Rodrick V. Chima and Stefanie M. Hirt

Glenn Research Center, Cleveland, Ohio

Robert Reger

University of Florida, Gainesville, Florida

Axisymmetric Calculations of a Low-Boom Inlet

in a Supersonic Wind Tunnel

NASA/TM—2011-217224

November 2011

AIAA–2011–3800
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Rodrick V. Chima and Stefanie M. Hirt

Glenn Research Center, Cleveland, Ohio

Robert Reger

University of Florida, Gainesville, Florida

Axisymmetric Calculations of a Low-Boom Inlet

in a Supersonic Wind Tunnel

NASA/TM—2011-217224

November 2011

AIAA–2011–3800

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
Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

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

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

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

either expressed or implied, by the National Aeronautics and 

Space Administration.

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. 
This paper describes axisymmetric CFD predictions made of a supersonic low-boom inlet 

with a facility diffuser, cold pipe, and mass flow plug within wind tunnel walls, and 

compares the CFD calculations with the experimental data. The inlet was designed for use 

on a small supersonic aircraft that would cruise at Mach 1.6, with a Mach number over the 

wing of 1.7. The inlet was tested in the 8-ft by 6-ft Supersonic Wind Tunnel at NASA Glenn 

Research Center in the fall of 2010 to demonstrate the performance and stability of a 

practical flight design that included a novel bypass duct. The inlet design is discussed here 

briefly. Prior to the test, CFD calculations were made to predict the performance of the inlet 

and its associated wind tunnel hardware, and to estimate flow areas needed to throttle the 

inlet. The calculations were done with the Wind-US CFD code and are described in detail. 

After the test, comparisons were made between computed and measured shock patterns, 

total pressure recoveries, and centerline pressures. The results showed that the dual-stream 

inlet had excellent performance, with capture ratios near one, a peak core total pressure 

recovery of 96 percent, and a large stable operating range. Predicted core recovery agreed 

well with the experiment but predicted bypass recovery and maximum capture ratio were 

high. Calculations of off-design performance of the inlet along a flight profile agreed well 

with measurements and previous calculations. 

I. Introduction 

In October and November of 2010 two similar low-boom supersonic inlets (LBSI) were tested in the 8-ft by 6-ft 

Supersonic Wind Tunnel (SWT) at NASA Glenn Research Center (GRC)1, 2. The tests were conducted by a team of 

researchers from NASA GRC, Gulfstream Aerospace Corporation (GAC), the University of Illinois at UrbanaChampaign (UIUC) and the University of Virginia (UVA). The inlets were designed for use on a small aircraft that 

would cruise at a Mach number of 1.6 at 45,000 feet3

. The inlets were designed to capture nearly 100 percent of the 

incoming supersonic flow, and to have low or zero external cowl angles to minimize external shock waves and 

thereby reduce the sonic boom signature of the aircraft. The two inlets, known as the single- and dual-stream inlets, 

are compared in Figure 1. They had identical centerbodies, struts, and cowl diameters. Both inlets were tested with 

and without flow control devices, including micro-ramps on the forebody, and micro-ramps or vane vortex 

generators in the subsonic diffuser. The flow control devices will be covered in other publications4  7 and are not 

discussed here. 

The single-stream inlet was tested to demonstrate a zero cowl angle design, and to provide CFD validation data. 

An inlet with an external cowl angle of zero degrees would theoretically produce no external disturbances, if it could 

be made infinitely thin. The single-stream inlet had a single flow path leading directly to the aerodynamic interface 

plane (AIP.) A glass window was installed in the cowl and a video camera was mounted externally and used to view 

pressure sensitive paint (PSP) and oil flow visualization on the centerbody8

. The simple axisymmetric geometry and 

the flow visualization data make this inlet especially useful for validation of CFD codes. The single-stream inlet and 

the results of these experiments will also be described elsewhere. 

