NASA Technical MemQrandum_ !0311 t

%o .,

Computationa! _Modeling and ......

Validation for Hypersonic Inlets

Louis A. Povinelli ....

Lewis Research Center

Cleveland, Ohio

(NA_A-TM-I0311I) C_MPUTATIONAL MQOELING AND

VALI_AIION FOR HYPERSONIC INLEIS (NASA)

ii D CSCL 20D

N90-22011

unclas

G3/34 0280100

Prepared for the

75th Symposium on Hypersonic Combined Cycle Propulsion

' sponsored by_Propuls_on and Energetics Panei of AGARD

MM!id, Spain, May 28--June 1, 1990

[W A

COMPUTATIONAL MODELING AND VALIDATION FOR HYPERSONIC INLETS

Louis A. Povinelli

National Aeronautics and Space Administration

Lewis Research Center

21000 Brookpark Road

Cleveland, Ohio 44135 U.S.A.

SUMMARY

Hypersonic inlet research activity at NASA is reviewed. The basis for the paper

is the experimental tests performed with three inlets: the NASA Lewis Research Center

Mach 5, the McDonnell Douglas Mach 12, and the NASA Langley Mach 18. Both threedimensional PNS and NS codes have been used to compute the flow within the three inlets.

Modeling assumptions in the codes involve the turbulence model, the nature of the

boundary layer, shock wave-boundary layer interaction, and the flow spilled to the outside of the inlet. Use of the codes in conjunction with the experimental data are helping to develop a clearer understanding of the inlet flow physics and to focus on the

modeling improvements required in order to arrive at validated codes.

INTRODUCTION

This paper presents a review of the hypersonic inlet activities at NASA. Generally,

the NASA work has covered the range from Mach 5 to 18. The research has been of a computational and experimental nature with a two-fold objective: (I) to develop an improved

understanding of the physics and chemistry of inlet flow fields and {2) to validate the

numerical codes used for high-speed inlets. The objectives result from the basic philosophy developed for the National Aerospace Plane project; namely, that numerical codes can

be validated using ground test data and extrapolated to higher velocities. Since the

ground experimental data base is limited to Mach 8 and lower, extrapolation of the computational methods must be made from Mash 8 to the Mach number, approximately 16, where

airbreathing propulsion terminates.

A number of codes have been used to date within NASA for inlet flows. Code development and modifications have taken place over the last several years and the emphasis

today is on the application to various inlets that have been tested. Those inlets

include the Mach 5 inlet at NASA Lewis Research Center, a General Dynamics Mach 12 configuration and a Mach 18 NASA Langley inlet. Only a limited amount of data is available

for each configuration.

The numerical methods used have included a PNS code, the PARC NS code, the SCRAM3D

and CFL3D. Generally, the codes employ a Baldwin-Lomax turbulence model. Assumptions

are made regarding the state of the boundary layer and spillage has generally not been

computed. Thick upstream boundary layers are computed in some cases having a forward

extenslon surface. In this paper a comparison of the experimental and computational

results will be reviewed.

RESULTS

Mach 12 Generic Inlet

The simple rectangular inlet configuration shown in Fig. 1 was tested at Mach 12.26.

A flat plate of 30-in. length preceded the entrance to the inlet in order to simulate

the boundary layer growth on the forebody of a hypersonic aircraft. Compression wedges

form the top and bottom walls of the inlet and the contraction ratio was equal to 5.

Swept sidewalls which connect the upper and lower walls prevent compressed flow from

spilling over the inlet sides.

Computations were made with a three-dimensional PNS LBI implicit scheme (Ref. I)

with grids of 80 by 60 by 750 on a Cray X-MP. This solver includes real gas effects

(Ref. 2) as well as dissociation and ionization modeling (Ref. 3). For this experiment,

however, the inlet air was only heated sufficiently to avoid condensation, and the real

gas modeling was not required. The issues that are of importance in this computation

are the assumptions regarding the state of the boundary layer, the turbulence model,

spillage of flow around the sideplates and shock boundary layer interaction. For the

PNS computation _t was assumed that the boundary _ayer was turbulent starting on the

leading edge of the flat plate, the cowl leading edge and the sidewall leading edges.

