A procedure is presented, as well as some results, to calculate the flow over the winged orbiter. This necessitates the use of two computer codes. A parabolized marching Navier-Stokes code is used to obtain the solution up to the bow shock-wing shock interaction region and for the region after the interaction. An unsteady Navier-Stokes code is to be used in the region of the shock interaction. Only resuls for the marching code are presented. For the flow conditions calculated, M infinity = 7.9, alpha = 25 deg, T(wall) = 540 R, Re(L) = 60728 per inch, laminar or turbulent, the PNS code was marched up to an X/L = 0.7 which is where the bow shock-wing shock interaction region occurs

Buning, P. G.

Chaussee, D. S.

Rizk, Y. M.

English

NASA Technical Reports Server

NASA Technical Memorandum 85977 
NASA-TM-85977 19840022156 
Viscous Computation of a Space 
Shuttle Flow Field 
Denny S. Chaussee, Yehia M. Rizk and 
Pieter G. Buning 
June 1984 
NI\5/\ 
National Aeronautics and 
Space Administration 
LIBRARY COP~ 
. ~?'" '984 
LANGLEY RESEARCf-' ~fI'lTER 
LII~RA~Y, NAS;' 
HAMPlO"" VIRGINIA 
111111111111111111111111111111111111111111111 
NF00832 
https://ntrs.nasa.gov/search.jsp?R=19840022156 2020-03-20T22:44:17+00:00Z
NASA Technical Memorandum 85977 
Viscous Computation of a Space 
Shuttle Flow Field 
Denny S Chaussee 
Pleter G Bunmg, Ames Research Center, Moffett Field, California 
Yehla M RIZk, Informatics General Corporation, 1121 San Antonio Road, 
Palo Alto, California 
NI\S/\ 
National Aeronautics and 
Space Administration 
Ames Research Center 
Moffett Field, CalifOrnia 94035 
I. INTRODUCTION 
VISCOUS COMPUTATION OF A SPACE SHUTTLE FLOW FIELD 
D. S. Chaussee, Y. M. Rizk, and P. G. Buning 
NASA Ames Research Center 
Moffett Field, CA 94035 USA 
Recent research efforts [1-3] have confl.rmed the ab~h.ty of the "parabo11Zed" 
Navier-Stokes (PNS) codes to predict accurately and rap~d1y the aerothermodynamics 
of the actual Space Shuttle Orb~ter up to an axial station that is 50% of the overall 
length. Th~s corresponds to a 10cat~on where the strake ends and the 45° swept w~ng 
beg~ns. In the past, the geometry of the Orbiter usually has been modified [4-7] 
such that a solution over the complete body was possible. These modifications take 
the form of chang~ng the sweep angle of the w~ng, remov~ng the canopy, and altering 
the lee of the body so that the cross sect~on ~s composed of two elliptical shapes. 
In one ~nstance [8], an inviscid solut10n was obta1ned for the complete Orbiter. In 
order to perform th1s calculation, a "fix" had to be used 1n the vicinity of the bow-
shock/w1ng-shock interact10n reg10n. Phys1cally, what occurs is a region of embedded 
subson1c flow oW1ng to the bow-shock/wing-shock interaction which causes the result-
1ng coalescing shock wave to be more normal to the flow for a short streamwise dis-
tance. Since the above 1nv1sc1d procedure was a marching code, 1t failed when the 
Mach number 1n the marching direct10n became subsonic. 
A V1SCOUS numer1cal procedure 1S descr1bed, to compute the flow over the Shuttle. 
Results are presented that demonstrate the capab~lity of the method. Obtainment of 
these results requires the use of two computer codes. A PNS code [9,10] is used to 
'obta1n the solut10n up to the bow-shock/wing-shock interact10n region, and an unsteady 
continuation code is used for the region after the shock 1nteraction. The unsteady 
Navier-Stokes code [11] 1S also used to obtain the blunt-body starting solution. Only 
results from the march1ng code w1ll be presented. For the flow condit10ns calculated, 
that is, Moo = 7.9, a = 25°, Twall = 540"R, ReL = 60,728/in., laminar or turbulent, 
the PNS code has been marched up to an X/L = 0.7 wh1ch 1S where the bow-shock/w1ng-