Axisymmetric Calculations of a Low-Boom Inlet 

in a Supersonic Wind Tunnel 

Rodrick V. Chima and Stefanie M. Hirt 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Robert Reger 

University of Florida 

Gainesville, Florida 32603 

Abstract 

NASA/TM—2011-217224 1
The dual-stream inlet was intended to 

model flight hardware. It had a low external 

cowl angle of eight degrees, and had two 

concentric flow paths. The inner, or core, flow 

path led to the 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 

the engine and gearbox, and back to the nozzle 

stream. The dual-stream inlet is discussed in 

this paper. 

Models of the inlets were constructed at 

1/4.86 scale for tests in the SWT. To throttle 

the inlets and to measure the mass flow 

through them, they were mounted on a facility 

diffuser followed by a long cold-pipe and a 

mass flow plug (MFP). On the dual-stream 

inlet the bypass channels were throttled with 

interchangeable exit plates vented to the 

freestream. 

Research team members performed many CFD analyses of the two inlets prior to the test. The analyses were 

done with several different CFD codes and at different levels of fidelity, ranging from design of the compression 

spike using a method of characteristics code3

 to 3-D analyses of the inlets including struts and bypass geometry6, 9. 

Although many CFD analyses were performed of isolated inlets before the test, little was known about how they 

would perform in a wind tunnel environment with all the associated hardware attached. In particular, several 

questions were posed before the test: 

• What size exit plates were needed to give the desired bypass/core mass flow ratio? 

• How would the flow behave after leaving the exit plates? 

• What range of travel of the MFP would give the desired inlet operating range? 

To answer these questions, an axisymmetric CFD analysis was performed of the dual-stream inlet, bypass duct 

and exit plates, facility diffuser, cold pipe, and mass flow plug within the SWT walls. This paper describes the inlet 

model, the SWT test, and the axisymmetric CFD analysis. The results of the analysis are compared to the 

experimental data to validate the predictions, to answer the previous questions, and to explain the operation of the 

dual-stream inlet in more detail. 

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 aircraft3

. Their reference aircraft is 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 Spike10 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 3 also describes an innovative approach for designing the external compression surface of the inlet. 

This approach gives higher recovery and lower sonic boom overpressures than conventional designs. Inlets designed 

with this approach have been tested experimentally at small scale11, modeled computationally12, 13, and shown to 

have good recovery and stability characteristics. 

In the present study a large-scale dual-stream inlet was designed, analyzed computationally, and tested 

experimentally. The inlet 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 over pressures. 

The dual-stream inlet consists of an isentropic compression spike, a curved throat region, and a subsonic 

diffuser. 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 on the centerbody, which is low enough to avoid boundary-layer separation. 

The Mach number at the cowl is close to the free stream value of 1.7, which generates a strong normal shock with 

high total pressure loss. A novel bypass duct was used to capture the high-loss flow near the cowl and divert it 

Figure 1. Comparison of the single- and dual-stream inlets. 

NASA/TM—2011-217224 2
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 fan face. 

The dual-stream inlet was sized for a Rolls-Royce Tay engine, whose gearbox extends almost 160 degrees 

around the perimeter. The bypass duct was used to keep the engine gearbox out of the external flow, where it would 

contribute to boom. The duct used 10 curved vanes to direct the flow around the gearbox region. To keep the bypass 

flow subsonic it was necessary to increase the cowl area downstream of the lip, which led to an eight-degree 

external cowl angle. 

Figure 2. Dual-stream inlet in the NASA Glenn 

8-ft by 6-ft Supersonic Wind Tunnel. 

Figure 3. Equivalent axisymmetric model of the inlet 

in the wind tunnel. 

III. 8-ft by 6-ft Supersonic Wind Tunnel Test 

The inlets were tested in the SWT at NASA Glenn Research Center14. 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. Figure 2 shows a photograph of the dual-stream inlet in the SWT test section. 