The turbulence model used was a Baldwin-Lomax model and spillage was not considered.

Modeling of the shock boundary layer interaction involved the use of a flare approximation in order to allow the PNS to march through the region of flow separation. The

results of the PNS solution are shown in Fig. 2. Contour plots of constant Mach number

within the inlet are shown. The concentration of lines near the walls indicate the

boundary layers, while concentrated contours in the freestream indicate shock wave locations. The flow features seen are boundary layer buildup on the flat plate followed by

thickening on the sidewalls and ramp surface. Shocks generated by the compression wedges

are seen as horizontal lines, and the sidewall shocks are vertical lines.
Basically, the important physics occurring are that the low energyflow in the sidewall boundarylayer has beensweptup the sidewall by the rampshock, and then down the

sidewall by the cowl shock. Further downstream, the shock waves cross and are distorted

by interaction with the sidewall boundary layers and the expansion fan on the ramp surface. Additional complex interactions then occur as the flow moves downstream. The PNS

solution fails when the ramp shock wave reflects from the cowl and strikes the ramp surface, resulting in large corner separation of the low energy flow.

An alternate view of the three-dimensional flow is obtained with sidewall particle

tracing (Fig. 4). Interaction of the ramp and cowl shocks with the sidewall boundary

layer causes the particles to converge near the shock interaction point. The particles

are then displaced due to the vortex motion. Flow migration details are evident in this

computational simulation. As a sidenote, since the vortex persists downstream, it has

been proposed that enhanced fuel mixing could occur with judicious injector locations

downstream (Ref. 4).

Navier-Stokes computations have also been carried out for the generic inlet at NASA

Langley with CFL3D (Ref. 5). In this case, the boundary layers were assumed turbulent

on all surfaces from the leading edges. The turbulence model used was a Baldwin-Lomax

model and spillage over the sideplates was not considered. In the vicinity of the shock

boundary no special modeling was employed. Figure 5(a) shows the pressure distributions

for the ramp and centerline cowl surfaces. Figure 5(b) shows the side plane distributions. Comparison of the CFL3D results and the experimental data show good agreement,

particularly along the centerline where shock locations appear to be well resolved by

the code. The viscous interactions occurring along the side plane are not accurately

resolved. There is a significant underprediction of the pressure on the ramp side

(Fig. 5(b)). In addition, CFL3D was used to compute the heat transfer on the ramp and

cowl surfaces (Figs. 6(a) and (b)). The experimental peak heat fluxes are underpredicted for the ramp centerline but well predicted for the cowl surface.

For the ramp and cowl side planes (Fig. 6(b)), the peak prediction is lower on the

ramp whereas the cowl side prediction is not qualitatively correct. AgaiN, strong V{s _

cous effects are predominating along the side walls of the inlet in agreement with the

complex behavior shown in Figs. 2 to 4. Further analysis of the Mach 12 inlet is underway at the NASA Centers and industry.

Mach 5 Inlet

A rectangular mixed compression inlet designed for Mach 5 operation and tested at

NASA Lewis is shown in Figs. 7(a) and (b) (Ref. 6). A series of ramps generate oblique

shock waves external to the cowl. An oblique shock from the cowl leading edge reflects

from the ramp surface and terminates in a normal shock downstream of the inlet throat.

Operation in the wind tunnel was such that a Mach number of 4.1 occurred on the first

ramp. The inlet incorporates variable geometry with collapsible ramp and variable bleed

exits on the cowl, sidewalls and ramps. Bleed of 0.5 percent was removed on the ramp

upstream of the shoulder. Additional bleed from the cowl and sidewalls was approximately

8.8 percent of the capture mass flow. Figure 8 shows the location of pressure rakes and

probes in the model. A 0.5-in. strip of grit was applied near the leading edges of the

ramp and sidewall to ensure that a fully turbulent boundary layer was ingested by the

inlet.

