characteristics of the bypass duct. 

NASA/TM—2011-217223 4
B. Computational Grids 

Computational grids for the inlet 

were generated using a combination of 

codes. Axisymmetric (x,r) grids for the 

core, bypass, and external flow were 

generated using Pointwise14. The core 

and external flow grids were rotated 

through 360 degrees using Pointwise, 

and the bypass grid was sheared 

between individual vanes using a 

custom code. The full, 360-degree inlet 

was modeled to allow calculations of 

yaw conditions. C-grids around the 

struts were generated using TCGRID15. 

The grid blocks were assembled using 

Pointwise, and converted to Wind-US16

format using Gridgen14. Boundary 

conditions for Wind-US were applied 

using Gridgen and GMAN, one of the 

Wind-US utilities. 

The computational grid is shown in 

Figures 3 – 5, which show the grid on 

the cowl, bypass vanes, and centerbody 

and struts, respectively. The final grid 

had about 24 million points in 25 

blocks, as given in Table 2. The spacing 

at the walls was 1 105

inches, giving 

y

+

= 1to 2 at the first point off the 

walls. Leading edges of the cowl, 

splitter, struts, and vanes were modeled 

as 2:1 ellipses, with 6-10 points along 

each surface, to give adequate 

resolution of bow shocks or incidence 

effects. 

Grid refinement studies were done 

on the (x,r) grids in previous, 

unpublished work. The grids described 

in Table 2 were regenerated with about 

half as many points in each direction, 

while maintaining spacing at the walls. 

Axisymmetric calculations showed that 

recoveries predicted on the coarse grid 

were within 0.13 percent of the fine 

grid results, so the fine (x,r) grid used 

here was assumed to be adequate. 

Tangential grids had the same spacing 

at the wall and roughly square cells 

near passage centers. Tangential grid 

resolution was tested in Wind-US by running solutions with sequencing, in which every other grid point was used. 

Results with sequencing showed little difference between the coarse and fine grid solutions. 

Figure 3. Computational grid on the cowl. 

Figure 4. Computational grids on the bypass vanes. 

Figure 5. Computational grids on the centerbody and struts. 

NASA/TM—2011-217223 5
Table 2. Computational grid sizes for the inlet. 

Region Blocks i j k Points 

Core 2 279 121 65 4,388,670 

Bypass 2 274 121 45 2,983,860 

External Flow 4 314 121 65 9,878,440 

Struts 5 181 45 65 2,647,125 

Core Exit 1 17 441 65 487,305 

Bypass Channels 10 193 33 45 2,866,050 

Bypass Exit 1 51 321 45 736,695 

Totals 25 23,988,145 

C. CFD Solution Scheme 

The dual-stream inlet was analyzed using the Wind-US code16. The Reynolds-averaged Navier-Stokes (RANS) 

equations were discretized using the Harten /Lax/Van Leer/Contact (HLLC) scheme with a minmod limiter and the 

SST turbulence model. The equations were solved using an alternating-direction-implicit (ADI) time-marching 

scheme with a Courant number of 2.0. 

Boundary conditions were specified as follows: 

• Upstream conditions were supersonic inflow with M = 1.7, and pitch and yaw angles specified. 

• Freestream conditions were specified on the outer boundaries. 

• The bypass exit was choked to the freestream pressure using a converging-diverging bump on the upper wall. 

• The core stream exit pressure was specified for each case to vary the inlet capture ratio. Previous calculations 

of a coupled inlet and fan showed that the static pressure at the AIP was indeed nearly constant5

. 

The solution was initialized to M = 0.6 and run a few hundred iterations to establish subsonic flow in the 

diffuser. Then the freestream conditions were reset to M = 1.7, and all the external flow blocks were reinitialized. 

The solution was run 10,000 iterations, which converged the capture ratio and recovery to plotting accuracy. 

Subsequent cases were restarted from previous solutions, usually at a higher capture ratio, and run 7,500 – 

10,000 iterations to convergence. All calculations were run on a cluster of 11 CPUs running at 3.2 GHz, and took 

2433 hours per case. 

