Document
CONSIDERATIONS REGARDING
THE EVALUATION A N D REDUCTION
OF SUPERSONIC SKIN FRICTION
by John B. Peterson, Jr*, and WiZZiam J. Monta
Langley Research Center
Langley Station, Humpton, Vk
i I N A T I O N A L A E R O N A U T I C S A N D SPACE A D M I N I S T R A T I O N W A S H I N G T O N , D. C . OCTOBER 1966 i li
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TECH LIBRARY KAFB, "I
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0079997 N A S A 'I" Ll-33UU CONSIDERATIONS REGARDING THE EVALUATION AND REDUCTION O F SUPERSONIC SKIN FRICTION By John B. Peterson, Jr., and William J. Monta Langley Research Center Langley Station, Hampton, Va.
NATIONAL AERONAUT ICs AND SPACE ADMINISTRATION For sale by the Clearinghouse for Federal Scientific and Technical Information Springfield, Virginia 22151 - Price $1.00 CONSIDERATIONS RFGARDING THE EVALUATION AND REDUCTION
OF SUPERSONIC SKIN FRICTION^
By John B. Peterson, Jr., and W i l l i a m J. Monta Langley Research Center SUMMARY ' A comparison i s made between previously published experimental data f o r supersonic turbulent boundary-layer skin f r i c t i o n and t h e skin-friction pre dictions obtained by using t h e Sommer and Short T ' and Spalding and Chi methods. A l s o , various methods f o r reducing skin f r i c t i o n on t h e supersonic transport a r e discussed.
INTRODUCTION Although t h e wave drag and t h e drag due t o l i f t of the proposed super sonic transport configurations have been greatly reduced as the design has progressed, t h e skin-friction drag has remained r e l a t i v e l y constant.
Because the skin-friction drag of a t y p i c a l supersonic transport i s a large part of t h e t o t a l drag, reduction of the skin-friction drag i s p o t e n t i a l l y a good means of obtaining drag reductions.
Figure 1 shows a breakdown of the drag of a t y p i c a l supersonic trans port cruising at a Mach number of 2.7 and an a l t i t u d e of 65 000 f t . As can
I
be seen, skin f r i c t i o n accounts f o r about 40 percent of the t o t a l drag.
Since a t y p i c a l transport has about 100 counts of t o t a l drag at cruise, the skin-friction drag i s about 40 counts.
For each count t h a t t h e drag can be reduced, the l i f t - d r a g r a t i o can be increased by about 0.1.
Several methods which can be used t o reduce the skin-friction drag on a supersonic transport will be reviewed i n t h i s paper. I n addition, methods f o r calculating skin-friction drag w i l l be reviewed.
SYMBOLS b wing span wing chord C CD drag coefficient drag coefficient due t o skin f r i c t i o n CD .F ._ lpresented a t the- c l a s s i f i e d "Conference on Aircraft Aerodynamics," Langley Research Center, May 23-25, 1966, and published i n N A S A SP-124.
i.ncrement in drag coefficient due to skin friction ~ D , F average skin-friction cqefficient CF adiabatic wall average skin-friction coefficient ' F , aw average skin-friction coefficient with air injection C~,injection average skin-friction coefficient without air injection CF,no injection average skin-friction coefficient for n = 0 cF, n=O local skin-friction coefficient Cf incompressible local skin-friction coefficient (see Cf,i section "Methods of Evaluating Skin Friction") h altitude M Mach number
m injection-air mass-flow rate
n planform exponent (see fig. 10) Reynolds number based on distance from leading edge R l Reynolds number based on distance f r o m virtual origin RX of turbulent boundary layer S reference surface area adiabatic wall temperature Taw wall temperature TW free-stream air velocity
v m
W injection-air weight-flow rate distance from center line in spanwise direction Y recovery factor qr free-stream air density pm Subscripts : experimental value exp theor t h e o r e t i c a l value DISCUSSION Methods of Evaluating Skin Friction Several years ago, a comparison between the various theories of super sonic skin f r i c t i o n i n use at t h a t time and the available experimental data showed t h a t the Sommer and Short T ' method generally gave t h e best prediction A new method of compressible turbulent boundary-layer skin f r i c t i o n ( r e f . 1).
f o r the prediction of compressible turbulent skin f r i c t i o n has since been Also, some new experimental data f o r developed by Spalding and Chi ( r e f . 2).
supersonic skin f r i c t i o n at high Reynolds numbers have extended the range of Reynolds numbers over which experimental r e s u l t s are available ( r e f s . 3 t o 9).