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 to 1.8. The SWT is an atmospheric tunnel that 

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

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

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

Inc. in Huntington Beach, CA did the detailed design and fabrication. The models were 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 ARP1420 total pressure rakes15 were located at the AIP, using 8 rakes with 5 probes located at the centers of 

equal areas. 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 inlet models were 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 core stream of each model expanded through a facility diffuser into 

a 16-inch diameter cold pipe. The core flow was throttled using a hydraulically actuated conical mass flow plug at 

the exit of the cold pipe. The mass flow plug was calibrated to give the core-stream mass flow to within a few tenths 

of a percent. 

The dual–stream bypass channels were throttled using interchangeable choke plates at the channel exit. Baseline 

exit plates were designed to produce a mass flow ratio of m bypass / m core = 0.7 . These plates were designated the Aex

= 1.0 plates. Since the plates had sharp, semi-annular openings, they were not expected to pass the full 1-D flow but 

some lesser amount. Thus, four sets of plates were made with relative areas Aex = 1.0, 1.1, 1.2, and 1.3. The CFD 

described here was used to investigate the effects of exit plate area, and to predict the exit plate area that would give 

the desired mass flow ratio. The choke plate area was then verified by trying the different plate sets early in the test. 

The dual–stream inlet also had total pressure rakes at the exit of each bypass channel, with five probes in each 

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

NASA/TM—2011-217224 3
temperature, and by assuming that M = 1.0 at the exit plate. A discharge coefficient CD = 0.873 was needed to 

correct the overall capture ratio to 1.0 at M = 1.8 with the mass flow plug fully open. Since the discharge coefficient 

probably varies with operating conditions, the accuracy of bypass flow rate measurement is uncertain. 

Large 26.5-inch diameter windows provided optical access to the test section, and a schlieren system was used to 

visualize the flow ahead of the inlet. A high-speed Phantom camera operating at 20004000 frames per second 

recorded steady or unsteady schlieren images. 

IV. Computational Model 

A. Computational Grids 

For the axisymmetric CFD model the 8- by 6-foot cross section of the SWT was modeled as a circle with an area 

of 48 ft2

, i.e., a radius of 3.91 ft. as shown in Figure 3. 

Figure 4. Computational grid for the inlet, facility diffuser, cold pipe, and mass flow plug. 

The computational grid for the inlet, bypass, facility diffuser, cold pipe, and mass flow plug within the tunnel 

walls was generated using Pointwise16. The grid, shown mirrored top-to-bottom in Figure 4, had 144,525 points in 7 

zones. The proportions of this side view of the axisymmetric grid are almost the same as the side view of the tunnel. 

The grid spacing at the walls was 1105 inches, giving y+

 = 1 to 2 at the first point off the walls. Leading edges 

of the cowl and splitter were modeled as 2:1 ellipses with 6-10 points along each surface, to give adequate resolution 

of bow shocks. 

The inlet, bypass, and external zones were 

comparable to grids used in Ref. 13, where a 

grid refinement study showed that the grid 

resolution was sufficient to predict the inlet 

recovery within 0.13 percent of a much coarser 

grid. The remaining zones had similar 

resolution, but they were not tested explicitly 

here. 

The bypass duct of the actual inlet model 

was not axisymmetric. It was split into 10 

curved passages designed to route the inlet 

flow around a 160-degree area where the 

gearbox of an engine would be located. Half 

way through the duct the 10 passages merged 

into five passages that were partially closed by 

the exit plates described earlier. To model the Figure 5. Bypass duct exit plate model. 

NASA/TM—2011-217224 4
gearbox blockage using an axisymmetric CFD model, the radius of the cowl inner mold line was reduced to make 

the axisymmetric passage area equal to the actual 3-D area at each axial location. The outer mold line was not 

modified. This is seen in Figure 5, which shows an enlargement of the grid near the bypass exit plates. While the 

actual cowl on the model was a thin shell, the CFD cowl was very thick to account for gearbox blockage. 