V. Comparison of CFD and Test Results 

A. Design Point Performance 

This section compares the CFD predictions of the dual-stream inlet with the experimental data at the design 

point, M = 1.7, = 0o. First the CFD results are used to describe the overall flow, and then detailed comparisons are 

made of centerbody pressure 

distributions, recovery characteristics, 

and AIP profiles. 

Computed Mach number contours 

along the symmetry plane of the inlet 

are compared with an overlaid schlieren 

image in Figure 6. Note that the 

schlieren optics resolves the effects of 

the strong lip shock around the entire 

axisymmetric inlet. This results in an 

apparent radial shock between the lip 

and centerbody (at the cut off edge of 

the schlieren image), while the actual 

shock curves into the inlet. For this 

figure the core exit pressure was low, 

i.e., the MFP was wide open. The 

schlieren image shows that the cowl 

captures most of the weak compression Figure 6. Schlieren image and computed Mach number contours on 

the inlet symmetry plane, M = 1.7, = 0°. 

NASA/TM—2011-217223 6
waves leaving the spike, confirming that the inlet is operating near its full-capture design point. The eight-degree 

cowl angle generates weak oblique shock waves on the exterior of the inlet, and the normal shock curves far into the 

inlet. The centerbody boundary layer does not separate, but it does become large (about 20 percent span) at the AIP 

due to the long subsonic diffuser. The bypass duct is choked to the freestream static pressure at the exit at all 

supersonic operating points. 

Pressure distributions along 

the centerbody are shown in 

Figure 7. The CFD solution is at 

a Mach number of 1.7, but the 

data was taken at a tunnel Mach 

number of 1.674. The difference 

in Mach numbers was accounted 

for by plotting pressure 

coefficient 

Cp = ( ) p  p/( ) p0 p. 

The two results are shown at the 

same capture ratio of about 0.97, 

indicated by the black diamonds 

on Figure 8. The CFD and data 

are in excellent agreement up to 

the start of the struts at x = 13 

inches. As discussed in Section 

IV A, the wind tunnel model had 

longer and thicker struts than the 

flight design analyzed here. The 

thicker struts in the wind tunnel 

model leave less area for the 

flow, and consequently increase 

the velocity and decrease the 

pressure below the computed 

values. Aft of the struts the 

measurements and CFD are 

again in good agreement. 

Overall core total pressure 

recovery is shown as a function 

of capture ratio in Figure 8. 

Capture ratio was varied 

experimentally by moving the 

MFP in small increments. The 

capture ratio was maximum 

when the MFP was wide open 

and decreased as the MFP was 

closed, until the inlet went into 

buzz at a capture ratio below 0.7 (not shown). AIP total pressures were measured using a standard SAE rake13 with 

an additional probe near the hub. Overall recovery was calculated by area-averaging the measured total pressures. 

The peak recovery was 96 percent, which is 10 percent higher than the normal shock recovery of 0.856 at M = 1.7. 

Computations were made at six capture ratios that covered the nominal operating range of the engine on an 

aircraft. The capture ratio was varied computationally by varying the static pressure at the exit of the core grid. No 

attempt was made to find the buzz point computationally. The maximum computed capture ratio is slightly higher 

than the measured value, but, as mentioned previously, the accuracy of the bypass mass flow is uncertain. Computed 

recoveries were evaluated at the same rake locations, and averaged using the same area-average as the 

measurements. Computed recoveries are 0.3 – 0.5 percent less than the measured values. 

Figure 7. Pressure distribution along the centerbody, M = 1.7,  = 0°. 

 

Figure 8. Core recovery vs. capture ratio, M = 1.7,  = 0°. 

