NASA Technical Memorandum 1023 17 

AIAA-89-2355 

Mach 5 Inlet CFD and Experimental Results 

Lois J. Weir 

National Aeronautics and Space Administration 

Lewis Research Center 

Cleveland, Ohio 

and 

D.R. Reddy and George D. Rupp 

Sverdrup Technology, Inc. 

NASA Lewis Research Center Group 

Cleveland, Ohio 

Prepared for the 

25th Joint Propulsion Conference 

cosponsored by the AIAA, ASME, SAE, and ASEE 

Monterey, California, July 10-12, 1989 

[fASA-TH-102317) HACH-5 INLET CFD aND ~89-27670 

- 

EXPERXHENTAL BESULTS {NASA. Lewis Research 

Center) 16 p CSCL 21E 

Unclas 

63/07 0225968 
MACH 5 INLET CFD AND EXPERIMENTAL RESULTS 

Lois J. Weir* 

National Aeronautics and Space Administration 

Lewis Research Center 

Cleveland, Ohio 44135 

and 

D.R. Reddyq and George D. Ruppi 

Sverdrup Technology, Inc. 

NASA Lewis Research Center Group 

Cleveland, Ohio 44135 

ABSTRACT 

An experimental research program was conducted 

in the NASA Lewis Research Center 10 ft. by 10 ft. 

supersonic wind tunnel. The two-dimensional inlet 

model was designed to study the Mach 3.0 to 5.0 speed 

range for an "over-under" turbojet plus ramjet 

propulsion system. The model was extensively 

instrumented to provide both analytical code 

validation data as well as inlet performance 

information. Support studies for the program include 

flow field predictions with both 3-dimensional 

parabolized Navier-Stokes (PNS) and 3D full NavierStokes (FNS) analytical codes. Analytical 

predictions and experimental results are compared. 

h 

M 

P 

P 

STA 

X 

Y 

Subscripts 

cl 

P 

1 

m 

NOMENCLATURE 

angle of attack 

height 

Mach number 

static pressure 

total pressure 

station 

axial dimension 

vertical dimension 

cowl lip 

pitot 

first ramp 

freestream 

*Aerospace Engineer; Member AIM. 

?Supervisor, Turbomachinery Analysis Section; 

*Wind Tunnel Test Engineer. 

Member AIM. 

BACKGROUND 

In 1980, A joint research program led by NASA 

Langley with NASA Lewis as a partner, Lockheed 

California as the prime contractor, and Pratt & 

Whitney as subcontractor was initiated which would 

address critical technology issues in the Mach 3 to 

6 speed range. Prior to this time, little focussed 

research had been done in this flight arena. The 

specific purpose of the program was to develop a 

concept for a Mach 5 cruise aircraft, identify the 

propulsion system required to power it, and define 

its integration with the aircraft. The aircraft 

configuration chosen as a result of the study, 

discussed in Reference 1 and shown in Figure I, would 

employ 4 propulsion modules (two under each wing). 

The propulsion system chosen for the aircraft is an 

over/under turbojet plus ramjet system with dualflow, two-dimensional inlet and nozzle. 

Figure 1: Artist's conception of Mach 5 cruise 

aircraft 

The modes of operation for the over-under 

turbojet-ramjet system are shown for various flight 

speeds in Figure 2. The upper flow path of this 

1 ORIGINAL PAGE 

BLACK AND WHITE PIiOTOGRAPH 
TURBOJET 

COLD-FLOW RAMJET 

MACH 0 TO 1 

TURBOJET a RAMJET 

MACH 2.9 

- RAMJET ONLY 

MACH 3 

RAMJET ONLV 

MACH 6 

Figure 2: Propulsion system operation 

dual-flow system provides airflow to the turbojet 

engine, which alone powers the aircraft at takeoff, 

with cold flow through the ramjet in the lower flow 

path. Near Mach 1, the ramjet is ignited, and both 

systems operate simultaneously. At Mach 2.5 the 

turbojet begins to spool down, and near Mach 3, the 

upper duct closes off completely. From Mach 3 to 

Mach 5 (cruise speed), the ramjet alone powers the 

vehicle. This propulsion system is designed to take 

maximur advantage of the turbojet engine in the lower 

speed range where it is most effective. and the 

ramjet in the high speed range, where it is most 

efficient. A discussion of the issues associated 

with the design of inlets for aircraft in the high 

supersonic speed range may be found in Reference 2. 

STATION (in.) 

-40 -20 0 20 10 60 11,11,,,1,,,1 

+,I 

-3 -2 -1 0 1 2 3 4 

Figure 3: Inlet compression system to ramp 

shoulder 

An inlet test program was planned for the Lewis 

Research Center 10 ft. by 10 ft. supersonic wind 

tunnel (10x10 SWT) as a result of the aircraft study. 

Only the lower (ramjet) portion of the dual flow duct 

was to be tested, representing supersonic flight 

between Mach 3 and 5. The inlet for the overhnder 

turbojet-ramjet propulsion system was designed 

inviscidly usingthe method of characteristics (MOC), 

with the compression surfaces adjusted to compensate 

for the viscous displacement effects as predicted by 

viscous boundary layer codes. The inlet employs four 

ramps and the cowl to provide the required external 

compression for operation at design (Mach 5) 

conditions. The fourth ranp angle is varied for offdesign operation. The oblique shocks generated by 

the ramps are designed to intersect very near the 

cowl lip, and the cowl shock is cancelled at the 

inlet shoulder. A cross-section showing the 

aerodynamic contours of the inlet to a station just 

aft of the shoulder is shown in Figure 3. The X- and 

Y-dimensions are nondimensionalized to the cowl lip 

height, h,l. At cruise conditions, the freestream 

airflow is oriented at a 9' angle relative to the 

first ramp surface. The wedge angles of the last 

three external ramps with reference to the first are 5'. lo', and 15', respectively. The cowl lip provides an additional 5' of turning. Oblique shock wave 

locations are indicated by the dashed lines. The 

local Mach numbers, predicted inviscidly, are shown 

for various regions in the flow field. Specifics of 

the inlet design are fully discussed in Reference 3. 

