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RESEARCH MEMORANDUM
RESEARCH MEMORAN DUM
BOUNDARY -LAYER-TRANSITION MEASUREMENTS EOUNDARY -LAYER-'i'RANSITION NIEBSUREMEKU'TS ill FULL-SCALE FLIGHT DJ FULL-SCALE FLIGHT By Richard D. Banner, Jor..n G. McTigue, By RichardD. Banner, Jobin G. McTigue, and Gilbert Petty, Jr.
and G i l b e r t Petty, Jr.
High-Speed Flight Station High-speed Flight Station Edivar& c - a i . Edwards, C alii.
CLASS I F1 CAT I ON CHANGED Tf ~jr:n.ASS' FlED EfFECTIVE JUN~ 12 196~ A~O~tTy }.IASA ~:, 4 Bi J, J, CARROLL
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revelation of which in any manner to an unautb:.ru
NATIONAL ADVISORY COMMITTEE
NATIONAL ADVISORY COMMITTEE
FOR AERONAUTICS
F O R AERONAUTICS
WASHINGTON WASHINGTON July 28, 1958 July 28, 1958 NACA RM H58U8 NATIONAL ADVISORY COMMI'I'TEE FOR AERONAUTICS RESEARCH MEMQRANDUM BOUNDARY-LAYER-TRANSITION MEASUREMENTS IN FULL-SCALE F L I G W By Richard D. Banner, John G. McTigue, and Gilbert Petty, Jr.
SUMMARY
/ 90 4 3
Chemical sublimation has been employed for boundary-layer-flow visualization on the wings of a supersonic fighter airplane in level flight at speeds near a Mach number of 2.0. The tests have shown that laminar flow can be obtained over extensive areas of the wing with practical wing-surface conditions.
In addition t o the flow visualization tests, a method of contin- uously monitoring the conditions of the boundary layer has been applied to flight testing, using heated temperature resistance gages installed in a Fiberglas "glove" installation on one wing. Tests were conducted at speeds from a Mach number of 1.2 to a Mach number of 2.0, at alti- tudes from 35,000 feet to 56,000 feet.
Data obtained at all angles of attack, from near Oo to near loo, have shown that the maximum transition Reynolds number on the upper s u r - face of the wing varies from about 2.5 x 1 0 at a Mach number of 1.2 to about 4 x 1 0 at a Mach number of 2.0. On the lower surface, the maxi- mum transition Reynolds number varies from about 2 x lo6 at a Mach n u - ber of 1.2 to about 8 x 1 0 at a Mach number of 2.0.
INTRODUCTION Because of the greatly increased need for knowledge of full-scale boundary-layer transition and the difficulty of simulating actual flight conditions, a program has been initiated to provide a better understanding of the boundary-layer flow as it exists in supersonic flight. This paper shows the results obtained in the early flight tests which determined the extent of laminar flow that could be obtained with practical wing-surface conditions.
*Title, Unclassified.
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NACA RM ~ 5 8 ~ 2 8 SYMBOLS V m / v , per ft R Iieynolds number per foot,
free-stream ve loc ity , f t / se c
V ' , kinematic viscosity V nondimensional Reynolds number based on x
%
X distance from leading edge sweep angle
n
U angle of attack altitude hP M Mach number 7 thickness, in.
INSTRWNTATION AND TECHNIQUES A fighter airplane was instrumented, as shown in figure 1 , for trrnsition investigations on the wings. The basic wing has a modified biconvex sirfoil xith a thickness ratio of 3.4 percent, a sharp leading A l/l0-inch-thick edge, 2nd a sllght ranount of sweep (about 27O).
F'iberglas glove w3c installed on the right wing and was instrumented with one row of transition detectors on both the top and bottom surfaces.
These detectors provided continuous monitoring of the laminar and tur- bulent boundary-layer-flow conditions (ref. 1) .
Chcnicctl sublimation was employed for boundary-layer-flow visualiza- tion on both wings, and cameras (fig. 1) were installed for recording the chemical indications. Many investigators have used the chemical sublima- tion technique in both wind tunnels and in flight (refs. 2, 3, and others).
of this technique in flight to speeds These tests have extended the use ;z lhch number of 2.0.
near The trznsition-detector signals (see fig. 2) were multiplexed and conform to recorded on an oscillograph. The sequencing was scheduled to the lccations of the detectors on the wing. This arrangement allowed 0 0 0.. . . 0 0 . 0 . 0 0.. 0 0.0 0 .