NASA/TM—2011-217223 7
Overall bypass recovery is 

shown as a function of capture 

ratio in Figure 9. The computed 

recoveries were evaluated at the 

bypass rake probe locations, and 

overall are 1–1.5 percent lower 

than the measurements. The 

discrepancy could be due in part 

to differences between the 

tested and modeled geometry 

discussed earlier, but a more 

likely cause will be discussed 

below. For  = 0° the bypass 

recovery decreases almost 

linearly with capture ratio. Quite 

different behavior was seen at 

= 5°, which will be discussed in 

Section B, Effects of Angle of 

Attack. 

Figure 10. Total pressure recovery contours at the AIP, and profiles at rake locations, M = 1.7,  = 0°. 

A comparison of rake profiles reveals the reasons for the discrepancy in bypass recovery noted above. Both the 

bypass and core rakes were numbered counterclockwise from top dead center, as shown in Figure 10. Computed 

total pressure recovery contours in that figure show uniform core flow with a thick hub boundary layer and a thin tip 

boundary layer. This is also seen in the core rake profiles at the bottom of the figure, where symmetric left and right 

rakes are overlaid. The computed profiles are shown at the same tangential locations as the rakes. The computed 

profiles are perfectly symmetric, and the computed recoveries are generally one percent less than the measured 

values. 

The plot at the bottom left compares the top and bottom core rakes 1 and 5. Rake 1 is at the top, where the 

predicted recovery is again about one percent less than was measured. Rake 5 is at the bottom and sits directly in the 

wake of a centerbody strut. Its measured recovery is 5-6 percent lower than rake 1. In this one location the CFD 

predicted a higher recovery than was measured. As discussed in Section IV A, the wind tunnel model had longer 

struts than the flight design modeled with CFD, and would be expected to have a lower recovery than was predicted. 

Figure 9. Bypass recovery vs. capture ratio, M = 1.7,  = 0°. 

NASA/TM—2011-217223 8
Bypass rake profiles shown along the top of Figure 10 do not agree as well as the core profiles. The bypass rakes 

all lie directly downstream of long bypass vanes that develop thick wakes. The contour plot shows that bypass rake 

1 at the top lies exactly in the center of a wake, and the computed rake recoveries shown at top left are 4-5 percent 

lower than the data. If the CFD profiles are plotted just outside the wake the agreement with the data is much better, 

suggesting that the CFD model is not mixing the wake sufficiently with the freestream flow. Wakes from the other 

bypass vanes cross rakes 2 – 5 at an angle, and the agreement between the CFD and the data varies with radius. In 

general the computed recoveries are lower than the data, which explains the discrepancy in overall bypass recovery 

discussed above. 

B. Effects of Angle of Attack 

The dual-stream inlet was tested at 

angles of attack between -2° and +5°. Here 

CFD predictions are compared with the 

experimental data at  = 5°, an extreme 

condition that an aircraft would be very 

unlikely to encounter. First the CFD results 

are used to describe the overall flow, and 

then detailed comparisons are made of 

recovery characteristics and AIP profiles. 

Computed Mach number contours along 

the symmetry plane of the inlet are compared 

with an overlaid schlieren image in Figure 

11. For this figure the core exit pressure was 

high, i.e., the MFP was partially closed. The 

weak compression waves from the bottom of 

the spike are captured inside the cowl, but 

the waves from the top of the spike spill 

outside the cowl, confirming that the inlet is 

operating at a lower capture ratio. The 

asymmetric bow shock causes the 

centerbody boundary layer to separate on 

top. This is apparent in both the schlieren 

image and the computations. 

Computed total pressure recovery 

contours at the AIP in Figure 12 show the 

asymmetric wave pattern around the inlet 

and the thick boundary layer at the top of the 

centerbody. Thus, operation at angle of 

attack introduces circumferential distortion 

at the fan face; however, the distortion is 

mostly confined to the hub, while modern 

fans tend to be more sensitive to tip 

distortion. 

Figure 11. Schlieren image and computed Mach number 

contours on the inlet symmetry plane, M = 1.7,  = 5°. 

Figure 12. Total pressure recovery contours at the AIP, 

M = 1.7,  = 5°. 