The first three-dimensional computational fluid 

dynamics (CFD) calculation of the Mach 5 inlet flow 

field was accomplished by Benson (Reference 4) using 

a 3D parabolized Navier-Stokes (PNS) code which is 

discussed in Reference 5. This analysis predicted a 

strong three-dimensional effect caused by the 

interaction of the ranp shock waves with the sidewall 

boundary layers. These interactions generated a 

vortical flow field within the sidewall boundary 

layer, which led to boundary layer separations 

downstream of the cowl leading edge. This effect is 

illustrated in Figure 4, which shows predicted crosssectional total pressure contours at selected 

stations between the cowl lip and the shoulder. In 

this figure, the cowl surface is on top, with the 

ramp surface below. Each plane represents half of 

the inlet area at its respective station, with the 

left hand border of each cross-section being the 

centerline of the inlet. Massive separations such as 

that shown in the cowl corner would likely lead to 

an inlet unstart. The inlet model was later modified 

to include porous boundary layer bleed areas on the 

sidewalls in the vicinity of the shock-boundary layer 

interactions, and on the cowl in the corners near the 

leading edge and sidewalls, as shown in Figure 5. 

This bleed was intended to control the threedimensional effects, and prevent separation on the 

cowl. 

Figure 4: PNS analysis: total pressure contours at 

selected stations between the cowl lip 

and shoulder 

As the development and use of hypersonic CFD 

codes grew, interest in the use of the Mach 5 inlet 

as a tool for code calibration and validation 

developed. The inlet, which was originally intended 

as a technology demonstrator, was modified to include 

additional instrumentation on the ramp and sidewalls 

in regions where 3-dimensional effects were 

predicted. Data from this instrumentation is now 

available to be used for comparison with analytical 

predictions. Data upstream of the inlet shoulder and 

some initial CPD / data comparisons are to be 

presented in this paper. 

2 
Bleed regions in 

Bleed reglons added as 

result of CFD anolysls 

Figure 6: Bleed regions upstream of the ramp 

shoulder 

APPARATUS 

A photograph of the Mach 5 inlet, a 1/3 scale 

model, installed in the Lewis 10x10 SWT is shown in 

Figure 6. The inlet was mounted on a large 

trapezoidal plate. By actuating this "accelerator" 

plate to an 8.5' negative angle of attack, the Mach 

3.5 tunnel airflow was expanded to approximately Mach 

4.1 on the inlet first ramp. This simulated the flow 

conditions present on the inlet first ramp when it is 

operating at Mach 5 flight conditions with the 

incoming airflow at an angle of attack of go, as 

previously indicated in Figure 3. The first oblique 

shock was not reproduced, but the data was corrected 

for the loss that would occur across it. (Recovery 

across this initial oblique ahock is 0.901.) The 

accelerator plate was 100 inches wide, the inlet 

capture height (with the first ramp at 0' relative to 

freestream) was 16 inches, capture width was 16 

inches, and the overall length of the model was 20 

feet. 

Figure 6: Mach 5 inlet model installed in Lewis 

Research Center 10 ft. by 10 ft. 

supersonic wind tunnel 

The inlet made use of extensive variable 

geometry, including a collapsible ramp and remotely 

variable bleed exits on the ramp, cowl, and 

sidewalls. The flow through the upstream bleed 

regions shown earlier (in Figure 5) was changeable 

(between runs) by installing various sizes of flow 

restrictors in the exit ducts, (Ramp bleed flow from 

the region upstream of the shoulder shown in the 

sketch was remotely variable.) Variation in engine 

airflow was provided by a remotely actuated choked 

exit plug. Model instrumentation included surface 

static taps, fixed total pressure rakes, translating 

pressure probes, and dynamic pressure transducers. 

The locations of instrumentation upstream of the 

shoulder are presented in Figure 7. Static taps on 

the ramp and cowl were located on the inlet 

centerline; and 4 inches, 6 inches, and 7.5 inches 

off the centerline. Static taps on the sidewalls 

were located in approximately vertical rows as shown. 

Eight boundary layer rakes on the ramp were located 

on the centerline and approximately 7.5 inches off 

the centerline. Translating pitot probes were 

located on all surfaces as indicated in the sketch. 

Four corner rakes, mounted to the cowl, were oriented 

in a 45' angle between the cowl and sidewalls. A 0.5 

inch strip of grit was applied near the leading edges 

of the ranp and sidewall to ensure that a fully 

turbulent boundary layer was ingested by the inlet. 