In a l l of these references except reference 9, l o c a l skin f r i c t i o n w a s measured.
w i l l be made by using l o c a l rather than average Therefore, comparisons herein skin-friction measurements.
Figures 2 t o 7 show comparisons between presently available experimental data and the skin-friction predictions of Sommer and Short ( f i g s . 2, 4, and 6) and Spalding and Chi ( f i g s . 3 , 3 , and 7). (The experimental data presented i n
the figures were obtained from references 3 t o 8 and i o t o 12.) The skin f r i c
t i o n i s presented i n the form of the r a t i o - Cf , where the value of c f , i i s
Cf,i predicted by the method involved. A s will be shown, neither method gives com over the e n t i r e range of Reynolds numbers.
pletely satisfactory r e s u l t s with M
I n figures 2 and 3 are shown the experimental variations of - Cf
'f,i
f o r a value of Rx of 10 x lo6 compared with the predictions of Somer and
The curves and data shown a r e a l l Short ( f i g . 2 ) and Spalding and Chi ( f i g . 3 ) .
f o r adiabatic w a l l temperatures. There is some s c a t t e r i n the experimental data, but generally the Sommer and Short T ' prediction agrees s l i g h t l y b e t t e r with the data at t h i s Reynolds number than the Spalding and Chi prediction.
When the new experimental data f o r skin f r i c t i o n at higher Reynolds numbers of 50 x 10 6 and 100 x 106 are used i n the same type of comparison, the
r e s u l t s are not the same, as shown i n figures 4 and 5 . (The data of Hopkins
and Keener were obtained f o r a Reynolds number based upon momentum thickness.
These data a r e converted i n figures 4 and 5 t o values f o r Rlr by using the
method of reference 1.) For these conditions, the data agree b e t t e r with the prediction Spalding and Chi prediction and l i e above the Sommer and Short T ' curve.
The reason f o r the agreement of the data with the Sommer and Short T' prediction at low Reynolds numbers and with the Spalding and Chi prediction at high Reynolds numbers can be shown by plotting t h e skin-friction r a t i o as a function of' Reynolds number f o r a constant Mach number. The Sommer and Short T ' curve is shown i n figure 6 and the Spalding and Chi curve is shown i n figure 7.
a l l obtained f o r Mach numbers between 2.20 and 2.95 and trans The data were formed t o values f o r a Mach number of 2.7 by using the equation Cf
The values obtained f o r - by using t h i s equation a r e not the same f o r the
i two methods since t h e parameter is dependerit upon the p a r t i c u l a r Cf,theor.
method involved. Therefore, the values of t h e r a t i o - Cf are s l i g h t l y dif-
C f , i
ferent i n figures 6 and 7. However, t h i s procedure allows a d i r e c t comparison
t o be made between t h e data and the predicted curves by preserving the relation of experimental values t o predicted values. It appears from these data t h a t the skin-friction r a t i o is almost independent of Reynolds number. Both methods predict some variation of the skin-friction r a t i o with Reynolds number and, therefore, neither prediction curve matches the data over the e n t i r e range of Reynolds numbers. I n order t o predict the average skin f r i c t i o n , it is impor t a n t t h a t t h e method give accurate r e s u l t s f o r the l o c a l skin-friction l e v e l at a l l Reynolds numbers up t o the Reynolds number of i n t e r e s t , since the average is obtained by integrating the l o c a l values.
Therefore, even though t h e Spalding and Chi method gives accurate r e s u l t s f o r the l o c a l skin f r i c t i o n at high Reynolds numbers, it does not necessarily give accurate results f o r the average skin f r i c t i o n a t these Reynolds numbers. There is a l s o some doubt as t o the v a l i d i t y of the Spalding and Chi method f o r use at the high temperature levels encountered on a supersonic transport, since t h e constants i n t h i s method were obtained by comparison with wind-tunnel data and no provision w a s made t o account f o r the e f f e c t of temperature l e v e l on the viscosity r a t i o of air.
Most other methods of predicting skin f r i c t i o n , including the Sommer and Short T ' method, do have such a provision.
As can be seen, only a limited amount of experimental data i s available f o r a supersonic transport. More data the very high Reynolds numbers encountered by are needed t o increase confidence i n the prediction of skin f r i c t i o n at high Reynolds numbers.