NASA/TM—2011-217223 9
Figure 13. Core recovery vs. capture ratio, M = 1.7, = 5°. 

The overall core total pressure recovery plot shown in Figure 13 is similar to the recovery plot at zero angle of 

attack shown in Figure 8, but with two significant differences. First, the maximum measured capture ratio has 

decreased to 96.7 percent due to the spillage over the top of the inlet as noted previously. Second, the measured buzz 

point has increased from 0.60.7 to 0.790.89. This is close to the engine operating range, but again, = 5° is an 

extreme condition. The reason for the decrease in stable operating range is unknown. The computations are within 

0.5 percent of the measurements, and they were completely stable at the lowest capture ratio computed. 

Figure 14. Bypass recovery vs. capture ratio, M = 1.7, = 5°. 

Bypass recoveries for = 5° shown in Figure 14 behave quite differently than those for Figure 9. At = 0° the 

bypass recovery decreases almost linearly with capture ratio, but at = 5° the measured recovery has a maximum at 

a capture ratio around 0.90, and the CFD predictions have a maximum at a capture ratio of 0.975. The following 

discussion examines the CFD predictions to explain the cause of the discrepancy. 

NASA/TM—2011-217223 10
Figure 15. Computed streamlines and recovery contours in the bottom bypass channels, M = 1.7,  = 5°, high 

capture ratio. 

At  = 5° and high capture ratios the CFD solutions showed that the lower (windward) cowl lip was separated, 

and that the separated flow was confined to the two lower bypass channels. This is illustrated in Figure 15, which 

shows particle traces and recovery contours in the bottom bypass channels. Particle traces released near the 

separated cowl lip make one complete revolution around the channel and pass through the lowest recovery regions at 

each channel cross section. The large lip separation and the strong secondary flows are both challenging flow 

phenomena for any turbulence model to predict. 

Figure 16. Cowl lip recovery contours and bypass rake profiles at two capture ratios, M = 1.7,  = 5°. 

NASA/TM—2011-217223 11
Details of the lip separation at the start of the channel and the bypass recovery at the end of the channel are 

shown in Figure 16 at  = 5° for both low and high capture ratios. Recovery contours on the center plane near the 

lower lip are shown at the top, and measured and computed bypass recoveries from the bypass rakes are shown at 

the bottom. The recovery plot from Figure 14 is repeated in the center for reference. 

At low capture ratios (left), the recovery contours show that the bow shock is pushed outside of the inlet, and the 

flow is attached to the lip. The measured rake profile at the exit of the corresponding bypass channel shows 

recoveries between 0.8 and 0.86, and the CFD agrees reasonably well. At high capture ratios (right), the recovery 

contours show that the bow shock is pulled into the inlet, causing the flow to separate from the lip. The measured 

rake profile at the exit of the corresponding bypass channel shows much lower recoveries between 0.7 and 0.8, 

suggesting possible separation. Thus, for  = 5° these results suggest the following conclusions: 

• At low capture ratios the flow is attached to the lower cowl lip, and the bypass recovery in the bottom channels 

is similar to the recovery at  = 0°, about 84 percent overall. Here the CFD agreement is similar to the results 

at  = 0°. 

• At high capture ratios the flow separates at the lower cowl lip, and the bypass recovery in the bottom channels 

is much lower, about 74 percent overall. Here the CFD over predicts the recovery in the bottom channels, 

probably because of deficiencies in the turbulence model. 

Even though the lower cowl lip separates at high flow and high angle of attack, the CFD results show that the 

separated flow is captured in the bypass duct, leaving undistorted, high-recovery flow for the engine. 

C. Effects of Yaw Angle 

The dual-stream inlet was not tested at 

yaw, but CFD calculations were made to 

investigate these effects. In this section 

predictions are shown for = 5°, which is 

also an extreme condition for a 

commercial aircraft. 