RESULTS 

In Figures 8 - 30, experimental data for 

instrumentation upstream of the inlet shoulder are 

presented. For all data presented here, the tunnel 

freestream Mach number, M,, is 3.49; and tunnel total 

pressure, Po, is 35.1 psia (this is also total 

pressure on the first ramp, PI.) The angle of attack 

of the accelerator plate, a, is -8.66', providing a 

Mach number on the 1st ramp, Mi, of 4.1. For all 

data, boundary layer bleed was removed from the ramp 

bleed region upstream of the shoulder. Bleed from 

this area removes approximately 0.5% of the total 

inlet capture mass flow. Data labeled "bleed" 

represents a configuration in which bleed was also 

removed from the regions on the sidewall and cowl 

shown in Figure 5. This bleed represents 

_approximately 8.8% of the capture mass flow.) A "no 

bleed" label indicates that porous bleed holes on the 

sidewalls and cowl were sealed. 

Figures 8 and 9 show the static pressure 

distributions on the ramp for conditions both with 

and without bleed. Static pressure is 

nondinenaionalized to total pressure on the inlet 

first ramp, PI, and the station number is divided by 

the cowl lip height, h,l. The locations of the cowl 

lip and shoulder are indicated by vertical dashed 

lines on the plot. Similarly, cowl static pressure 

distributions are shown in Figures 10 and 11. For 

both of these sets of data, little difference between 

the bleed and no bleed configurations may be 

observed. Sidewall static pressure profiles for 

vertical rows of taps located at the stations 

indicated are shown in Figures 12 - 14. Figure 12 

shows pressure distributions for the three most 

upstream rows of static taps. Since bleed and no 

bleed distributions were very similar, only the bleed 

data are shown. Figures 13 and 14 show sidewall 

static pressure distributions at five stations 

further downstream for both bleed and no bleed 

configurations. The sidewall static pressure 

profiles do not seem to indicate the locations of 

oblique shocks from the ramps, except near the ramp 

in the cowl compression region. This is most likely 

due to the boundary layer flow migration on the 

walls caused by the sidewall / shock interactions. 

Figures 15 - 19 show ramp boundary layer rake 

3 
M 

Figure I: Instrumentation upstream of inlet shoulder 

0 CENTERLINE 

0 7.5" OFF CENTERLINE 

Shoulder 

Cowl lip- 

> 

- -3 

1 

I 

I 

I 

I 

I 

I 

Axial distance from beginning 

of second ramp, x/h,l 

I Figure 8: Ramp static pressure distribution: no bleed 

0 CENTERLINE 

0 7.5" OFF CENTERLINE 

Shoulder 

Cowl lip- 

- 

I 

I 

I 

I 

I 

I 

I 

Axial distance from beginning 

of second ramp) x/hcl 

Figure 9: Ramp static pressure distribution: bleed 

4 
.30 

oa 

OE 

.r( .10 4Jk 

.05 

.oo2 

0 CENTERLINE 

0 7.5" OFF CENTERLINE 

Cowl 

Shoulder 

lipI I 

I I 

I I 

I I 

I 

I I I I I1 11 I I 

1 2.5 3.0 3.5 4.0 4.5 5.0 

Axial distance from beginning 

of second ramp, x/h,l 

Figure 10: Cowl static pressure distribution: no 

bleed 

.30 

0 CENTERLINE 

0 7.5" OFF CENTERLINE 

ShoulderI 

I Cowl lip-, 

I I 

I I 

I I 

I 

I 

I 

I 

I 

I 

I 

I 

I 1m 

I I 

I I 

'002!0 215 310 315 410 415 20 

Axial distance from beginning 

of second ramp, x/h,l 

Figure 11: Cowl static pressure distribution: 

bleed 

profiles. Pitot pressure (nondimensionalized to PI) 

is plotted vs. height from the ramp 

(nondimensionalized to h,l). Higher pitot pressures 

generally indicate a lower supersonic local Mach 

number, except when the tube is located in the 

subsonic boundary layer. Circular symbols represent 

rakes located on the inlet centerline, and square 

symbols denote rakes located 0.5 inch away from the 

sidewall. The profiles obtained without bleed were 

r 0 STA4.4 

0 STA23.0 0 STA35.4 

Flagged symbols denote 

nearest ramp stotic pressure 

Ratio of local static to 

first ramp total pressure, p/pI 

Figure 12:Sidewall static pressure distribution, 

stations 4.4, 23.0, and 35.4 

.6 

.5 7 

v 

c 

\ 

c 

a 

E 

(d 

k 

- .4 

E .3 

k 

u-l 

P) 

0 

(d 

+J 

VI 

'4 

CI 

E .2 

.I 

.o 

P 

0 STA54.7 

0 STA57.5 0 STA60.5 

a STA63.7 

V STA67.0 

Flagged symbols denote 

nearest ramp static pressure 

Ratio of local static to 

first ramp total pressure, p/p, 

Figure 13:Sidewall static pressure distribution 

stations 54.7, 57.5, 60.5, 63.7, and 

67.0: no bleed 

not significantly different from those with bleed, 

so only bleed data are shown. The first set of 

rakes, shown in Figure 15, are located near the end 

of the first ramp. The next eet (Figure 16) is 

located near the start of the second ramp. The drop 

in pitot pressure on the outermost tubes of the rake 

indicates the location of the intersection of the 

oblique shock from the first ramp. A considerable 

5 
I 

.6 

.5 

.-. " 

e 

s 

a 

E 

iu 

k 

0 

k 

w 

Q 

iu 

w 

u) 