The e f f e c t of w a l l temperature on the average skin f r i c t i o n at M = 3.0 and Rx = 94 x 106 i s shown i n figures 8 and 9 f o r an ogive-cylinder body of _-
Tw revolution ( r e f . 9). The variations of - with - are presented f o r
CF , a w Taw
experiment and theory, where Taw i s based on a recovery f a c t o r qr of 0.89 and t h e value of C F , ~ ~ is obtained by extrapolating the experimental values t o adiabatic conditions. A comparison of these figures shows t h a t the Sommer and Short T ' method b e t t e r predicts the e f f e c t of wall temperature on skin f r i c t i o n at Mach 3 . However, the heat-transfer correlations in reference 13 indicate t h a t the Spalding and Chi method is more accurate at hypersonic speeds.
It is apparent t h a t there is much room f o r improvement i n the accuracy of predictions of turbulent skin-friction drag. However, t h e Sommer and Short T' method is considered t o provide the best predictions of skin f r i c t i o n under conditions encountered by the supersonic transport. Therefore, t h i s method is used t o calculate skin f r i c t i o n i n the r e s t of t h i s paper.
Methods of Reducing Skin Friction Most of the methods discussed i n t h i s paper f o r reducing the skin f r i c t i o n on a supersonic transport have been presented before i n various conference papers and NASA reports. These methods are presented herein without regard t o the design considerations involved, o r the e f f e c t s they might have on other characteristics of the a i r c r a f t . Application t o a supersonic transport w i l l require careful and ingenious design i n order t o obtain favorable overall r e s u l t s .
Configuration changes and blending.- One way t o reduce skin f r i c t i o n i s t o take advantage of the f a c t t h a t skin f r i c t i o n is low at high Reynolds numbers.
(See r e f . 14.) Figure 10 i l l u s t r a t e s the changes i n skin f r i c t i o n which occur as the wing planform i s changed so as t o remove areas from the t i p s and add areas i n the center, where they w i l l be i n high Reynolds number flows. The skin f r i c t i o n w a s calculated at a Mach number of 2.7 and an a l t i t u d e of 65 000 f t f o r a wing with a planform area of 8000 f t 2 and an aspect r a t i o of 1.7. The wing chord w a s determined by a power-law formula, and the midchord sweep of the wing w a s held constant at W o . As can be seen, the areas near the t i p s are progressively movedtoward t h e center of the wing. This process r e s u l t s i n a reduction i n the t o t a l skin f r i c t i o n , even though the t o t a l area and the aspect r a t i o of the wing remain the same.
Another obvious way of reducing skin f r i c t i o n i s decreasing the wetted area of the a i r c r a f t . The method used t o decrease the wetted area, which is called blending, is accomplished by deforming the a i r c r a f t into a shape t h a t is as close as possible t o a body of revolution. Such a shape, of course, would have the l e a s t surface area f o r a given volume distribution. Although t h i s type of blending i s used t o reduce skin f r i c t i o n only, it is not incom patible with the type of blending which can be used t o reduce wave drag and s t r u c t u r a l weight.
An example of a configuration shape which resulted from blending and changing the planform of a delta-wing type supersonic transport is shown i n figure 11. The wing planform has been changed t o remove areas near the t i p s and add areas near the center i n such a way t h a t t h e t o t a l area and the wing aspect r a t i o are constant. The wing and t a i l have been blended into the fuse lage with large fillets; the nacelles have been blended together and a s p l i t t e r p l a t e used t o separate the i n l e t s . The data of reference 15 show t h a t a
I I 1
s p l i t t e r p l a t e prevents mutual interference between i n l e t s when they are The nacelle i n l e t s have not been blended i n t o the wing because unstasted.
such blending would have produced problems of diverting the wing boundary layer around the i n l e t s .
The skin-friction-drag reductions t h a t might be obtained by these config uration changes are shown i n t a b l e I. The reductions due t o planform changes result only from removing areas i n low Reynolds number flows and replacing them i n high Reynolds number flaws. There i s no change i n the t o t a l wetted area.
The skin-friction-*% reductions due t o blending r e s u l t from changes i n the t o t a l wetted area which occur as the components of the a i r c r a f t are blended.
Changes i n the skin f r i c t i o n caused by three-dimensional e f f e c t s i n the corners w e r e neglected i n these calculations. Although each individual increment i s smal1,'the t o t a l increment can be a significant reduction i n the skin-friction For an a i r c r a f t with 40 counts of skin-friction drag, the t o t a l reduc drag.
t i o n shown i n table I amounts t o about 2 counts.