0 . 0 0 . 0 0 0 0 . 0 . 0 . 0 .
0 . 0 . 0 0 . . 0 0 0 0 .
. 0 . 0 0 . 0 .
NACA RM ~ 5 8 ~ 2 8 : e ' 0. . 0 . 0 0 . 0 .
location of the laminar and turbulent flow areas, within about 5 percent of t h e chord, by inspection of the records. The reasons f o r using flow v i s u a l i z a t i o n a r e i l l u s t r a t e d i n figure 2 . Turbulent wedges, originating A s can upstream of the detectors, cause l o c a l areas of turbulent flow.
be seen, t h e t h i r d detector indicates turbulent flow i n an area t h a t would otherwise be laminar.
RESULTS AND DISCUSSION From t h e 33-millimeter f l i g h t f i l m of the chemical indications, photographic enlargements have been made and a t y p i c a l i n - f l i g h t photo- graph of t h e lower surface of the Fiberglas covered wing i s shown i n figure 3 . The w h i t e chemical remaining in the v i c i n i t y of t h e leading edge indicates the extent of laminar flow being experienced on t h e w i n g .
The f i e l d of view of the camera includes the area of the wing from t h e leading edge rearward t o j u s t behind t h e a i l e r o n hinge i n the outboard In a l l the t e s t s no area and some of the inboard area of t h e w i n g .
laminar flow had been observed i n the inboard area, and f o r t h a t reason t h i s area i s omitted i n subsequent photographs of t h i s presentation.
The area shown i s outboard of the 47-percent-exposed-span s t a t i o n .
I n some cases the airplane returned from t h e flight with a chemical indication remaining on the wing. It w i l l be of i n t e r e s t t o look a t one such indication before proceeding with the main p a r t of the discussion.
Figure 4 presents an enlargement of t h e leading-edge region of the w i n g .
The section seen i s about 1 square f o o t . Note the s t r i a t i o n s t h a t can be seen i n the chemical. Other investigators have a l s o observed these s t r i a t i o n s i n a laminar boundary layer, both i n wind-tunnel tests (ref. 4) and i n f l i g h t ( r e f . 5 ) a t subsonic speeds. The s t r i a t i o n s have been a t t r i b u t e d t o the presence of vortices which a r e shed from t h e swept leading edges and contribute t o the breakdown of the normally laminar flow. Although it could not be determined when t h i s phenomenon occurred during the f l i g h t , it i s believed t o be worth mentioning since it appears t o be a problem t h a t must be considered i n determining t h e extent of laminar flow that could be expected on swept wings.
Turning now t o t h e f l i g h t photographs t h a t were taken during t h e t e s t s , figure 5 shows the e f f e c t of t h e leading-edge-flap "piano type" A s can be seen, the hinge tripped t h e hinge on producing t r a n s i t i o n .
laminar boundary layer producing turbulent wedges which merge rearward of the hinge t o form completely turbulent flow over the remainder of the wing. The laminar area i s approximately 15 percent of t h e test area.
This condition of the wing is referred t o as unfinished. I n improving the wing-surface conditions the f l a p hinge was f i l l e d t o eliminate any abrupt d i s c o n t i n u i t i e s . Also, a l l rivetheads and screwheads were ground flush with the wing skin and f i l l e r material w a s applied t o f i l l any 0 . 0.. . ... . 0 . 0 . . . . 0.. 0 .
0 . .. 0 . . 0 . . 0 . . 0 . .
4 NACA RM ~ 5 8 ~ 2 8 p i t s or small depressions. The whole surface w a s then sanded. This condition of the wing i s r e f e r r e d t o as t h e f i n i s h e d wing. Following the tests with the finished wing, t h e wing w a s painted and polished.