The bypass exit plates were modeled computationally by applying inviscid wall boundary conditions over part of 

the radial grid line at the location of the plates, also shown in Figure 5. The height of the inviscid wall was chosen to 

leave an axisymmetric area equivalent to the semi-annular area of the exit plates. Slight modifications were made to 

the grid clustering in this region to accommodate the four different plate areas with Aex ranging from 1.0 to 1.3. 

Due to limitations of the axisymmetric model, changes in the cowl geometry, and the crude exit plate model, the 

CFD analysis was not expected to give detailed information about the bypass flow. However, it was expected to give 

reasonable predictions of bypass recovery characteristics, which will be shown later. 

In the experiment, the core flow was 

throttled using a 16-inch diameter 

hydraulically actuated mass flow plug17, 

shown in outline in Figure 6. For the 

computations, a grid was generated 

around the MFP, shaft, and hydraulic 

cylinder using Pointwise16. The MFP 

geometry was specified using a database 

surface, and the grid was attached to the 

database. Translating the database in 

Pointwise automatically transformed the 

surrounding grid, allowing grids to be 

produced at new MFP positions with a 

single operation. When the calculations 

were made, Pointwise did not export grids in a format suitable for Wind-US, so the grids were imported to Gridgen, 

where boundary conditions were applied using a glyph script. Thirty different grids were generated to model all of 

the exit plates and MFP positions considered in this paper. 

B. CFD Solution Scheme 

The dual-stream inlet was analyzed using the Wind-US code18, 19. The Reynolds-averaged Navier-Stokes 

(RANS) equations were discretized using the Roe upwind scheme with a minmod limiter and the Menter SST 

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

with a Courant number of 1.0. 

Boundary conditions were specified as follows: 

• Tunnel inlet conditions were supersonic inflow with Mach number specified. 

• Tunnel exit conditions were all extrapolated. 

• The actual tunnel had porous walls to remove wall boundary layers and reduce shock/boundary-layer 

interaction. The bleed flow was unknown, so the tunnel walls were modeled using inviscid wall boundary 

conditions. 

• Viscous wall boundary conditions were used for the entire inlet model, except for the bypass exit plates 

described above. 

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

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

reinitialized. The solution was run 15,000 iterations, which converged the capture ratio and recovery to plotting 

accuracy. 

Subsequent cases were restarted from previous solutions and run 15,000 iterations to convergence. All 

calculations were run on a cluster of 6 CPUs running at 3.2 GHz, which took about 1.5 hours per case. 

V. Comparison of CFD and Test Results 

A. Bypass Exit Plate Selection 

One objective of the present work was to determine the size of bypass exit plates that would give the desired 

flow rates through the inlet. CFD predictions of inlet performance maps were made for four relative plate sizes, 

Aex = 1.0, 1.1, 1.2, and 1.3. 

Figure 6. Mass flow plug translation. 

NASA/TM—2011-217224 5
The computed performance maps, sometimes known as cane curves because of their shape, are shown in Figure 

7. For each exit plate area, the MFP position was varied to produce a curve of core total pressure recovery versus 

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

2 ) . 

The Aex = 1.0 plate had low recovery 

and a maximum capture ratio less than 

0.95. The Aex = 1.1 plate had better 

recovery and a maximum capture ratio 

of 0.98. The two largest plates, Aex = 1.2 

and 1.3, had the highest recoveries and 

maximum capture ratios of 0.99. It was 

clear that the two smaller plate areas 

could not pass the desired flow through 

the bypass stream, so the Aex = 1.2 

plates were recommended for the test. 

Some initial experimental data was 

taken using the plates with Aex = 1.2. 

Later these plates were replaced with 

each of the other three sets, and one 

cane curve was generated for each set. 

The results were similar to the 

predictions, and the plates with Aex = 

1.2 were used for the remainder of the 

tests. 