Effects of emissivitg on ___ wall temzerature a n d _ s k i n f r i c t i o n . - A s shown before, the wall-temperature r a t i o has a large e f f e c t on skin f r i c t i o n . Both the theory and the experimental data showed t h a t the skin f r i c t i o n increased as the wall-temperature r a t i o decreased. This trend i s shown i n figure 12, where the skin-friction drag i s plotted as a function of the wall-temperature flying at a Mach number of 2.7 and r a t i o f o r a t y p i c a l supersonic transport an a l t i t u d e of 65 000 f t . Also showh i n the figure a r e v e r t i c a l dashed l i n e s a t the temperature r a t i o s corresponding t o the equilibrium w a l l temperatures f o r various w a l l emissivities. (See also r e f . 16.) The range of emissivities being considered f o r presently proposed supersonic transports i s shown as the crosshatched region. As i s w e l l known, radiation of heat from the w a l l caused w a l l temperature. For t h i s particular con by high emissivities reduces t h e figuration, an emissivity of 0.5 reduces the wall-temperature r a t i o t o about it t o about 0.91. It can be seen t h a t 0.95, and an emissivity of 1.0 reduces increasing the emissivity reduces the wall temperature but increases the skin- f r i c t i o n drag. Low emissivities have the opposite e f f e c t of increasing the w a l l temperature and reducing the skin-friction drag. Therefore, low values of the emissivity, which can be controlled t o a certain extent by the choice of a supersonic paint o r surface coating used, reduce the skin-friction drag of transport. Determination of the best w a l l emissivity t o use w i l l depend on the exact configuration and s t r u c t u r a l design chosen.
Boundary-laxer control.- The i d e a l way t o reduce the skin f r i c t i o n on a supersonic transport, of course, would be with laminar-flow control. Research on laminar-flow control, however, is s t i l l continuing and very l i t t l e p r a c t i c a l experience has been obtained so f a r . Therefore, laminar-flow control does not appear feasible f o r the first-generation supersonic transport.
Theoretically, it i s possible t o obtain about 2 f e e t of natural laminar flow on unswept leading edges, such as the engine nacelles, and about 1.2 f e e t of natural laminar flow on swept leading edges, such as the wing and t a i l
( r e f . 17). With these amounts of laminar flow, the skin-friction drag could be
However, large extents of natural laminar flow on reduced by about 3 percent.
since t h i s condition would require very t h e supersonic transport appear unlikely, accurate construction, extensive maintenance, and some method of avoiding insect contamination during service.
There is, however, a method of reducing the turbulent skin-friction drag.
It has been shown experimentally t h a t the turbulent skin f r i c t i o n can be reduced by injecting air into the boundary layer through rearward-inclined flush s l o t s i n the surface (refs. 18 and 19). The e f f e c t of air injection on the drag is shown i n figure 13, i n which the model drag coefficient is plotted as a function of the injection mass-flow parameter. The lower curve presents the variation of the measured values of with the rate of air injection CD through a rearward-inclined flush s l o t at M = 3.0. These measured values of include the reduction i n skin f r i c t i o n as well as the t h r u s t recovered from CD the injected air. The upper curve represents the calculated values of CD t h a t could be obtained if the momentum t h r u s t of the same air were recovered with a convergent nozzle. The difference i n the levels of the two curves indi cates t h a t a reduction i n skin f r i c t i o n occurred. The physical process behind t h i s skin-friction reduction is not yet f u l l y understood.
A possible application of air injection t o one of the supersonic transport configurations i s presented i n figure 14. The abscissa i s the injection-air weight-flow r a t e throigh inclined flush s l o t s . The a i r f o r injection can be obtained from the i n l e t bleed air, which i s already available f o r use onboard the airplane. The ordinate is the r a t i o of the average airplane skin-friction coef f i c i e n t with air injection t o the average skin-friction coefficient without air injection. The flow r a t e of the i n l e t bleed air is estimated t o be about 80 lb/sec. With t h i s amount of air, the skin f r i c t i o n can be reduced by 5 percent.