The e f f e c t of these improvements can be seen by comparing f i g u r e 6 w i t h f i g u r e 5 . Although the Mach number f o r the t e s t with the painted wing i s s l i g h t l y d i f f e r e n t , the v a r i a t i o n i n t h e a l t i t u d e s resulted i n same angles of a t t a c k .
the same free-stream Reynolds number and the In comparing the unfinished and finished w i n g lower surfaces, it can be seen t h a t considerably more laminar flow was obtained on the finished wing. This i s primarily due t o smoothing over t h e leading- edge-flap hinge. Painting t h e wing surface reduced the average rough- ness from about 25 to 13 microinches, but the e f f e c t on t r a n s i t i o n was not appreciable on e i t h e r t h e top o r bottom surface. The extent of laminar flow on the painted wing i s about 25 percent of the t e s t area f o r the upper surface and about 35 percent of the t e s t area f o r the lower surface.
Realizing t h a t the standards t h a t had been set f o r roughness were a r b i t r a r y and t h a t they might d i f f e r from those set i n the wind rather tunnel, it was f e l t , nevertheless, t h a t t h e maximum i n p r a c t i c a l improve- The,extent of laminar flow ments t o t h e wing surface had been reached.
considered t o be t h a t w a s observed on the finished and painted wing i s representative of the m a x i m u m t h a t might reasonably be expected f o r these f l i g h t conditions. This conclusion was arrived a t because the extreme c a r e . t h a t was taken i n producing t h e Fiberglas surface f i n i s h had resulted i n an average roughness of only 7 microinches.
A comparison of the finished and painted wing and the Fiberglas covered wing i s shown i n figure 7. For c l a r i t y , the leading edges a r e a l l shown t o the l e f t . Covering the wing with Fiberglas had s l i g h t l y a l t e r e d t h e wing p r o f i l e , and the leading edge had been rounded t o l/lO-inch radius, instead of the sharp leading edge of the basic wing.
A l s o , waviness measurements a t l/2-inch increments indicated an average deviation of about 0.003 inch on the Fiberglas covered wing as compared ‘ w i t h 0.006 inch on the basic wing. Exactly what e f f e c t these changes produced l o c a l l y could not be determined; however, as can be seen, no large differences i n t h e overall extent of laminar flow i s evidenced.
I n order t o d.etermine the e f f e c t of Mach number and a l t i t u d e on the extent of laminar f l o w , the transition-detector i n s t a l l a t i o n on the Fiberglas covered wing was u t i l i z e d .
Tests were conducted a t speeds from a Mach number of 1.2 t o a Mach number of 2.0 a t a l t i t u d e s from 33,000 t o 36,000 f e e t . The free-stream Reynolds number varied from 1.3 t o 4.3 x 10 per f o o t . The maximum t r a n s i t i o n Reynolds numbers (based on free-stream conditions and t h e NACA RM "j8E28 distance t o the point of t r a n s i t i o n ) that were obtained on the Fiberglas t e s t area are shown i n figure 8.
Data obtained a t a l l angles of attack, from near 0 ' t o near loo, have been used t o construct the curves. A s can be seen, the m a x i m t r a n s i t i o n Reynolds number on t h e top surface of t h e wing varied from 6 6 about 2.5 x 10 a t a Mach number of 1.2 t o about 4 x 10 a t a Mach num- ber of 2.0. The trend on the lower surface is generally t o more laminar flow, w i t h the m a x i m t r a n s i t i o n Reynolds nwher varying from about 6 6 2 x 10 a t a Mach number of 1.2 t o about 8 x 10 at a Mach number of 2.0.
Although no attempt has y e t been made t o separate the e f f e c t s of t h e variables t h a t contribute t o t h e results presented herein, the r e s u l t s a r e encouraging i n that laminar flow has been obtained over extensive areas of a wing surface a t supersonic speeds w i t h p r a c t i c a l wing-surface conditions.
Further f l i g h t t e s t i n g should include investigations t o determine what e f f e c t s on the boundary layer are experienced when the leading edge i s a l t e r e d , when the angle of a t t a c k i s varied, when shock-wave-boundary- and when other f a c t o r s enter the problem l a y e r i n t e r a c t i o n takes place, a s important variables.
CONCLUS IClNS Chemical sublimation has been employed f o r boundary-layer-flow v i s u a l i z a t i o n on the wings of a supersonic f i g h t e r airplane i n l e v e l flight a t speeds near a Mach number of 2.0. The t e s t s have shown t h a t laminar flow can be obtained over extensive areas of the w i n g w i t h p r a c t i c a l wing-surface conditions.