.rl 

a 

. .4 

E .3 

; .2 

.1 

0 STA54.7 

0 STA57.5 0 STA60.5 

A STA 63.7 

V STA67.0 

Flagged symbols denote 

nearest ramp static pressure 

- I &I I I I1 1 I I 

I .oY -04 .06 .08 .10 

Ratio of local static to 

first ramp total pressure, p/Pl 

Figure 14: Sidewall static pressure distribution, 

stations 54.7, 57.5, 60.5, 63.7, and 

67.0: bleed 

.20 

.18 

. .16 

.14 

7 

5 

s 

Q 

0 

k 

2 .12 4 

m 

0 

0 .10 

6 .08 

I+ 

k 

W 

Q 

u 

C 

w 

.06 

m .04 

0 CENTERLlNE 

0 7.5" OFF CENTERLINE 

u) 

';1 .02 tRatio of local pitot to first 

ramp total pressurer PdPl 

Figure 15: Ramp boundary layer profiles. station 

-3.0 

difference between the profile on the centerline and 

that near the wall can be seen at this 

station. Figure 17 shows the pitot pressure profile 

on a rake located on the inlet centerline near the 

end of the second ramp, and Figure 18 shows rake 

profiles near the start of ramp 3. Little difference 

~~ 

.20 

5 .18 

.16 

(d .14 

e 

c 

Q 

0 

vc 

k 

2 .12 

I+ 

I3 

0 

4 

0 .lo 

.08 

.06 

2 .04 

8 

k 

W 

PI 

0 

4J 

u) 

-4 a .02 

.oo 

0 CENTERLINE 

0 7.5" OFF CENTERLINE 

Ratio of local pitot to first 

ramp total pressure, PdP1 

Figure 16:Ramp boundary layer profiles, station 4.4 

.20 

PI 

0 

vc .14 

u) .12 $ 

r( 

m 

0 

I+ 

u .10 

.08 

.06 

5 .04 

8 

k 

vc 

01 

0 

4-j 

u) b" .02 

.oo 

Ratio of local pitot to first 

ramp total pressurel PdPl 

Figure 1'l:Ramp boundary layer profile, station 17.2 

between the near wall and centerline profiles is 

observed here. Figure 19 shows rake data near the 

end of the third ramp. All these rake profiles 

indicate well developed boundary layer on the ramp. 

6 
.20 

- 

cv .la \ 

c 

al 

u 

rc 

k 

.16 

m .14 

$ .12 

rl 

!Q 

0 

0 .10 

E .08 

.06 

4 -04 

rl 

0 

k 

VI 

al 

0 

Y 

.20 

.18 

.16 

.14 

.12 

in 

*r( 0 .02 

.oo 

- 

- 

- 

- 

- 

0 CENTERLINE 

0 7.5" OFF CENTERLINE 

.20 

7 .la 

5 

.16 

li) .14 

al 

W 

k 

; .12 

I+ 

Ratio of local pitot to first 

ramp total pressure, PdP! 

Figure 18: Ramp boundary layer profiles, station 

22.9 

- 

- 

- 

- 

- 

(d 

u 

0 

Pi 

E 

0 

k 

%I 

W 

B 

(d 

4J 

u) 

vi 

a 

.04 

.02 

9 

2 

.5 

Ratio of local pitot to first 

ramp total pressure, PdPl 

Figure 19: Ramp boundary layer profile, station 

30.5 

In Figures 20 - 28, pitot pressure (nondinensionalized to PI) profiles are shown for translating 

probes at various locations in the inlet. Distances 

from the local flow surfaces are nondimensionalized 

to h,l. Where probes are located downstream of bleed 

areas, both bleed and no hlced profiles are 

presented, with (with the exception of Figures 21 & 

22) circular symbols representing no bleed and 

squares denoting bleed configurations, respectively. 

Figure 20 shows survey data for a probe translating 

.10 

.oa 

.06 

.04 

.02 

.oo IIIIII 

.3 .4 .5 

Ratio of local pitot to first 

ramp total pressure, P,,/Pl 

Figure 20:Sidewall boundary layer profile, station 

10.0: bleed 

G .18 

*20 F 

.16 1 

a; 

W 

k : .12 I 

t -I 

0 

; .10 

Pi 

.08 

ICI .06 

E 

k 

Q 

u 

w 

2 .04 

u) 

a .02 

.oo .b 

0 CENTERLINE 

0 4" OFF CENTERLINE 0 7.5" OFF CENTERLINE 

Ratio of local pitot to first 

ramp total pressure, PdP I 

Figure 2l:Ramp boundary layer profiles, station 

35.5: no bleed 

out of the sidewall which is located at a station 

approximately 10 inches downstream of the start of 

the second ramp, and. 1.1 inches above the ramp 

surface (5.6 inches aft of the rakes shown in Figure 

16 and the first row of statics in Figure 12.) 

Figures 21 and 22 show sidewall probes which extend 

down from the fourth ramp at a location just 

downstrean of its leading edge. Figure 21 shows the 

no-bleed case, and Figure 22 the bleed case. The 

circles represent the probe located on the 

centerline; the squares indicate 
.20 

5 .18 

.16 

m .14 

s 

\ 

c 

01 

V 

yc 

k 

2 

rl 

(d 

0 

0 

E 

0 

k 

tec 

01 

u 

C 

m 

Y 

u) 