Recently published boundary-layer surveys behind flush s l o t s ( r e f . 19) have suggested t h a t perhaps even larger reductions i n skin f r i c t i o n could be obtained from two or three s l o t s distributed along the surface, instead of one s l o t near the leading edge. This hypothesis requires experimental verification, however, before it can be used. The f e a s i b i l i t y of using air injection t o reduce skin f r i c t i o n depends upon many considerations. The point t o be made, however, is t h a t the skin-friction reductions shown i n figure 1 4 indicate t h a t further study of the use of air injection on the supersonic transport i s warranted.
CONCLUDING IiEMARKS In summary, a comparison between theory and the l a t e s t experimental r e s u l t s f o r compressible turbulent skin f r i c t i o n shows t h a t more data are needed t o increase confidence i n the prediction of skin f r i c t i o n at supersonic speeds and high Reynolds numbers.
Planform changes and c o n f i w a t i o n blending can significantly change the Also, the w a l l emissivity t o t a l skin-friction drag of a supersonic transport.
of a supersonic transport can have a large effect on the skin-friction drag.
I The reduction i n turbulent skin f r i c t i o n obtainable with air injection through rearward-inclined flush s l o t s indicates t h a t further study is warranted.
Langley Research Center, National Aeronautics and Space Administration, Langley Station, Hampton, Va., May 25, 1966, 126-13-03-22-23.
1 . Peterson, John By Jr.: A Comparison of the Experimental and Theoretical Results for the Compressible Turbulent-Boundary-Layer Skin Friction With Zero Pressure Gradient. NASA TN D-1795, 1963.
2 . Spalding, D. B.; and Chi, S . W . : The Drag of a Compressible Turbulent Boundary Layer on a Smooth Flat Plate With and Without Heat Transfer.
J. Fluid Mech., vol. 1 8 , pt. 1, Jan. 1964, pp. 117-143.
3 . Jackson, Mary W.; Czamecki, K. R . ; and Monta, William J . : Turbulent Skin Friction at High Reynolds Numbers and Low Supersonic Velocities. NASA TN D-2687, 1965.
4 . Monta, William J; and Allen, Jerry M . : Local Turbulent Skin-Friction Meas
urements on a Flat Plate at Mach Numbers From 2 . 5 to 4 . 5 and Reynolds
Numbers up to 6 9 x 106. NACA TN D-2896, 1965.
5. Aircraft Div., Douglas Aircraft Co., Inc.: Investigation of Skin Friction
Drag on Practical Construction Surfaces for the Supersonic Transport.
FDL TDR 64-74, U.S. Air Force, Aug. 1964.
6. Moore, D. R.; and Harkness, J . : Experimental Investigations of the Compres
sible Turbulent Boundary Layer at Very High Reynolds Numbers. AIAA J., vol. 3, no. 4, Apr. 1965, pp. 631-638.
7 . Hopkins, Edward J.; and Keener, Earl R . : Study of Surface Pitots for Meas
uring Turbulent Skin Friction at Supersonic Mach Numbers - Adiabatic Wall.
NASA TN D-3478, 1966.
8. Winter, K. G.; Smith K. G . ; and Gaudet, L.: Measurements of Turbulent Skin Friction at High Reynolds Numbers at Mach Numbers of 0.2 and 2.2. Recent Developments in Boundary Layer Research, Pt. I, AGARDograph 97, May 1965, PP- 97-l-24.
9. Czarnecki, K. R.; Jackson, Mary W.; andMonta, William J.: Studies of Skin
Friction at Supersonic Speeds. NASA Cmference on Supersonic-Transport Feasibility Studies and Supporting Research. NASA TM X-905, 1963, PP- 177-189.
10. Matting, Fred W; Chapman, Dean R.; Nyholm, Jack R.; and Thomas, Andrew G . : Turbulent Skin Friction at High Mach Numbers and Reynolds Numbers in Air and Helium. NASA TR R-82, 1961.
ll. Coles, Donald: Measurements in the Boundary Layer on a Smooth Flat Plate in Supersonic Flow. 111. Measurements in a Flat-Plate Boundary Layer at the Jet Propulsion Laboratory.
Rept. No. 20-71 (Contract No. DA-04-495-0rd 18), Jet Propulsion Lab., California Inst. Technol., June 1, 1953.
12. Shutts, W. H.; Hartwig, W. H.; and Weiler, J. E.: Final Report on Turbulent Boundary-Layer and Skin-Friction Measurements on a Smooth, Thermally Insulated F l a t Plate at Supersonic Speeds. Dm-364, CM-823 (Contract NOrd-9195), Univ. of Texas, Jan. 5, 1955.