Computed symmetry plane Mach 

number contours for M = 1.7, = 5° are 

shown in a top view in Figure 17. As with 

angle of attack, compression waves are 

captured on the windward side of the spike 

and spilled on the leeward side. Thus the 

maximum capture ratio also decreases with 

yaw. The case shown is for a high capture 

ratio, and, as at angle of attack, the cowl 

lip separates and the separated flow is 

captured in the bypass duct. On the 

windward side the bow shock is pulled far 

into the core and causes the centerbody 

flow to separate. The core flow 

reaccelerates to low supersonic speeds, and 

then shocks back down to subsonic. 

Total pressure recovery contours at the 

AIP are shown in Figure 18. Bypass 

channels on the left of the figure have 

clean flow, but bypass channels on the 

right show the results of the lip separation. 

The core flow shows an unusual 

circumferential distortion pattern at the 

hub. The effects of this type of distortion 

on engine operation require further study. 

Figure 17. Computed Mach number contours on the inlet 

symmetry plane, top view, M = 1.7, = 5°. 

Figure 18. Total pressure recovery contours at the AIP, 

M = 1.7, = 5°. 

NASA/TM—2011-217223 12
VI. Summary and Conclusions 

CFD predictions were made of a low-boom dual-stream inlet, and compared with experimental data. The inlet 

was designed by engineers at Gulfstream Aerospace Corporation for use on a conceptual small supersonic aircraft, 

and was tested in the 8x6 foot Supersonic Wind Tunnel at NASA Glenn Research Center. 

Test results showed that the dual-stream inlet had excellent performance, with capture ratios near one and peak 

core total pressure recoveries of 96 percent. The bypass recovery was much lower, between 84 and 87 percent. 

Gulfstream engineers feel that bypass losses are more than compensated for by gains in boom signature, core 

recovery, and engine operability. The inlet operated stably over a range of capture ratios much larger than the engine 

operating range. The peak core recovery remained constant with angle of attack, but the stable operating range 

decreased. 

CFD predictions were made with the Wind-US CFD code, on a grid with nearly 24 million grid points. The 

HLLC differencing scheme and Menter SST turbulence model were used. Predictions were compared with the 

experimental data, and the following results were found: 

• Computed shock positions agreed well with schlieren images. 

• Computed centerline pressures agreed well with measurements, except between the struts where differences 

between the tested and computed geometries caused differences in pressure distributions. 

• CFD calculations gave a slightly higher maximum capture ratio than was measured. However, the accuracy of 

the measured bypass flow is uncertain. 

• Predicted core recoveries were 0.30.5 percent lower than measured recoveries. 

• Predicted bypass recoveries were 11.5 percent lower than measured recoveries, probably due to insufficient 

mixing of the bypass vane wakes. 

• Total pressure profiles at the AIP agreed very well with the measurements in the core, but tended to be low in 

the bypass duct. 

• CFD predicted separation at the cowl lip at high capture ratios and high angles of attack or yaw. The separation 

is consistent with lower total pressures measured at bypass rakes far downstream. The separation is captured in 

the bypass ducts and would not affect the operation of an installed engine. 

CFD was used heavily to design the dual stream inlet. The good agreement between the CFD predictions and the 

experimental data shows that minor geometric differences between the flight inlet and the model used in the wind 

tunnel tests had negligible effects on inlet performance, and validates the use of CFD for future designs. 

References 

1. Conners, T. R., and Howe, D. C., “Supersonic Inlet Shaping for Dramatic Reductions in Drag and Sonic Boom Strength,” 

AIAA Paper 2006-30, Jan. 2006. 

2. Conners, T. R. and Wayman, T. R., “The Feasibility of a High-Flow Nacelle Bypass for Low Sonic Boom Propulsion 

System Design,” AIAA 29th Applied Aerodynamics Meeting, June 27-30, 2011 (to be published.) 

3. Hirt, S. M., Vyas, M. A., Chima, R. V., Reger, R. W., and Wayman, T. R., "Experimental Investigation of a Large-Scale 

Low- Boom Inlet Concept," AIAA 29th Applied Aerodynamics Conference, June 27-30, 2011 (to be published.) 