Navier-Stokes computations were carried out using the PARC3D solver (Ref. 7) on the

NAS Cray 2. Grid sizes of 151 by 81 by 41 were used with hyperbolic packing so that the

first point was at a y+ of 2. Bleed was simulated by imposing a constant mass flux

through the porous bleed surfaces based on the experimental data. The boundary layer was

assumed to be turbulent throughout, and the turbulence model was that of Baldwin-Lomax.

Flow spillage over the sideplates was also not computed in this case.

The computed ramp pressure results are compared with experimental data in Fig. 9.

The agreement of the computations with the data is very good throughout the computed

length of the inlet. Figure i0 shows the comparison for the cowl pressure distribution.

The d_sagreement of the results at an x/h of 4.2 is believed to be due to the fact

that one of the translating probe assemblies is located in the same region where the

four static pressure taps are located. Because the retracting probe does not completely

retract into the wall, additional shocks are generated which biased the data. Pitot

pressure profiles were compared with data at various locations along the inlet. Figure Ii shows the pressure profile from rake 3 which was located on the centerline and in

the region of the second ramp (Fig. 8). The agreement of data and computation is very

good. Along the sidewall, however, the agreement is much poorer, as shown in Fig. 12

for rake 7. The corner effects are not being adequately simulated. An improved turbulence model may improve the comparison in these corner regions. Figure 13 shows the

pitot pressure comparison for rake 10 mounted at 45 ° from the corner of the cowl and

sidewall at station 59.6 from the start of the inlet. This region of the flow is dominated by low energy vortical flow as seen in Fig. 14. Large variations in the pitot

pressure are seen as one moves from the corner into the stream. Measurements in these

regions are also very difficult.

The Mach 5 inlet was also analyzed using the SCRAM3D Navier-Stokes code by Rose

(Ref. 8). A Baldwin-Lomax turbulence was used, assuming turbulent boundary layers.

These results also reveal strong glancing shock wave-boundary layer interaction leading

to large regions of low momentum flow on the sidewalls. Rose carried out a number of

numerical experiments to control the vortex phenomena in the corner regions. Figure 16
showsthe baseline or no control case, followed by cowl cutback, cowl bleed and removal

of a part of the sidewall. Thesemodifications weremadenear the inlet rampshoulder.

It maybe seenthat these modifications were ineffective in eliminating the vortex

region. Evenwith the cutback sidewall, the low momentumfluid exists along the entire

sidewall. Someattenuation is seenalong the cowl surface for that case. It is evident,

however,that the shock-boundarylayer physics within a rectangular shapedinlet will

lead to pressure losses in the corner regions. However,if these regions can be utilized

in an "integrated design approach"(Ref. 4), then combustor/nozzledesign maybenefit

substantially. Further computationsof the Mach5 inlet with improvedsimulation of the

bleed zonesis underway,as well as further analysis of the test data.

Mach18Sidewall CompressionInlet

A sidewall compressioninlet hasbeendesignedand tested at NASALangleyby

Trexler. As seen in Fig. 17, the compressionoccurson the sideplates. Theexperiments

were run at an entranceMachnumberof 18to 22with andwithout a flat plate upstreamto

simulate fuselage boundarylayer buildup. Theentering boundarylayer wasapproximately

one-third the cowl height. Computationsweremadeby Rose(Ref. 9) using the SCRAM3D

code. Laminarboundarylayers wereassumedon cowl andsidewalls, and a Baldwin-Lomax turbulence modelwasused.

Figure 18 showsthe Machnumbercontours along the vertical centerplane for an entranceMachnumberof 18.1. Thecontraction ratio was4 andthe cowl leading edgeis

located at the entrance to the constant area section. All of the convergenceoccurs

along the sidewalls, which generatea pair of shockwavesthat intersect on the vertical

centerplane. A large pressure rise is felt on the rampsurface. Further downstream,the shocksinteract with the sidewall boundarylayers and reflect andintersect again on the

centerplane at the indicated position. A further rise in pressure causesrampboundary

layer separation.