shock 1nteract10n reg10n occurs. In this work, L refers to the length of the 
veh1c1e. 
II. COMPUTATIONAL TECHNIQUE 
The PNS equat10ns are obta1ned from the complete Nav1er-Stokes equations by 
neglecting the unsteady terms and the streamw1se V1SCOUS derivative terms. The com-
plete deta1ls of all the terms and derivations can be found in Ref. 12. 
In the present formulation, ~ (the march1ng direction in computational space) 1S 
a function of x only (ax1s-normal marching planes). The governing equations are 
hyperbo11c-parabo11c in this ~-direct1on 1f the 1nviscid part of the flow field 1S 
supersonic, 1f there is no streamwise (axial) separat10n, and if the pressure gra-
d1ent in the viscous reg10n near the wall is treated correctly. However, the system 
of equations st11l allows for the separation 1n the crossf10w plane (n-~). 
The present PNS code uses the Beam-Warming implicit algorithm to update the 
inter~or of the region and characteristic, ~mplic~t, spatially second-order-accurate 
boundary conditions at the outermost shock wave. An elliptic gr~d generator of the 
type developed by Steger and Sorenson [13] and further specialized to w~ng bodies by 
Ra~ et al. [10] ~s used to generate the grid for the calculat~ons. 
If the conditions in a particular region are such that the march~ng procedure 
~s ~nval~dated, the unsteady Navier-Stokes (UNS) code is used for these reg~ons. In 
calculat~ng the flow over the Space Shuttle, one such region occurs ~n the vic~nity 
of the bow-shock/w~ng-shock interact~on (a pocket of subsonic flow is encountered). 
The complete details of the UNS code can be found in Ref. 11. The UNS code is 
extremely versatile and relat~vely easy to use. It uses either a Beam-Warm1ng 
impl~cit algorithm or a hybrid scheme due to Rizk and Chaussee [11]. The outer shock 
wave is e~ther fitted or captured. Usually the initial guess is furn~shed by the PNS 
code, which is modified in some manner to march through reg~ons where ~t would not 
march before. This procedure is acceptable, since the unsteady code takes this rea-
sonable guess and ~terates in time until a steady-state solution ~s obta1ned. 
The domain of th~s unsteady calculation encompasses the subson~c flow. The 
outflow boundary cons~sts ent~rely of superson~c axial flow ~n the ~nviscid part of 
the flow f~eld. Th~s permits the PNS code, which ~s more efficient, to continue 
marching from th~s po~nt. 
III. RESULTS 
Numerical results have been obtained for the following w~nd-tunnel cond~tions: 
Moo = 7.9, a = 25°, Twall = 540 0 R, ReL = 60,728/in. turbulent flow. For this calcula-
t~on, the Shuttle surface coord~nates were obta~ned from Rockwell-International 
Corporation. The current geometry consists of the complete Shuttle; the canopy, 
OMS pods, and the vertical stab~l~zer are ~ncluded. 
The three-d~mensional blunt-body code or~g~nally developed by Kutler et al. [14] 
was used to obtain the blunt-nose solut~on which creates the necessary start~ng 
planes for the PNS code at X/L = 0.0522. Th~s solut~on was then marched downstream 
using the el11ptic grid generator to construct the gr1d between the body and the 
f1tted outermost shock wave. The grid cons1sted of e1ther 61 or 121 points 1n the 
merid10nal direct10n and 45 geometr1cally stretched radial p01nts. An example of the 
grid at an X/L = 0.66 is shown 1n F1g. 1. The outermost gr1d l1ne 1S the bow wave, 
which 1S fitted uS1ng an 1mplicit techn1que. 
The pressure contours 1n the reg10n of the canopy are Jresented 1n Figs. 2 and 3. 
In F1g. 2, the contours on the lee p1tch plane of symmetry uetween X/L = 0.067 and 
X/L = 0.4 are shown. In the canopy reg1on, the coalescence of contours details the 
canopy shock wave followed by an expansion wave on the lee of the Shuttle. The con-