*r( 

P 

rl 

0 CENTERLINE 

0 4" OFF CENTERLINE 0 7.5'' OFF CENTERLINE 

Ratio of local pitot to first 

ramp total pressure' PdPl 

Picure 22: Ramp boundary layer profiles, station 

35.5: bleed 

.20 r 7 2 .18 

c 

- .16 

01 

0 

.14 +I 

k 

\ 

2 .12 

0 NOBLEED 

0 BLEED 

Ratio of local pitot to first 

ramp total pressure, PdPl 

Pimure 23: Sidewall boundary layer profile, station 

35.4 

the probe situated 4 inches off the centerline; and 

the diamonds denote the probe located 0.5 inches off 

the sidewall. The drop in pitot pressure 

approximately halfway up the survey indicates the 

location of the intersection of the oblique shock 

from the fourth ramp. Figure 23 shows the survey 

data for a probe which extends out from the sidewall 

at station 35.4 (corresponding to the three probes 

shown in the previous two figures and the third row 

- 

U 

e 

c 

al 

0 

rn 

rcI 

m 

rl 

(d 

0 

0 

rl 

E 

0 

k 

tec 

al 

(d 

Y 

m 

'4 

c1 

5 

: 

.20 

.ia 

.16 

.14 

.12 

.10 

.08 

.06 

.04 

.02 

.oo 

0 NOBLEED 

0 BLEED 

.i 

- 

.4 .5 

Ratio of local pitot to first 

ramp total pressure, PdPl 

Figure 24:Sidewall boundary layer profile, station 

52.3 

.20 r 7 

.18 \ 

c 

01 

0 

k 

.16 

2 .14 

.12 rl 

m 

0 

rl 

0 

k 

y.l 

01 

0 

Y 

m 

'4 

u .lo 

E .oa 

.06 

9 .04 

CI .02 

L 

.OO .o 

J 

0 NOBLEED 

0 BLEED 

1- 

.3 

Lf.--L-J 

.4 .5 

Ratio of local pitot to first 

ramp total pressure, PdPI 

Figure 2S:Sidewall boundary layer profile, station 

57.0 

of sidewall statics presented in Figure 12), located 

approxiaately 1.7 inches from the local ramp surface. 

The probe presented in Figure 24 is also located on 

the sidewall at station 52.3 (just upstream of the 

first row of statics shown in Figure 14), and about 

3.3 inches from the ramp. The sidewall probe from 

which the survey in Figure 25 is recorded is located 

approximately at the same station as the row of 

sidewall statics denoted by squares in Figure 14, and 

is 7 inches from the local ramp surface. 

8 
7 

cv .18 \ 

c 

P) 

0 

- .16 

2 .14 

m .12 

5 

g .to 

r( 

m 

- 

- 

- 

- 

- 

::: 1 

.16 

.20 c 0 NOBLEED 

0 BLEED 0 NOBLEED 

0 BLEED 

.14 

.12 

.10 

.08 

.06 

.04 

.02 

- 

- 

- 

- 

- 

- 

- 

- 

.2 .3 .4 .5 

Ratio of local pitot to first 

ramp total pressure, PdPl 

Ratio of local pitot to first 

ramp total pressure, PdPl 

Figure 26: Ramp boundary layer profile, station 

62.5 

Figure 28:Cowl boundary layer profile, station 68.5 

.20 

.18 

.16 

m .14 

m .12 

- 

- c 

\ 

c 

Q 

u 

w 

k 

3 

- 

- 

- 

NO BLEED 

BLEED 

.20 

5 .18 

.16 

m .14 

c 

\ 

c 

Q 

0 

w 

k 

2 .12 

4 

m 

0 

u -10 

E .08 

.06 

4 .04 

rl 

0 

k 

w 

Q 

u 

4J 

m b" .02 

0 NOBLEED 

0 BLEED 

rl 

m 

0 

0 .10 1 rl 

E .08 0 

k 

w 

u 

5 .04 4J 

m 

L 

.OO .o 

u .1 I I .oo 

.2 

Ratio of local pitot to first 

ramp total pressure, PdPl 

Ratio of local pitot to first 

ramp total pressure, PdPl 

Figure 27: Sidewall boundary layer profile, station Figure 29:Cowl corner rake profiles, station 59.6 

63.8 

represented by right-side-down triangles), on the 

cowl centerline, and it extends up into the flow 

towards the ramp. In general, bleed and no bleed 

profiles for the probes downstream of the cowl lip 

are similar, with the exception of those for the 

probe of Figure 27. 

In Figures 29 and 30, pitot pressure profiles 

are presented for cowl corner rakes located at 

stations 59.6 and 68.5 inches, respectively. These 

The probe of Figure 26 is located on the ramp 

centerline at station 62.5. The sidewall probe 

in Figure 27 is at approximately the same station, 

and 6 inches from the ramp surface. This station 

corresponds to the row of sidewall statics denoted 

by right-side-up triangles in Figure 14. The probe 

of Figure 28 is located slightly aft of these, at 

station 68.5 (near the row of sidewall statics 

9 
.20 

-18 

-16 

.14 

.12 

.10 

.08 

.06 

.04 

.02 

.oo 

0 NOBLEED 

0 BLEED 

Ratio of local pitot to first 

ramp total pressure, PdPl 

Figure 30: Cowl corner rake profiles, station 68.5 

rakes cut across the 3D corner vortical flow region 

predicted by CFD analyses. Station 69.8 is 

approximately 6.1 inches aft of the cowl lip. For 

the no bleed configuration, at this station, great 

changes in pressure are observed, with a region of 

medium pitot pressure very near the corner, which 

decreases rapidly, and then rises with a decrease 

near the end of the rake. With bleed, the profile 

shows an additional small rise in pitot pressure 

after the initial Pall-off, and the pressure near the 

end of the rake falls off sooner. At station 68.5, 

which is about 15 inches aft of the cowl lip, there 

is only a small "bump" in the profile near the 

corner. Bleed and no bleed profiles are similar at 

this station. The data for this profile would tend 

to indicate that the vortical flow action is not 

present at this location (either it has moved toward 

the center of the flow or disappeared at this 

station. ) 