I n addition t o the flow visualization tests, a method of contin- uously monitoring the conditions of t h e boundary layer has been applied t o flight t e s t i n g , using heated temperature r e s i s t a n c e gages i n s t a l l e d Tests were conducted i n a Fiberglas "glove" i n s t a l l a t i o n on one wing.
a t speeds from a Mach number of 1 . 2 t o a Mach number of 2.0, a t a l t i t u d e s from 35,000 feet t o 36,000 f e e t .
Data obtained a t a l l angles of attack, f'rom near Oo t o near loo, have shown t h a t the maximum t r a n s i t i o n Reynolds number on the upper sur- face of the wing varies from about 2.5 x 10 at a Mach number of 1.2 t o about 4 x 10 a t a Mach n u b e r of 2.0. O n t h e lower surface, the maximum e. 0.. e e.. e e. e. . e e.. e.
e . e . e . e e . . e . . e . .
e . e . . e . . e e. . e
e . e. e.. e .e e e . . e e. e - . : 6 NACA RM H58E28 transition Reynolds number varies from about 2 x lo6 at a Mach number of 1.2 to about 8 x 10 at a Mach number of 2 .O .
High-speed Flight Station, National Advisory Committee for Aeronautics, Edwards, Calif., March 20, 1958.
REFERENCES 1. Richardson, Norman R., and Horton, Elmer A.: A Thermal System for Continuous Monitoring of Laminar and Turbulent Boundary-Layer Flows During Routine Flight. IACA TN 4108, 1957.
Chemic31 Solids as Diffusible Coating Films for 2 . kin-Smith, J. D. : of Boundary-Layer Transition in Air and Water.
Visual Indications R. & 1.1. No. 2755, British A.R.C., Feb. 1950.
Visual Trsnsition Tests in Flight on a Griffith Suction 3. Gr3y, W. E.: Rep. No. Aero.2076, British R.A.E., Oct. 1950.
Wing, 16% Thick.
4. Seiff, Alvin: The Prospects of Laminar Flow on Hy-personic Airplanes.
NACA R M A58D25 , 1958.
5. Owen, P. R., and Randall, D. G.: Boundary Layer Transition on a Sweptback Wing. Rep. No. Aero 277, British R.A.E., May 1952.
AIRPLANE - TRANSITION TEST AREAS
CAMERA LOCATIO SHARP PRODUCTION WING / Figure 1 TRANSITION TEST METHODS CHEMICAL SUBLIMATION TRANSITION DETECTORS I \ TURBULENT GALVANOMETER DEFLECTION LAMINAR - 0 1 2 3 4 5 X, FT Figure 2 n n n m - r n w n m - r A T NACA RM ~ 5 8 ~ 2 8 IN-FLIGHT PHOTOGRAPHS TYPICAL 35'MM ENLARGEMENT ARTISTS CLARIFICATION Figure 3 INDICATION OF THE PRESENCE OF VORTICES Figure 4 CONFTDENTIAL NACA RM ~ 5 8 ~ 2 8 LAMINAR FLOW ON UNFINISHED WlNG M= 2.0, hp z 56,000 FEET a ~ 4 . 5 4 R=1.8x106 PER FOOT L.E. FLAP HINGE LINE LOWER SURFACE LEFT WlNG Figure 5 EFFECTS OF SURFACE CONDITIONS R= 1.8x106 PER FOOT FINISHED 8 PAINTED WING; M= 1.8 FINISHED WING; M = 2.0 ROUGHNESS. 2 5 p in. R O U G H N E S S = I ~ ~ I I ~ .
UPPER SURFACE UPPER SURFACE LOWER SURFACE LOWER SURFACE Figure 6 C QNF IDENT IAL Restriction/Classification Cancelled NACA RM ~ 5 8 ~ 2 8 COMPARISON OF LEFT-& RIGHT-WING TRANSITION M=1.8, R= 1 . 8 ~ 106PER FOOT RIGHT WING LEFT WlNG FINISHED 8 PAINTED FIBERGLAS TEST AREA ROUGHNESS =13p in. ROUGHNESS=7p in.
UPPER SURFACE UPPER SURFACE LOWER SURFACE LOWER SURFACE Figure 7 MAXIMUM TRANSITION REYNOLDS NUMBERS FIBERGLAS TEST AREA- RIGHT WlNG 10°>(r v O O 8\106 TOP SURFACE 8 x lo6 BOTTOM SURFACE R x 4 Restriction/Classification Cancelled