13. Bertram, Mitchel H.; and N e a l , Luther, Jr. : Recent Experiments in.Hy-personic Turbulent Boundary Layers.
Presented t o the AGARD Specialists Meeting on Recent Developments i n Boundary-Layer Research (Naples, Italy), May 10-14, 14. Robins, A. Warner; H a r r i s , Roy V., Jr.; and Jackson, Charlie M., Jr.: Char a c t e r i s t i c s at Mach Number of 2.03 of a Series of Wings Having Various Spanwise Distributions of Thickness Ratio and Chord. NASA T N D-631, 1960.
15. Moseley, George W.; Peterson, John B., Jr.; and Braslow, Albert L.: An Investigation of S p l i t t e r Plates f o r the Aerodynamic Separation of Twin I n l e t s at Mach 2.5. NASA TN D-3385, 1966.
16. Allen, Jerry M.; and Czamecki, K. R.: Effects of Surface Ehnittance on Turbulent Skin Friction at Supersonic and Low Hy-personic Speeds. NASA TN D-2706, 1965.
17. J i l l i e , Don W.; and Hopkins, Edward J.: Effects of Mach Number, Leading-
Edge Bluntness, and Sweep on Boundary-Layer Transition on a Flat Plate.
NASA TN D-1071, 1961.
18. McRee, Donald I.; Peterson, John B., Jr.; and Braslow, Albert L.: Effect of A i r Injection Through a Porous Surface and Through Slots on Turbulent Skin Friction at Mach 3 . NASA TN D-2427, 1964.
19. Peterson, John B., Jr.; McRee, Donald I.; Adcock, J e r r y B.; and Braslow, Albert L.: Further Investigation of Effect of A i r Injection Through Slots and Porous Surfaces on Flat-Plate Turbulent Skin Friction a t Mach 3. NASA T N D-3311, 1966.
TABLE I
CALCULATED SKIN - FRICTION - DRAG REDUCTlONS
DUE TO CONFIGURATION CHANGES h = 6 5 0 0 0 FT M=2.7; CD,F PLANFORM CHANGES
WING ....................................................................... -0.5 9'0
TAIL ....................................................................... -0.1 9'0
TOTAL - 0.6 9'0
BLENDING
................................ 3.1 9'0
WING-FUSELAGE JUNCTURE
VERTICAL-TAIL-FUSELAGE JUNCTURE .............- 0.3 9'0
NACELLES .............................................................. - I . 3
TOTAL -4.7 '70 TOTAL CHANGE IN SKIN-FRICTION DRAG
DUE TO CONFIGURATION CHANGES ........................... -5.3 9'0
... .. . . . . . . . .
I DRAG BREAKDOWN OF A TYPICAL SUPERSONIC TRANSPORT Mz2.7; h = 6 5 000 FT DRAG DUE .0080 TO LIFT, 34%
t
ZERO-LIFT WAVE DRAG, CD Figure 1 EFFECT OF MACH NUMBER ON TURBULENT SKIN FRICTION SOMMER AND SHORT T' METHOD; Tw/Taw=l.O EXPERIMENT 1 . 0 0 JACKSON et al.
MONTA AND ALLEN A MATTING et al.
.8 OWINTER AND SMITH 1 7 COLES D SHUTTS et al.
.6 .4
'21 mMUWSONIC T T S P O R T
RANGE OF INTEREST //A I I 0 I 2 3 4 5 M Figure 2 EFFECT OF MACH NUMBER ON TURBULENT SKIN FRICTION SPALDING AND CHI METHOD; Tw/Taw=l.O EXPERlMENT et al.
OJACKSBN UMONTA AND ALLEN A MATTING et al.
- , .a OWINTER AND SMITH D COLES DSHUTTS et al.
-6 1 ‘9 9.i - PREDICTION .4 - . 2 SUPERSONIC TRANSPORT RANGE OF INTEREST .7%225/////// I I I I 2 3 4 5 M Figure 3 EFFECT OF MACH NUMBER ON TURBULENT SKIN FRICTION SOMMER AND SHORT T’ METHOD; Tw/Taw= 1.0 EXPERIMENT 1.0 0 JACKSON et 01.