CFD / DATA COMPARISONS 

Rose (Reference 6) calculated the flow field in 

the Mach 5 inlet using the Kumar explicit, timeaccurate implementation of MacCoraack's algorithm for 

solving the full Navier-Stokes equations. This 

calculation begins at the start of the second ramp; 

however, the incoming boundary layer (assumed to be 

turbulent) is accounted for by including in the 

incoming flow a boundary layer profile calculated for 

Mach 4.1 flow on the first ramp. The leading edge of 

the sidewalls begins at the start of the second ranp 

as well. The code assumes turbulent boundary layer 

on all surfaces from the start of their respective 

leading edges. For the comparisons shown here, bleed 

is removed in two places: 4% of the capture mass 

flow is removed from the ramp in the vicinity of the 

shoulder, and 3% is removed from the cowl through 

notches 3 inches wide on either side of the cowl and 

extending back 6 inches from the leading edge (not a 

configuration specifically tested.) 

Sidewall 

Sidewall vortex 

Cowl boundary layer 

Ramp boundary layer 

Cowl shock wave 

Ramp Cowl 

Symmetry plane, 

Figure 31:Rose FNS analysis: Mach number contours 

at a cross plane near the inlet shoulder 

Figure 31 shows Rose's predicted airflow 

conditions for one-half of a cross-section of the 

inlet. This cross-section is between the cowl lip 

and shoulder axial stations. The cowl surface is to 

the right and the ramp surface is to the left. The 

inlet centerline is indicated near the bottom of the 

section, and the sidewall is at the top. The Mach 

nuuber contours show a separation siuilar to that 

predicted by Benson's PNS analysis shown in Figure 4. 

despite the boundary layer bleed removed from the 

cowl. 

c o Experiment 

oa 

p0.25 

0.20 

0.15 

0.10 

0.05 

0.00 

L Shoulder-, 

-0.5 0.5 1.5 2.5 3.5 4.5 5.5 

Axial distance from beginning 

of second ramp, x/h,l 

Figure 32:ROSe FNS analysis / data comparison: 

ramp centerline static pressure 

Figures 32 and 33 show analysis / data 

comparisons for centerline static pressures on the 

ranp and cowl. Code / data agreement is reasonable, 

except on the ramp near the start of the ramp bleed 

region. The discrepancy may be due to differing 

amounts of bleed between the experiment and analysis. 

10 
- e 0.30 

0" '. 

'4 

w 0.00 

0 Experiment 

- Analysis 

Shoulder cow'lipl; 

I I 

I I 

I I 

I I 

I I 

I I 

OI 

I I I I I1 .I I 

-0.5 0.5 1.5 2.5 3.5 4.5 5.5 

Axial distance from beginning 

of second ramp, x/h,l 

Figure 33: Rose FNS analysis / data comparison: 

cowl centerline static pressure 

-I 

m 

V 

0 

-I 

E 

0 

k 

w 

(1 35 

(1.25 

0 15 

n nB 

0.00 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.e 

Ratio of local pitot to first 

ramp total pressure, PdPl 

Figure 34: Rose FNS analysis / data comparison: 

ramp centerline boundary layer profile, 

station -3.0 

Figures 34 - 36 show comparisons of boundary 

layer profiles on the inlet centerline at the end of 

the first, second and third ramps, respectively. 

Pitot pressures are nondimensionalized to Pi, and the 

distance from the ramp surface is divided by h,l. In 

Figures 35 and 36, the locations of the oblique shock 

intersections through the measurement plane can be 

observed in the analysis, which extends much further 

out from the ramp than does the actual rake. These 

three comparisons indicate that in flow areas where 

three-dimensional effects are negligible, the 

analysis agrees closely with the data. 

- 0.65 

0.55 

a 

0 

m 

W 0.45 

k 

rl 0.35 

m 

V 

0 

0 

k 

w 

0 Experiment 

- Analysis 

0 

0 

E 

I I I I I 

0.0 0.1 0.2 0.3 0.4 0.5 0.0 0.7 0.8 

Ratio of local pitot to first 

ramp total pressure, PdPi 

Figure 3S:Rose PNS analysis / data comparison: 

ramp boundary layer profile, station 17.2 

Ratio of local pitot to first 

ramp total pressure, PdPt 

Figure 36:Rose PNS analysis / data comparison: 

ramp boundary layer profile. station 30.5 

Figure 37 compares the analytical prediction of 

the corner rake profile at station 59.6 (6" aft of 

the cowl lip) to data from that rake. The circular 

symbols represent no bleed data, and the squares 

denote bleed data. The shape of the bleed profile is 

similar to the analysis; however, the pitot pressures 

for the analysis are generally lower. This 

difference may be due to the shorter length of the 

sidewalls (generating less boundary layer) than those 

of the actual inlet. 