tours for a cross sect10n at an X/L = 0.2 are presented 1n F1g. 3. The canopy 
shock ~s once aga~n viewed at the p01nt on the lee where the pressure contours 
coalesce. The expansions which are visible on the windward are due to discontinui-
ties 1n geometry. 
The Mach number contours at an X/L = 0.66 are presented in F1g. 4. The main 
features are the wing shock and the crossflow shocks on the wing and upper body, 
respectively. These are denoted by the coalescence of the Mach contours. 
In Fig. 5, the crossflow veloc1ty vectors are presented at an X/L = 0.66. Two 
1nterest1ng features seen in this f1gure are the recirculation region in the w1ng-
body Juncture and the lee vortex. 
The density contours in the vicin1ty of the wing tip at an X/L = 0.667 are 
shown 1n Fig. 6. The wing shock and the bow shock have 1nteracted as characterized 
by the bulge in the outer boundary. This 1S due to the w1ng shock becoming the outer-
most surface, w1th the bow shock being captured. The bow shock appears as the coales-
cence of the dens1ty contours near the outer surface. 
By numerically s1mulating 011 flow on the surface of the vehicle, as 1n Fig. 7, 
many 1nteresting features are observed. The lines of separation on both the strake-
w1ng and the lee of the body are eV1dent by the coalescence of the numerical oil flow. 
The reattachment l1ne 1S v1s1ble on the Shuttle as a ser1es of 01l-flow lines d1verg-
1ng toward the separat10n l1nes. 
The computer-generated part1cle paths of F1g. 8 exh1b1t the same trends 1n the 
flow f1eld that are v1sible on the Shuttle surface V1a the 011 flow. Spec1f1c fea-
tures are the vortices on the lee wh1ch are due to the strake-wing. At this angle of 
attack, the vort1ces that are generated on the W1ng 1mpact on the OMS pod. 
IV. SUMMARY 
A procedure has been presented for ca1cu1at1ng the flow over veh1cles that have 
embedded reg10ns of subson1c flow in the 1nvisc1d part of the flow f1e1d. A PNS 
march1ng code 1S used to obtain the solut10n up to the bow-shock/w1ng-shock interac-
t10n reg10n. In th1s 1nteract10n reg10n, the UNS code can be employed S1nce the 
reg10n has a pocket of subson1c flow. Currently, only the results for the marching 
code up to an X/L = 0.667 are 1nc1uded. In the future, the results for the bow-
shock/w1ng-shock reg10n will be ava1lable. 
V. REFERENCES 
1. Venkatapathy, E., Rak1ch, J. V , and Tanneh111, J. C., '~umerica1 Solution of 
Superson1c V1SCOUS Flow over Blunt Delta Wings," AIAA Paper 82-0028, 1982. 
2. Prabhu, D. K. and Tanneh111. J. C., '~umerical Solut10n of Space Shuttle Orbiter 
Flow F1eld Inc1ud1ng Real Gas Effects," AIAA Paper 84-1747, 1984. 
3. Balakr1shnan, A., "Computat10n of V1SCOUS Real Gas Flow Field for the Space 
Shuttle Orb1ter," AlAA Paper 84-1748, 1984. 
4. L1, C. P., "App11cat10n of an Imp11c1t Techn1que to the Shock-Layer Flow Around 
General Bodies." AIAA Journal, Vol. 20, 1982, p. 175. 
5. Szema, K. Y., Griffith, B. J., Maus, J. R., and Best, J. T., "Lam1nar Viscous 
Flow F1e1d Pred1ct10n of Shuttle-like Veh1cle Aerodynam1cs," AIAA Paper 83-0211. 
1983. 
6. We1lmuenster, K. J., "H1gh Angle of Attack Invisc1d Flow Calculations Over a 
Shuttle-like Veh1cle with Compar1sons to F11ght Data," AIAA Paper 83-1798, 1983. 
7. Weilmuenster, K. J. and Hamilton, H. H., "Calculations of Inviscid Flow Over 
Shuttle-like Vehicles at High Angles of Attack and Comparisons with Experimental 
Data," NASA TP-2l03, 1983. 
8. Chaussee. D. S., Kutler, P., and Holtz, T., "Inviscid Supersonic/Hypersonic Body 
Flow Field and Aerodynamics from Shock-Capturing Technique Calculations," Journal 
of Spacecraft & Rockets, Vol. 13, 1976, pp. 325-331. 
9. Rai, M. M. and Chaussee, D. S., '~ew Implicit Boundry Procedures: Theory and 
Applications," AIAA Paper 83-0123, 1983. 