TheMachnumbercontours on the horizontal centerplane are shownin Fig. 19. The

intersecting sidewall shocksand the intersecting reflected shocksare visible. Since

the sidewall shocksstrike the sidewall well upstreamof the shoulder, shockcancellation is clearly not achieved. The reflected shockwaves,however,are seento cancel at

the shoulder. Thestrong viscous interaction effects are very evident at these flow conditions.

Calculated Machnumbercontours are shownin Fig. 20 for both the horizontal andthe

vertical center planes with an entranceboundarylayer. Theentrance plate reducesthe Machnumberfrom the entrancevalue to about Mach12. Separation of the boundary layer

on the ramp, caused by the sidewall shock waves, causes a large upstream influence. As

the ramp boundary layer thickens, an oblique shock occurs reducing the inlet flow to

Mach 8. Sidewall shocks and their intersection are seen in the horizontal centerplane.

The sidewall shock wave angle is substantially increased due to the reduced Mach number

entering the inlet. The ramp shock falls outside the cowl leading edge. Figure 21 shows

a comparison of the experimental and computed surface pressure distributions on the ramp

centerline, £or the case where the cowl is moved forward, in this particular comparison,

the numerical code yields results which are higher than the measured data and also rises

faster than measured. Further data analysis and comparisons are underway at the present

time, which will lead to a more complete understanding of the flow in this class of

inlets.

CONCLUSIONS

Through the use of a variety of numerical simulations and experiments, the basic

flow features within rectangular hypersonic inlets are becoming better understood. Fast

running PNS solvers in combination with much longer running but more sophisticated

Navier-Stokes codes are providing a clearer picture of shock structure and boundary

layer behavior in inlets. Clearly, the flow fields are highly three-dimensional, viscously dominated and contain significant flow separations. Shock wave-boundary layer

interactions persist do_ to the throat and beyond. As the propulsion community moves

towards the validation of these codes, a number of issues still remain which will impede

the application of the methods for the design of hypersonic inlets. Perhaps the first

concern is that regarding the nature of the boundary layer within the inlet. Although

attempts are made to ensure the presence of turbulent layers, for example, questions

still remain regarding the existence of transitional layers. Installation of hot film

gauges on the walls would provide the numerical analyst with the proper information to

use within his computer code; be it laminar, transitional or turbulent. An additional

issue is concerned with turbulence modeling and the ability of the Baldwin-Lomax model

in regions of glancing shock wave-boundary layer interaction. Current turbulence models

appear to yield good qualitative flow characteristics, but may be inadequate for quantitative predictions. Alternate models are needed. A third concern deals with the necessity of including the zone outside of the inlet in order to provide proper boundary

conditions for the computation. In spite of these concerns, the understanding and agreement (i.e., on centerline) are very significant. The Mach 5 inlet presents the same concerns but with the complication of bleed flow on all four surfaces. Modeling issues

related to turbulence, boundary layer transition and spillage are made more complex with

a distributed mass flux boundary condition. Again, it is remarkable that such good

agreement, with limited data, was achieved for that inlet. A great deal more effort will

be required to analyze the test data and develop the proper modeling for a bulk of the

data.
Thesidewall compressionpresents a unique approachto inlet design. Computations of the three-dimensional flow field and associatedshockstructure provide us with a

great deal of physics. Thelimited data provides a basis for an improvementin current

modeling. Use of the experimental data and empirically derived correlations may serve

as a basis to produce improved viscous modeling. It is important to point out that in

all of these inlet tests, the tunnel flow was sufficiently low to avoid real gas effects.

Hence, the comparisons presented in the paper are only aerodynamic in nature. More significant modeling issues will surface as test data is acquired in high-enthalpy flows.

Finally, the need to analyze variable geometry with transient disturbances will provide

a significant challenge.

REFERENCES

i. Reddy, D.R., Smith, G.E., Liou, M.-F., and Benson, T.J., "Three Dimensional Viscous

Analysis of a Hypersonic Inlet," AIAA Paper 89-0004, Jan. 1989. (Also, NASA TM-!01474,

1989.)