A second objective of this work was to determine the range of MFP travel that would be needed to cover the 

desired inlet operating range. A nominal core mass flow rate was estimated by assuming a core recovery of 0.95 and 

an AIP Mach number of 0.65. The MFP plug exit area was found with the continuity equation, and the plug position 

was found from the plug angle. A CFD solution for this position gave a capture ratio around 0.97. Subsequent 

calculations were made by moving the MFP in increments of 0.1 or 0.2 inches. The maximum opening occurred 

when the core flow choked. The minimum opening occurred when the capture ratio was around 0.9, which was 

thought to be near the stall point of the engine. The total range of MFP travel was found to be one inch for all the 

cases shown in Figure 7. 

In the experiment the MFP was set fully open to minimize blockage during SWT startup. The plug was then 

closed in regular increments until the inlet went into buzz, at a capture ratio of 0.60 for this case. No attempt was 

made to determine the buzz point computationally. 

B. Design Point Performance 

This section compares the CFD predictions of the dual-stream inlet using the exit plates with Aex = 1.2 to the 

experimental data at the design point, M = 1.7. First the CFD results are used to describe the overall flow, and then 

detailed comparisons are made of AIP profiles, recovery and mass flow characteristics, and a centerbody pressure 

distribution. 

Figure 7. Computed core recovery vs. capture ratio for 

four bypass exit plate areas. 

NASA/TM—2011-217224 6
Figure 8. Schlieren image and computed Mach number contours, M = 1.7. 

The overall flow field is shown in Figure 8, where a schlieren image is compared to computed Mach number 

contours. The images were taken at equivalent MFP positions near the peak recovery operating point, and the shocks 

line up nicely. The compression waves from the spike extend slightly outside the cowl, indicating that the inlet is 

operating just below full capture. 

The core flow decelerates to M  0.65, the desired fan face Mach number, at the AIP. The flow decelerates 

further in the facility diffuser, which has a large recirculating region along the centerline. The computed size of this 

region depends on MFP position and on the turbulence model, and it probably contributes to an under prediction of 

the computed core flow discussed later. The core flow chokes near the exit of the MFP, and then expands to 

supersonic speeds downstream. 

The bypass flow chokes at the exit plates, over expands to M  3.0 just downstream, and then shocks to subsonic 

speeds along the cold pipe. This high loss flow hugs the cold pipe to the exit of the grid. 

Recoveries were measured at the AIP using eight 

total pressure rakes spaced 45 degrees apart. Six 

static probes and one dynamic probe were located 

on each rake, and wall static pressures were 

measured on the centerbody. Figure 9 shows 

spanwise profiles of recovery at the operating points 

indicated by diamonds in Figure 7. Seven of the 

eight rakes were located away from wakes from the 

centerbody struts, and they measured nearly 

identical recoveries (blue circles). The axisymmetric 

solution (solid black line) agrees closely with these 

seven rakes, except in the hub boundary layer, 

where the computed recovery is 2-3 percent higher 

than the measured recoveries. One rake was located 

at bottom dead center immediately behind a strut, 

and measured recoveries 5-6 percent lower than the 

rest (red squares). The axisymmetric CFD model 

does not include the struts and cannot predict this 

feature. 

Figure 9. Recovery profiles along the AIP rakes. 

NASA/TM—2011-217224 7
Overall core recoveries are compared in 

Figure 10. Two curves are shown for 

measured recovery – one computed using 

all eight rakes, and one computed by 

ignoring the wake behind the strut. The 

CFD results are about 0.3 percent higher 

than the 8-rake curve, and about 0.3 percent 

lower than the 7-rake curve, consistent with 

the spanwise profiles shown above. 

The maximum computed capture ratio is 

about one percent higher than the measured 

value. As mentioned in Section III, there is 

some uncertainty in the measured bypass 

flow rate that may contribute to the 

discrepancy. Later we will show that there 

is some discrepancy in the computed mass 

flow split between the two channels that 

also affects these results. 