11 
- 

0 c 0.35 

\ 

r: 

a; 0.30 

0 

m 

w 

k 0.25 

3 

ul 

7 0.20 

V 

0 

r( 

E 

0 

k 

u 0.10 

al 

0 9 0.05 

4J 

0.15 

Shoulder .-, 

Cowl lip-, I 

II 

I I 

1 

I 

- Analysis 

0 Experiment, no bleed 

0 Experiment, Meed 

A 

J 

I 

0 0 0 1 0.2 0 3 04 0.5 0 6 0.7 0.8 

Ratio of local pitot to first 

ramp total pressure, PdPl 

Figure 37: Rose FNS analysis / data comparison: 

Cowl corner rake profile, station 59.6 

Recently, analyses of the Mach 5 inlet have been 

completed using a three-dimensional, full NavierStokes time-marching code. The PARCSD code, 

described in Reference 7, solves the full threedimensional Reynolds-averaged Navier-Stokes equations 

in strong conservation form with the Beam and Warming 

approximate factorization. The turbulence model used 

in the code for this study is the Baldwin-Lomax 

model. The solution presented here assumes no bleed 

flow is removed from the inlet on any surface. 

Figure 38 shows Mach number contours in cross 

planes at selected axial stations. The figure shows 

the interaction of the shocks with the boundary layer 

on the sidewalls. The strong secondary flow set up 

by this shock/boundary layer interaction and the 

migration of the low energy fluid towards the center 

plane of the inlet as the flow approaches the throat 

region can be seen. 

Figure 38: PARC3D FNS analysis: Mach number 

contours at selected axial stations 

12 

A comparison of the predicted static pressures 

along the ramp centerline with the experimental data 

is shown in Figure 39. In this and the following 

figures, pressures are nondimensionalized to ramp 1 

total pressure, PI. and heights and distances are 

divided by cowl lip height, h,]. The agreement of the 

solution with the data is reasonable in the region 

upstream of the ramp bleed area. The comparison aft 

of that differs due to the absence of bleed in the 

analysis. Figure 40 shows a similar comparison for 

the cowl centerline. 

0 25C 

'a 0.2oc 

v2.. 

(0 005C 

k 

.rl 

w 

A Experiment 

__ Analysis 

I 'I 

I 

I 

I 

I 

-3-2-1 0 12 3 4 5 

Axial distance from beginning 

of second ramp, x/h,l 

Figure 3@:PARC3D analysis / data comparison: Harnp 

centerline static pressure 

Pitot pressure profiles are compared at locations 

along the ramp centerline in Figures 41 - 43. The 

agreement of the solution with the data is reasonably 

good. Figure 41 shows the profile for a rake near 

the end of the first ramp, and Figure 42 shows the 

profile corresponding to a rake located near the end 

of the second ramp. For both comparisons (though to 

a lesser extent for the first,) The computed maximum 

pitot pressure is slightly larger than the measured 

value. This disagreement is attributed to a slightly 

lower Mach number of the calculated flow approaching 

the rake. This lower Mach number is the result of 

a weak compression wave generated due to the boundary 

layer growth on the sidewall at the entrance of the 

inlet. This phenomenon does not seem to be present 

in the experimental data. The pitot pressure profile 

for a rake located near the end of the third ramp is 

shown in Figure 43 along with the experimental data. 

Once again for the reasons mentioned previously, the 

maximum computed values are slightly larger than the 

measured values in the two-dimensional region of the 

flow. 

Figure 44 shows a comparison of the predicted 

pitot pressure profile of the corner rake located 6 

inches aft of the cowl leading edge with bleed and no 
0.25C 

m4J 

a v) 0.050 

k 

4 

vc 

Shoulder 

0.ooc I I I I , 1 I1 I I I 

-3-2-1 0 12 3 4 5 

Axial distance from beginning 

of second ramp, x/h,l 

Figure 40: PAUC3D analysis / data cornparison: cowl 

centerline static pressure 

7 

V 

A 

\ 0.4 

c 

a, 

V 

m 

vI 5 0.3 

u) 

PI 

m 

0 

0 - 0.2 

E 

0 

k 

vI 

a, 

m 

w 

u) 

'rl 

2 0.1 

n 

0.0 

0 

A Experiment 

- Analysis 

1 

L 

L 

I I I I I I I 

I 0.02 0 04 0.06 0.08 0.10 0.12 0.1 

Ratio of local pitot to first 

ramp total pressure, PdPl 

centerline boundary layer profile, 

station -3.0 

Figure 41: PAUC3D analysis / data comparison: ramp 

0.6 

.-- 

0 

c 

\ 

0.5 

Q 

u 

(d 

vc 

k 0.4 

4 

(d 

; 0.2 

5 

4 

k 

vc 0.2 

Q 

0 

c 

(d t; 0.1 

'Ii 

n 

0.c 

00 0.04 0.08 0.12 0.16 

Ratio of local pitot to first 

ramp total pressure, PdPl 

Figure 42:PARC3D analysis / data comparison: ramp 

centerline boundary layer profile, 

station 17.2 

O.( 

- 

V 

c 2 O.! 

a, 

u 

m 

Wl 0.r 

k 

9 

v) 

r( u" 0.2 

0 

E 

0 

k 0.; vc 

(u 

0 

c 

w 

u) 