10. Rai, M~ M., Chaussee, D. S., and Rizk, Y. M., "Calculation of Viscous Supersonic 
Flows over Finned Bodies." AIAA Paper 83-1667, 1983. 
11. Rizk. Y. M. and Chaussee, D. S., "Three-Dimensional Viscous-Flow Computations 
Using a Directionally Hybrid Implicit-Explicit Procedure," AIAA Paper 83-1785, 
1983. 
12. Schiff, L. B. and Steger, J. L., "Numerical Simulation of Steady Supersonic 
Viscous Flow," AIAA Paper 79-0130, 1979. 
13. Steger, J. L. and Sorenson, R. L., "Automatic Mesh-Point Clustering Near a 
Boundary in Grid Generation with Elliptic Partial Differential Equations," 
Journal of Computational Physics, Vol. 33, 1979, pp. 405-410. 
14. Kutler, P., Pedelty, J. A., and Pulliam, T. H., "Supersonic Flow Over Three-
Dimensional Ablated Nosetips using an Unsteady Implicit Numerical Procedure," 
AIAA Paper 80-0063, 1980. 
14 
12 
10 
8 
6 
Z 
4 
2 
0 
-2 
-4 
0 2 4 6 8 
Y 
Fi.g. 1 Elliptic grid at X/'L ... O. 66 ~ 
9 
61 x 45 GRID 
8 
7 
6 
Z5 
4 
3 
2 
2 3 4 
ROCKWELL 
INTERNATIONAL 
SHUTTLE 
5 
X 
6 7 8 9 
Fig. 2 Pressure contours in the lee pitch 
plane of symmetry. 
5 
4 
3 
2 
Z 
1 
0 
-1 
-2 
0 1 2 3 4 
Y 
Fig. 3 Pressure contours at X/L = 0.2. 
4 
3 
2 
Z 1 
0 
--1 
--2 
0 1 2 3 4 5 
Y 
Fig. 5 Crossflow veloc~ty vectors at 
X/L = 0.66. 
6 
14 
12 
10 
8 
6 
Z 
4 
2 
0 
-2 
-4 
0 2 4 6 8 
Y 
Fig. 4 Mach contours at X/L = 0.66 •. 
-.2 
-.4 
-.6 
Z -.8 
-1.0~~.~ 
-1.2 
-1.4 
61 x 45 GRID 
.~;~ 
-1.6 ! .................. - .. ---.---.------------
3.4 3.6 3.8 4.0 4.2 4.4 4.6 4.8 5.0 
Y 
Fig. 6 Density contours near the wing tip 
at X/L = 0.667. 
61 x 45 GRID 
Fig. 7 Computational oil flow on the Shuttle surface up to X/L 0.66. 
61 x 45 GRID 
Fig. 8 Computational particle paths up to X/L = 0.66. 
1 Report No I 2 Government Accession No 3 Recipient's catalog No 
NASA TM 85977 
4 Title and Subtitle 5J~~rt f~~4 
VISCOUS COMPUTATION OF A SPACE SHUTTLE FLOW FIELD 6 Performing Organization Code 
A-9799 
8 Performing Organization Report No 
Denny S. Chaussee, Yeh1a M. R1Zk* and 
P1eter G. Bun1ng 10 Work Unit No 
9 Performing Organization Name and Address T-6458 
Ames Research Center 11 Contract or Grant No 
Moffett F1eld, CA 94035 
13 Type of Report and Period Covered 
12 Sponsonng Agency Name and Address 
Techn1cal Memorandum 
Nat10nal Aeronaut1cs and Space Adm1n1strat1on 14 Sponsoring Agency Code 
Wash1ngton DC, 20546 505-31-01 
15 Su pplementary Notes 
*Informat1cs General Corporat1on, Palo Alto J CA 
POlnt of contact. Denny S. Chaussee, Ames Research Center, MS 292A-14 
Moffett F1e1d, CA 94035, (415) 965-6742 or FTS 448-6742 
16 Abstract 
A procedure 1S presented, as well as some results, to calculate the flow 
over the w1nged Orblter. Th1S necessitates the use of two computer codes. 
A parabo11zed march1ng Nav1er-Stokes code 1S used to obta1n the solution up 
to the bow shock-w1ng shock 1nteract1on reg10n and for the region 'lfter 
the lnteract1on. An unsteady Nav1er-Stokes code 1S to be used 1n the reg10n 
of the shock 1nteract1on. Only results for the march1ng code are presented. 
For the flow cond1t1ons calculated, Moo = 7.9, a. = 25°, Twa 11 = 540 0 R, 
ReL = 60728 per 1nch, 1am1nar or turbulent, the PNS code has been marched up 
to an X/L = 0.7 which is where the bow shock-w1ng shock interact10n reg10n 
occurs. 
17 Key Words (Suggested by Author(sl) 18 Distribution Statement 
Cornputatlonal f1u1d dynamlcs 
Supersonlc flow Un11m1ted 
Shuttle 
Three-d1mens1ona1 V1SCOUS flow Sub] ect category: 34 
19 Security Classlf (of thiS report) 
j20 Security Classlf (of thiS page) 21 No of Pages 
22 Price· 
UnC1aSS1f1ed Unc1ass1f1ed 8 A02 
·For sale by the National Technical Information Service Springfield Virginia 22161 
End of Document 


Viscous computation of a space shuttle flow field

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