The minimum computed capture ratio 

was chosen to cover the expected operating 

envelope of an aircraft engine. The 

minimum experimental capture ratio 

occurred when the inlet went into buzz at a 

capture ratio of 60 percent for this case. The 

experimental buzz points were omitted here 

for clarity. 

The two black diamonds indicate 

computed and measured points at the same 

MFP position. These two points were 

compared in the schlieren image (Figure 8), 

the spanwise profiles (Figure 9), and later 

for centerbody pressures (Figure 13). 

Although the two points differ by three 

percent in capture ratio, other quantities 

compare well at the same MFP position. 

Overall bypass recoveries are compared 

in Figure 11. The measured recovery is an 

area-weighted average of 25 probes in five 

bypass channels, while the computed 

recovery was integrated across the grid at 

the axial location of the rakes. The 

computed bypass recoveries are about 0.8 

percent higher than the data. Averaging the 

computed results at the approximate rake 

locations increases the discrepancy 

somewhat. Since the axisymmetric CFD 

model does not include sidewall losses, 

secondary flows, or other 3-D effects, it is 

not surprising that it over predicts the 

bypass recovery. 

Discrepancies in the mass flow split are 

examined in Figure 12, where corrected 

mass flow for the two streams is plotted 

against MFP position. The corrected mass 

flow is defined by m c = m /  , where 

Figure 10. Core recovery vs. capture ratio. 

Figure 11. Bypass recovery vs. capture ratio. 

Figure 12. Corrected mass flow rate vs. mass flow plug 

position. 

NASA/TM—2011-217224 8
 = T0 / T0ref is the total temperature ratio and  = P0 / P0ref is the total pressure ratio. Nominal tunnel operating 

conditions were used both for the reference conditions and for the CFD. Use of corrected flow removes variations in 

tunnel total conditions between data points. 

The core flow was measured with a calibrated mass flow plug, and is believed to be correct to within a fraction 

of a percent17. The computed core flow matches the measured core flow almost exactly at low MFP positions (plug 

closed), but under predicts the maximum core flow by 1.1 lb/sec (plug open). This discrepancy was not seen in 

calculations of the isolated inlet13, and may be caused here by the large recirculating region in the facility diffuser 

discussed earlier. 

The experimental bypass flow was estimated from bypass total pressure rake measurements and a discharge 

coefficient calibrated at M = 1.8, using the procedure described earlier. Since the discharge coefficient probably 

varies with Mach number and capture ratio, the accuracy of the bypass flow measurement is unknown. 

The computed bypass flow rate is about 11 percent high, which could be due partly to experimental uncertainty, 

but it is more likely due to the simplified exit plate model used in the computations. In the experiment the exit plates 

had semi-annular slots covering 40 degrees of arc, with a sharp lip on each edge. In the computations the exit plates 

were modeled as a full annular slot covering 360 degrees, with a sharp lip on the bottom edge modeled as an 

inviscid wall. The experimental and modeled slots had the same areas, and thus would have the same 1-D flow rates. 

However, viscous and 3-D effects in the experiment would probably give a lower actual flow rate. 

Thus the CFD under predicts the core flow and over predicts the bypass flow and total capture ratio. Throughout 

this paper the CFD and measurements have been compared at equivalent MFP positions rather than capture ratios. 

As shown in the next figure, that strategy 

tends to match static pressures closely. 

Figure 13 compares computed and 

measured pressure coefficients along the 

centerbody, where 

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

The two results were taken at the same 

MFP position, shown by the black diamonds 

in Figure 10. The computations agree very 

well with the measurements, except for a 

region where five struts were attached to the 

centerbody. These struts are not included in 

the axisymmetric CFD model and their 

effect is not seen in the computations. 

C. Off-Design Performance 

A series of calculations was run to 

simulate the behavior of the inlet along a flight profile. The MFP position was fixed, and the upstream Mach number 

was varied from nominally 1.8 to 1.3 in increments of 0.1. This resulted in approximately constant corrected core 

flow for the six operating points. 