2. Liou, M.-F., "Three Dimensional PNS Solutions of Hypersonic Internal Flows With

Equilibrium Chemistry," AIAA Paper 89-0002, Jan. 1989.

3. Yu, S.-T., McBride, B.J., Hsieh, K.-C., and Shuen, J.-S., "Numerical Simulation of

Hypersonic Inlet Flows with Equilibrium or Finite Rate Chemistry," AIAA Paper 88-0273,

Jan. 1988.

4. Povinelli, L.A., "Advanced Computational Techniques for Hypersonic Propulsion," NASA

TM-102005, 1989. Presented at the Ninth International Symposium on Air Breathing

Engines, Athens, Greece, Sept. 4-9, 1989.

5. Diley, A., Eppard, W., and Switzer, G., "Zonal Analysis of Two High Speed Inlets,"

Presented at the CFD Symposium on Aeropropulsion, NASA Lewis Research Center,

Cleveland, OH, April 24-26, 1990.

6. Weir, L.J., Reddy, D.R., and Rupp, G.D., "Mach 5 Inlet CFD and Experimental Results,"

AIAA Paper 89-2355, July 1989.

7. Reddy, D.R., Benson, T.J., and Weir, L.J., "Comparison of 3-D Viscous Flow Computations of Mach 5 Inlet with Experimental Data," AIAA Paper 90-0600, Jan. 1990.

8. Rose, W,C., and Perkins, E.W., "Innovative Boundary Layer Control Methods in High

Speed Inlet Systems-Final Report," Contract NAS3-25408, Sept. 9, 1988 (NASA CR in

publication).

9. Rose, W.C., Perkins, E.W., and Benze, D.P., NASA Ames Research Review, Dec. 1989.

COWL

CONTRACTION RATIO = 5 ____ 24 ir_

Figure 1. - Generic inlet (ref. 5).

Figure 2. - Mach number contours, M = 12.25 (re[ 1).
250_

150

loo

5O

_- 0

t',r

-50 --

no..

LU

_6O--

rr

80

Figure 3. - Mach number contours viewed from aft (ref. 1).

Figure 4. - Sidewall particle tracing, M = 12.25 (ref. _}.

RAMPCENTERLINE

200-- O

O EXPERIMENTAL

DATA I

CFL3D O_

6O

4O

2O

0

2O

COWL CENTERLINE

(a) Ramp cenlerline and cowl centerline.

RAMP SIDE PLANE 0

0

0

o -

30 40 50 6O 30

DISTANCE DOWNSTREAM, in.

(b) Ramp side plane and cowl side plane.

Figure 5. - Pressure comparisons {ref. 5).

_L____j

COWL SiDE PLANE

0

40 50 60
o,i

(30

LU

u.

@J

Z

<

rr

w

I

c)

<

n60

50

40

30

20 --

10 --

0 --

-10

50 --

40 --

30 --

20 --

10

0

2O

"6

RAMP CENTERLINE

O EXPERIMENTAL

DATA

CFL3D O

O

O

Oo

1 I l I

COWL CENTERLINE

I 1

(a) Ramp centerline and cowl centerline.

RAMP SIDE PLANE

O

30 40 50

\

-40

I I I

-3 -2

E

COWL SIDE PLANE

m

60 30 40 50

DISTANCE DOWNSTREAM, in.

(b) Ramp side plane and cowl side plane.

Figure 6. - Heat transfer comparisons (ref. 5).

STATION, in.

-20 0 20 40 60

I I I I I I I I I

X/hd

-1 0 1 2 3 4

I I t I I

oI ..............M.,o., = 4.1_////_'j/'_

I M. =s.o ...... _" -'.L.'_./3

t _ 3.7 --1 ......... L'.._'--_.: 1 ._"_(-J- _

1 L ........ 3.4 "J ---:t,,_¢=_=;_,¢.o _= go

(a) Inlet compression system to ramp shoulder,

I

6O

LEFT

STA - 47.89 SIDEWALL

STA 0.01 I _-"TRANSLATING

I , _ _ !STA2020 PROBES

I ,' ' '

! v , ', ,

_;ps'_ /_,_STA '53.48 i

BOUNDARY,'" ! _ U._" _ STA'67.25

_>- TRANSLATIN_

OORNERRA_E_I__/OowLPROBES

(b) Instrumentation upstream of inlet shoulder.