4. Chima, R. V., Hirt, S. M., and Reger, R., “Axisymmetric Calculations of a Low-Boom Inlet in a Supersonic Wind Tunnel,” 

AIAA 29th Applied Aerodynamics Meeting, June 27-30, 2011 (to be published.) 

5. Chima, R. V., Conners, T. R., and Wayman, T. R., “Coupled Analysis of an Inlet and Fan for a Quiet Supersonic Jet,” AIAA 

Paper 2010-479, Jan. 2010. Also NASA TM-2010-216350. 

6. Rybalko, M., Loth, E., Chima, R. V., Hirt, S. M., and DeBonis, J. R., “Micro-ramps for External Compression Low-Boom 

Inlets,” AIAA Paper 2009-4206, June 2009. 

7. Rybalko, M. and Loth, E., “Vortex Generators for a Single-Stream Low-Boom Inlet,” AIAA 29th Applied Aerodynamics 

Conference, June 27-30, 2011 (to be published.) 

8. Gillen, T., and Loth, E., “Vortex Generators for a Dual-Stream Low-Boom Inlet,” AIAA 29th Applied Aerodynamics 

Meeting, June 27-30, 2011 (to be published.) 

9. Wilson, J. R., “Quiet Spike Softening the Sonic Boom,” Aerospace America, Oct. 2007, pp. 38- 42. 

10. Conners, T. R., Merret, J. M., Howe, D. C., Tacina, K. M., and Hirt, S. M., “Wind Tunnel Testing of an Axisymmetric 

Relaxed External Compression Inlet at Mach 1.97 Design Speed,” AIAA Paper 2007-5066, July, 2007. 

11. Hirt, S. M., Tacina, K. M., Conners, T. R., Merret, J. M., and Howe, D. C., “CFD Results for an Axisymmetric Isentropic 

Relaxed Compression Inlet,” AIAA Paper 2008-0092, Nov. 2008. Also NASA TM-2008-215416. 

12. Soeder, R. H., “NASA Lewis 8- by 6-Foot Supersonic Wind Tunnel User Manual,” NASA TM-105771, Feb. 1993. 

NASA/TM—2011-217223 13
13. SAE S-16 Committee, ARP 1420, Revision B, “Gas Turbine Inlet Flow Distortion Guidelines,” Society of Automotive 

Engineers, Feb. 2002. 

14. Pointwise and Gridgen CFD mesh generation software, Pointwise, Inc., 213 S. Jennings Ave., Fort Worth, TX 76104. 

15. Chima, R. V., “TCGRID 3-D Grid Generator for Turbomachinery, User’s Manual and Documentation,” July, 2003, 

http://www.grc.nasa.gov/WWW/5810/rvc/docs.htm

16. Towne, C. E., “Wind-US User’s Guide, Version 2.0,” NASA TM-2009-215804, Oct. 2009. 

NASA/TM—2011-217223 14
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Computational Analysis of a Low-Boom Supersonic Inlet 

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14. ABSTRACT

A low-boom supersonic inlet was designed for use on a conceptual small supersonic aircraft that would cruise with an over-wing Mach 

number of 1.7. The inlet was designed to minimize external overpressures, and used a novel bypass duct to divert the highest shock losses 

around the engine. The Wind-US CFD code was used to predict the effects of capture ratio, struts, bypass design, and angles of attack on 

inlet performance. The inlet was tested in the 8-ft by 6-ft Supersonic Wind Tunnel at NASA Glenn Research Center. Test results showed 

that the inlet had excellent performance, with capture ratios near one, a peak core total pressure recovery of 96 percent, and a stable 

operating range much larger than that of an engine. Predictions generally compared very well with the experimental data, and were used to 

help interpret some of the experimental results. 

15. SUBJECT TERMS

Supersonic flow; Sonic boom; Computational fluid dynamics (CFD); Inlet 

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