0 MONTA AND ALLEN A DOUGLAS AIRCRAFT 0 MOORE AND HARKNESS DHOPKINS AND KEENER a MATTING et al.
0 WINTER et al.
:f .4
SUPERSONIC TRANSPORT
/- RANGE OF INTEREST
- 2 t Y / / / / / / / / / / U
0 I 2 3 4 5 M Figure 4 EFFECT OF MACH NUMBER ON TURBULENT SKIN FRICTION SPALDING AND CHI METHOD; Tw/Taw =LO EXPERIMENT 0 JACKSON et al.
0 MONTA AND ALLEN A DOUGLAS AIRCRAFT 0 MOORE AND HARKNESS DHOPKINS AND KEENER a MATTING et at.
0 WINTER et at.
SUPERSONIC TRANSPORT
/- RANGE OF INTEREST
0 I 2 3 4 5 M Figure 5 EFFECT OF REYNOLDS NUMBER ON TURBULENT SKIN FRICTION SOMMER AND SHORT T' METHOD; M=2.7; Tw/Taw=I.O .8 . 6 C f - '3,i EXPERIMENT .4 0 WINTER et 01.
0 MOORE AND HARKNESS A MATTING et 01.
.2 0 MONTA AND ALLEN 0 1 I I I 1 I I 5 IO 50 100 500 1000 2OO0x1O6 Rx Figure 6 EFFECT OF REYNOLDS NUMBER ON TURBULENT SKIN FRICTION SPALDING AND CHI METHOD; Mz2.7; Tw/Tow ~1.0
1 . 0 r
-
EXPERIMENT 0 WINTER et al.
0 MOORE AND HARKNESS A MATTING et al.
0 MONTA AND ALLEN .2
cf'-i
I 1 1 I -
O5 I O 50 loo 500 1000 2OoOx106 R * Figure 7 EFFECT OF WALL TEMPERATURE ON TURBULENT SKIN FRICTION SOMMER AND SHORT T' METHOD; Mz3.0; Rx= 9 4 ~ 1 0 ~ 0 CZARNECKI et al.
I -PREDl CTI ON
1 . 4 - 1.2 CF cF, ow I .o 1 I I I I I I 0 .2 .4 . 6 .a 1.0 1.2 Tw Taw Figure 8 I 111 I . . I , . , I 1 . 1 . 1 . .I .,,,,,.,.,, 111 I.. 1. 1 . m . 1 . I. .,.,,,,.. .,. .-. ..... -.. --.--. -.
EFFECT OF WALL TEMPERATURE ON TURBULENT SKIN FRICTION SPALDING AND CHI METHOD; Mz3.0; Rx=94 xi0' 0 CZARNECKI et ai.
I -PREDICT1 ON
1 ; 0 .2 .4 . 6 .8 1.0 1.2 Tw/ Taw Figure 9 EFFECT OF WING PLANFORM ON SKIN FRICTION WING AREA AND ASPECT RATIO CONSTANT I I I I I 0 .5 I .o 1.5 2.0 PLANFORM EXPONENT, n I- I I I I .o 1.5 2.0 2.5 3.0 RELATIVE ROOT CHORD Figure 10 CONFIGURATION CHANGES TO REDUCE SKIN FRICTION BLENDED CONFlGURATION Figure 11 EFFECT OF EMISSIVITY ON WALL TEMPERATURE AND SKIN FRICTION M = 2.7; h= 65 000 FT �M/SS/WTY
.0°4T /.o 0 .5
-0040 CD,F -
I
.0038 - I
- RANGE O F S S T PROPOSALS .0036L L a I I
0 .90 .95 I .oo
Tw/ Taw I I I 350 400 450 AVERAGE WALL TEMP., O F Figure 12 I - EFFECT OF AIR INJECTION ON DRAG M= 3.0; Rl = I6 x IO6 -0020 r CALCULATED DRAG WITH CONVERGENT - .OOlO REARWARD-INCLINED FLUSH SLOT - .0005 I I I 1 .I 0 2 4 6 8 x 10-4 Figure 13 POSSIBLE SKIN-FRICTION REDUCTION ON AN S S T WITH AIR INJECTION M = 2 . 7 ; h 65 000 FT I
-94 t
I I I I 1 . 1 I I 1 0 20 40 60 80 w , LB/SEC Figure 14 NASA-Langley, 1966 L-3250 “The aeronautical and spare activities of the United States shall be conducted so as to contribute . . . to the expansion of human knowl edge of phenomena i n the atmosphere and space. T h e Administration shall provide f o r the widest practicable and appro&ate dissemination o f information concerning its activities and the results thereof .” AERONAUTICS -NATIONAL AND SPACE ACT OF 1958
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