Figure 7. - (ref. 6).
r-RAKE13

I

I_-'_! F- PROBE 1"0

PROBE 11 ---__ RAKE 10

I.A \ "_PROBE5

RAKE 6

E 7 ,

\J\

RAKE 3 "--" _,_ _

RAKE 1

Figure 8. - Location of fixed rakes and translating probes used for pitot

pressure measurement (ref. 7).

.20 I'

.t5

J_

O

o

-_ .10

-r"

o_

uJ

"T"

.05

A EXPERIMENT

ANALYSIS l

2 4 6 8 10 12 14

PITOT PRESSURE, Pp/pinf

Figure 11. - Rake 3 pitot pressure profile (ref. 7).

E

u]

n"

t_LU

(3

Q.

20 --

15

10

5

0

-2

_ O

EXPERIMENT

ANALYSIS } CENTERUNE

EXPERIMENT

) 7.5 in, FROM CENTERLINE ANALYSIS

/k

-1 0 1 2 3 4 5

DISTANCE ALONG THE RAMP, X/COWL ht

Figure g. - Ramp static pressure distr_ution (ref. 7).

.20 A EXPERIMENT \

-- ANALYSIS ,t__

-_, .15

g A

0 4 8 12 16

PITOT PRESSURE, Pp/pinf

Figure 12. - Rake 7 pitot pressure profile (ref. 7).

1

18

2O

18

16

. 14

uJ

rr

03u.I

" 6

-- A

-- O

EXPERIMENT

ANALYSIS } CENTERLINE

EXPERIMENT

ANALYSIS 7.5 in. FROM CENTERLINE

4

2

o

3.0

Z_

-l---_' I I I I _

3.5 4.0 4.5 5.0 5.5

DISTANCE ALONG THE COWL, X/COWL ht

Figure 10,. Cowl static pressure distribution (ref. 7).

J

6.0

.35

.3O

.25

.E

.20

O

I- .15

W

-r

.10

.O5

/k EXPERIMENT /

ANALYSIS

,,k

_ A

_ A

A

-- /k /_

I

5 10 15 20 25

PITOT PRESSURE, Pp/pinf

Figure 13. - Rake 10 pitot pressure profile (ref, 7).

I

3O
PITOTPRESSURE RATIO

CONTOUR LEVELS

0.00000 2.50000

5.00000

7.50000 I_ _ RAKE 10

10.00000 12.50000 t \

t5,ooooo \

17.50000

20.00000

22.50000

25.00000

27.50000

30.00000

32.50000

35.OOOOO

37.50000 \

40.00000

42.50000

FIGURE 14. - Pitot pressure contour at rake 10 and rake 13 (ref. 7).

,,r SYMMETRY PLANE

COWL SHOCK WAVE --

COWL BOUNDARY

LAYER

_- RAMP

BOUNDARY

LAYER

SIDEWALL VORTEXSIDEWALL

Figure 15. - Orientation of crossfiow plane 150 located near the

ramp shoulder showing plane of symmetry, sidewall ramp and

cowl surfaces (ref. 8).

I

NO CONTROL CUTBACK COWL COWL BLEED CUTBACK SIDEWALL

Figure 16. - Comparison of effect of various control methods on mach number contours

near ramp shoulder (ref. 8).

FREESTREAM CASE

Pt = 2000 psi Pt = 300 psi

Tt = 520 °R Tt = 520 °R .075

30deg _ J"" '.050

HORIZONTAL

CENTERPLANE _" _

" -.025

,,," L_ RAMP

/

FLOW z._VERTICAL

CENTERPLANE -.050

ALL LAMINAR BOUNDARY LAYER

Figure 17. - Sidewall compre_ion inlet schematic (ref, g).