Figure 13. Pressure distribution along the centerbody, 

 M = 1.137. 

NASA/TM—2011-217224 9
M = 1.779 M = 1.452

M = 1.664 M = 1.352

M = 1.555 

(Schlieren image at different MFP position)

M = 1.253 

(No data was taken at this Mach number) 

Figure 14. Variation of schlieren images and computed Mach number contours with tunnel Mach number. 

Color bar ranges from M = 0.0 to 1.9. 

Mach number contours with corresponding schlieren images at the six operating points are shown in Figure 14. 

At Mach 1.779 all the compression waves from the spike enter the cowl because the inlet is at full capture. As the 

Mach number decreases the wave angles decrease, the normal shock moves forward on the spike, and the inlet spills 

more flow. In all cases the computed shock positions match the schlieren images, except at M = 1.555 where no 

schlieren image was taken at the correct MFP position. And in all cases the Mach number stays around 0.65 at the 

AIP, which is the design fan face Mach number. At M = 1.253 a normal shock spans the tunnel at the bypass exit. 

No data was taken at this Mach number, but the normal shock was observed to pass across the inlet every time the 

tunnel was started or stopped. 

Core recovery is plotted as a function of Mach number in Figure 15. The computed recovery curve (green 

triangles) is in excellent agreement with the measured recovery (black circles) and a previous inlet-fan calculation12

(red dots.) For comparison, the recovery of a pitot inlet for an F-16 aircraft is also shown20 (dashed line). The 

recovery of the F-16 inlet is very similar to the present inlet up to M = 1.4, but then it decreases rapidly with normal 

shock losses. The core recovery of the present inlet stays high up to M = 1.6, partly due to the isentropic 

compression design, and partly because losses are confined to the bypass stream. 

NASA/TM—2011-217224 10
VI. Summary and Conclusions 

This paper describes axisymmetric CFD 

predictions made of a low-boom dual-stream 

inlet, a facility diffuser, cold pipe, and mass flow 

plug in a supersonic wind tunnel. The inlet was 

designed for use on a small supersonic aircraft 

that would cruise at a Mach number of 1.6, with 

an over-wing Mach number of 1.7. It was tested 

in the 8-ft by 6-ft Supersonic Wind Tunnel at 

NASA Glenn Research Center in the fall of 2010. 

The paper discusses the inlet design and the test 

briefly, and discusses the calculations in detail. 

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 at the design point. The bypass 

recovery was much lower, between 84 and 87 

percent. 

CFD predictions were made with the Wind-US CFD code on a grid with 144,525 grid points. The Roe upwind 

differencing scheme and SST turbulence model were used. Bypass duct blockage was modeled by decreasing the 

radius of the outer bypass wall. Bypass exit plates were modeled by applying inviscid wall boundary conditions 

along grid lines normal to the inner bypass wall. The mass flow plug grid was transformed for each plug location. 

Predictions were compared with the experimental data, and the following results were found: 

• Computed shock positions agreed well with schlieren images. 

• The computations were used to select the initial bypass plate area used in the experiment. This area was 

expected to give nearly full capture with the mass flow plug wide open. Four plate areas were tested, and the 

predicted plate area gave the desired capture ratio. 

• The calculations slightly under predicted the core flow, possibly due to blockage from a large recirculating 

region predicted in the facility diffuser. It is not known if this recirculation occurred in the experiment. 

• The calculations over predicted the bypass flow, probably due the simple exit plate model used in the 

calculations. 

• Predicted core recoveries were within 0.3 percent of the measured recoveries. 

• Predicted bypass recoveries were 0.8 percent higher than the measured recoveries, probably due to lack of 3-D 

effects and viscous walls in the axisymmetric CFD model. 