'4 

4 

0.1 

n 

0.c 

111111111111 

00 0.04 0.08 0 12 0.16 0.20 0.21 

Ratio of local pitot to first 

ramp total pressure, PdPl 

Figure 43:PARC3D analysis / data comparison: ramp 

centerline boundary layer profile, 

station 30.5 

13 
0.30 7 

0 

c 

\ 

c 

al 

0 

m 

y.c 

0.25 

!j 0.20 

u) 

w-l 

m 

u 

0 0.15 r( 

E 

0 

& 

al 

0 

C 

m 

v) 0.05 

0 

y.c 0.10 

w 

*I 

0.00 

- Analysis 

0 Experiment, no bleed 

0 Experiment, Meed 

- 

- 

- 

- 

- 

- 

- 

11 I I I I I I I I 

0.00 0.05 0.10 0.15 0.20 0.25 0.30 0.35 0.40 

Ratio of local pitot to first 

ramp total pressure, PdPl 

Figure 44: PARC3D analysis / data comparison: cowl 

corner rake profile, station 59.6 

bleed experimental data. The code predicts "bumps" 

in the profile similar to the data, but on a much 

different vertical scale. Discrepancies between the 

analysis and the experimental are currently being 

studied. An analysis which includes bleed 

representative of a configuration tested will be 

completed soon, and may provide closer comparisons to 

the data. 

CONCLUDING REMARKS 

The previous analyses would seem to indicate 

that the state of CFD has arrived to the point that, 

at least on a two-dimensional basis, it is doing a 

good job of predicting high speed flow fields. In 

areas of highly complex, three-dimensional flow, it 

is calculating flow fields which are somewhat similar 

to reality, since unusual trends ("burps" in pressure 

profiles on corner rakes) are being predicted which 

resemble the data in general shape, if not magnitude. 

Comparisons are, however, just getting underway, and 

the analyses presented here represent the current 

status of ongoing efforts to calibrate and refine the 

CFD codes. 

The Mach 5 inlet test has provided a set of data 

which is currently being employed by various 

organizations to calibrate high speed inlet analysis 

codes. The data is available to interested 

individuals, and nay be obtained by contacting the 

author. 

2. 

3. 

4. 

5. 

6. 

7. 

Coltrin. R.E: High-speed Inlet Research Program 

and Supporting Analyses. Presented at 

Aeropropulsion '87 Conference, NASA Lewis 

Research Center, Cleveland, Ohio, November 17- 

19, 1987. NASA CP-30003 

Perkins, E.W. ; Rose, W.C.; and Harie, G. : Design 

of a Mach 5 Inlet System Model, NASA CR-3830. 

August, 1984. 

Benson , T . J . : Three-dimensional Viscous 

Calculation of Flow in a Mach 5.0 Hypersonic 

Inlet, AIAA 86-1461, June, 1986 

Anderson, B.H.: Three-dimensional Viscous Design 

Methodology for Advanced Technology Aircraft 

Supersonic Inlet Systems. NASA TM-83558, 1984. 

Rose, W.C., and Perkins. E.W.: Innovative 

Boundary Layer Control Methods in High Speed 

Inlet System. Final Report, Phase I SBIR 

Contract No. NAS3-25408, NASA Lewis Research 

Center, September, 1988. 

Cooper, G.K., Jordan, J.L., and Phares. W.J.: 

Analysis Tool for Application to Ground Testing 

of Highly Underexpanded Nozzles. AIAA 87-2015, 

1987. 

REFERENCES 

1. Watts, J.D., et. al.: Mach 5 Cruise Aircraft 

Research. NASA CP-2398. Volume 11. 1985. 

14 
National Aeronautics and I Space Administration 

I 

Report Documentation Page 

[ 12. Sponsoring Agency Name and Address 

National Aeronautics and Space Administration 

Washington, D.C. 20546-0001 

2. Government Accession No. NASA TM-102317 

AIAA-89-2355 

1. Report No. 

4. Title and Subtitle 

Mach 5 Inlet CFD and Experimental Results 

I 

7. Author(@ 

Lois J. Weir, D.R. Reddy, and George D. Rupp 

17. Key Words (Suggested by Author(s)) 18. Distribution Statement 

I 9. Performing Organization Name and Address 

National Aeronautics and Space Administration 

Lewis Research Center 

Cleveland, Ohio 44135-3191 

19. Security Classif. (of this report) 20. Security Classif. (of this page) 21. No of pages 

Unclassified Unclassified 16 

22. Price’ 

A03 

3. Recipient’s Catalog No. 

5. Report Date 

6. Performing Organization Code 

8. Performing Organization Report No. 

E-501 1 

10. Work Unit No. 

505-80-2 1 

11. Contract or Grant No. 

13. Type of Report and Period Covered 

Technical Memorandum 

14. Sponsoring Agency Code 

15. Supplementary Notes 

Prepared for the 25th Joint Propulsion Conference cosponsored by the AIAA, ASME, SAE, and ASEE, 

Monterey, California, July 10-12, 1989. Lois J. Weir, NASA Lewis Research Center; D.R. Reddy and George 

D. Rupp, Sverdrup Technology, Inc., NASA Lewis Research Center Group, Cleveland, Ohio 44135. 

16. Abstract 

An experimental research program was conducted in the NASA Lewis Research Center 10 ft. by 10 ft. supersonic 

wind tunnel. The two-dimensional inlet model was designed to study the Mach 3.0 to 5.0 speed range for an 

“over-under” turbojet plus ramjet propulsion system. The model was extensively instrumented to provide both 

analytical code validation data as well as inlet performance information. Support studies for the program include 

flow field predictions with both three-dimensional parabolized Navier-Stokes (PNS) and three-dimensional full 

Navier-Stokes (FNS) anaytical codes. Analytical predictions and experimental results are compared. 

Inlet 

Computational Fluid Dynamics 

Hypersonic propulsion 

Unclassified - Unlimited 

Subject Category 07