M_=18.1 6°SIDEWALLS GCR=4.0 COWLAFT

VERTICALCENTER PLANE

F- SEPARATION

t SHOCK RAMP

FLOW \ '_ 'o

ti I COWL

INTERSECTING t L_ I

._j NTERSECTING

_._SIDEWALL SHOCKS REFLECTED

SHOCKS

EXPANDED VERTICAL SCALE

1 t I I I 1 1

-.05 0 .05 .10 .15 .20 .25 .30

x

Figure 18. - Mach number contours for freestream case (ref. 9).

I

.35
.050 --

.025 -

0 --

-.025 ]

-.05 0

M. = 18,1 6° SIDEWALLS GCR = 4.0 COWL AFT

HORIZONTAL CENTER PLANE

EXPANDED LATERAL SCALE

SIDEWALL //--SIDEWALL

SHOCKS/_ /'I/ SEPARATIONS

// I /

_,_,;__ ," / -._

CONSTANT AREA THROAT

I I I I I I I

.05 .10 ,15 .20 .25 .30 .35

X

Figure 19, - Mach number contours for a representative freestream

case in the horizontal center plane (ref, 9).

.200

.150

.100

,050

.025

-.025

M_= 17.7 6° SIDEWALLS GCR = 5.0 COWL AFT (100%)

__ COWL__.__

_2_'_-, RAMP

/// _ //

/ L.. INLET RAMP

,," LEADING EDGE

z_ ENTERING TURBULENT

BOUNDARY LAYER

HORIZONTAL CENTER PLANE

I I 1 I 1 I I I

1,2 1.3 1,4 1.5 1.6 1.7 t .8 1.9 2.0

X

Figure 20. - Mach number contours for entering boundary layer case

(ref. 9).

J

Q.

1000

750

500

250 --

0

1.2

M_ = 17.7 60 SIDEWALLS GCR = 5.0 WEDGE CASE

RAMP CENTERLINE

COWL FWD (40%) BALDWlN-LOMAX

TURBULENCE MODEL

1.3 1.4 1.5 1.6 1.7 t .8 1.9 2.0

x, m

Figure 21. - Comparison of experimental and numerical

surface pressure distributions (ref, 9).
National Aeronautics and

Space Adminisltalion

1. Report No.

NASA TM-103111

Report Documentation Page

2. Government Accession No. 3. Reciplent's Catalog No.

4. Title and Subtitle

Computational Modeling and Validation for Hypersonic Inlets

7. Author(s)

Louis A. Povinelli

9. Performing Organization Name and Address

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio 44135-3191

12. Sponsoring Agency Name and Address

National Aeronautics and Space Administration

Washington, D.C. 20546-0001

5. Report Date

6. Performing Organization Code

8. Performing Organization Report No.

E-5435

lO. Work Unit No.

505-62-21

11. Contract or Grant No.

13. Type of Reportand PeriodCovered

Technical Memorandum

14, Sponsoring Agency Code

15. Supplementary Notes

Prepared for the 75th Symposium on Hypersonic Combined Cycle Propulsion sponsored by the Propulsion and

Energetics Panel of AGARD, Madrid, Spain, May 28--June 1, 1990.

16. Abstract

Hypersonic inlet research activity at NASA is reviewed. The basis for the paper is the experimental tests performed

with three inlets: the NASA Lewis Research Center Mach 5, the McDonnell.Douglas Mach 12, and the NASA

Langley Mach 18. Both three-dimensional PNS and NS codes h_iv_ been used to compute the flow within the

three inlets. Modeling assumptions in the codes involve the turbulence model, the nature of the boundary layer,

shock wave-boundary layer interaction, and the flow spilled to the outside of the inlet. Use of the codes in

conjunction with the experimental data are helping to develop a clearer understanding of the inlet flow physics

and to focus on the modeling improvements required in order to arrive at validated codes.

18, Distribution Statement

Unclassified - Unlimited

Subject Category 34

22. Price"

A03

17. Key Words (Suggested by Author(s))

CFD

Hypersonic propulsion

Propulsion CFD

High speed inlets

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

Unclassified Unclassified 10

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