• Predicted centerbody pressure distributions agreed very well with the experiment, except near struts that were 

not included in the axisymmetric analysis. 

• The CFD was used to predict off-design performance of the inlet for a simulated flight profile. The predicted 

core recovery agreed well with the experiment, and with a previous inlet/fan calculation. 

Despite discrepancies between the computed and measured flow rates, the computations provided a useful 

system-level model of the inlet with all its associated wind-tunnel hardware. They were used to select the bypass 

exit plate areas, and to specify mass flow plug travel before the test. Finally, they have provided the only look at the 

SWT flow field away from the schlieren window. 

References 

1. 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.) 

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. 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. 

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

Inlets,” NASA TM-2010-216350, May 2010. 

Figure 15. Inlet core recovery vs. Mach number. 

NASA/TM—2011-217224 11
5. Vyas, M. A., Hirt, S. M., Chima, R. V., Davis, D. O., Wayman, T. R., “Experimental Investigation of Micro-Vortex 

Generators on a Low Boom Supersonic Inlet,” 29th AIAA Applied Aerodynamics Conference, June 27-30, 2011 (to be 

published.) 

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

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

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. Herges, T. G., Dutton, J. C., Elliott, D. S., Hirt, S. M., and Bencic, T. J., “Surface Flow and PSP Measurements in the LargeScale Low-Boom Inlet,” AIAA 29th Applied Aerodynamics Meeting, June 27-30, 2011 (to be published.) 

9. Chima, R. V., “Computational Analysis of a Low-Boom Supersonic Inlet,” AIAA 29th Applied Aerodynamics Meeting, 

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

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

11. 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. 

12. 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. 

13. 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. 

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

15. SAE S-16 Committee, ARP 1420, Revision B, “Gas Turbine Inlet Flow Distortion Guidelines,” Society of Automotive 

Engineers, Feb. 2002. 

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

17. Friedlander, D. J., Saunders, J. D., and Frate, F. C., “Computational Fluid Dynamics Analysis of the NASA Glenn Research 

Center 16 Mass Flow Plug,” (NASA TM to be published.) 

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

19. Nelson, C. C., “An Overview of the NPARC Alliance’s Wind-US Flow Solver,” AIAA Paper 2010-27, Jan. 2010. 

20. Hunter, L. G., and Cawthon, J. A., “Improved Supersonic Performance for the F-16 Inlet Modified for the J79 Engine,” J. 

Propulsion, Vol. 1, Jan. – Feb. 1985. 

NASA/TM—2011-217224 12
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Axisymmetric Calculations of a Low-Boom Inlet in a Supersonic Wind Tunnel 

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14. ABSTRACT

This paper describes axisymmetric CFD predictions made of a supersonic low-boom inlet with a facility diffuser, cold pipe, and mass flow 

plug within wind tunnel walls, and compares the CFD calculations with the experimental data. The inlet was designed for use on a small 

supersonic aircraft that would cruise at Mach 1.6, with a Mach number over the wing of 1.7. The inlet was tested in the 8-ft by 6-ft 

Supersonic Wind Tunnel at NASA Glenn Research Center in the fall of 2010 to demonstrate the performance and stability of a practical 

flight design that included a novel bypass duct. The inlet design is discussed here briefly. Prior to the test, CFD calculations were made to 

predict the performance of the inlet and its associated wind tunnel hardware, and to estimate flow areas needed to throttle the inlet. The 

calculations were done with the Wind-US CFD code and are described in detail. After the test, comparisons were made between computed 

and measured shock patterns, total pressure recoveries, and centerline pressures. The results showed that the dual-stream inlet had excellent 

performance, with capture ratios near one, a peak core total pressure recovery of 96 percent, and a large stable operating range. Predicted 

core recovery agreed well with the experiment but predicted bypass recovery and maximum capture ratio were high. Calculations of offdesign performance of the inlet along a flight profile agreed well with measurements and previous calculations. 

15. SUBJECT TERMS

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

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