Addendum to: A theoretical and experimental study of the boundary layer flow on a 45º swept back wing

CoA Report No. 109 Addend

nCHKISCHE ••

N/LIEGTUIC;

Kanaai»traoi 10 1=

-THE COLLEGE OF AERONAUTICS

CRANFIELD

ADDENDUM TO

A THEORETICAL AND EXPERIMENTAL STUDY

OF THE BOUNDARY LAYER FLOW ON A 45'

SWEPT BACK WING

by

;

J. WALTON

Addendum t o R e p o r t 109 November, 1957»

T H E C O L L E G E O F A E R O N A U T I C S

C R A N F I E L D

Addendum to College of Aeronautics Report -109:

A Theoretical and Experimental Stud;-- of the Boundary Layer Fluw on a 4-5 Swept

Back Wing,

T^y

J . Walton

( P r e p a r e d u n d e r M i n i s t r y of Supply C o n t r a c t No. 6 / A i r c r a f t / l 4 0 0 ^ C . B , 6 ( a )

SUlM/iRY

College of Aeronautics Report 109 (Ref.l) descrr^es Flight Tests carried out on a swept back half wing of double elliptic section to investigate the nature of the boiondary layer flow, with particular reference to Bounday Layer Instability and subsequent transition,

The wing, which had a chord of 7ft.2" was mounted as a dorsal fin on the m.id upper fuselage^-of an Avro Lancaster, which enabled a Reynolds N'Cimber range of 0.88 c 10 - 1,92 x 10 per foot to be achieved.

There v/as some doubt about the validity of applying the results of these tests to wings of orthodox section because of the possible occurrence of wake instability associatea with the bluff trailing edge. This Addendum gives the results of a few check tests on the same wing with

a short trailing edge extension having a trailing edge angle of approxims.tely 12 . Unfortionately wing surface deterioration near the L.E, from mid semi span to the tip prevented conclusive i-esults being obtained but some evidence is presented to show that the results of Ref. 1 are not invalidated by the cho:^ce of section.

C0^rT5^^T3 P a g e IRTROmCTION 3 EXPERIMENTAL EQUIPMENT 5 2 . 1 . The A i r c r a f t 5 2 . 2 , The T e s t Wdng 5 2,3a I n s t r u m e n t a t i o n 6 2,4o Boundary Layer Combs and T r a n s i t i o n I n d i c a t o r s 6

2 , 5 . P r e s s u r e Leads 6 DETAILS OP TESTS 7 PRESENTATION OP RESULTS 8 DISCUSSION OF RESULTS 9 5 . 1 . S t a t i c P r e s s v r e D i s t r i b u t i o n 9 5 . 2 . Tuft O b s e r v a t i o n s 10 5 . 3 . Boimdary Layer Measurements 10

5 . 4 . Shape P a r a m e t e r and Boiondary Layer T r a n s i t i o n 11

CONCLUSIONS 12 REFERENCES 13 TABLES 14 & 15

FIGURES: 1 . Arrangement of A i r c r a f t and Wing. 2. Layout of Equipment i n A i r c r a f t , 3 . P l a n f orm of Swept Back H a l f Wing. 4 . S e c t i o n of Sv/ept Back Half Wing. 5 a . P r e s s \ i r e E r r o r C o r r e c t i o n s . 5 b . V a r i a t i o n of A i r c r a f t I n c i d e n c e w i t h I n d i c a t e d A i r Speed. 6 . A i r c r a f t Speecl/Test Wing I n c i d e n c e E n v e l o p e . 7 . P e r c e n t a g e c h o r d l e n g t h s i n t e r m s of d i s t a n c e from L . E .

COWTENTS ( c o n t d . )

2

-FIGURESï 8. S t a t i c p r e s s u r e d i s t r i b u t i o n , 9. Chordwise L o a d i n g s .

10. Tuft O b s e r v a t i o n s .

1 1 . Boundary Layer P r o f i l e s Root, (^/s 0.223

12. « " " Mid Semi-Span. ( ^ / s = 0.497) 1 3 . " " " T i p . ( ^ / s = 0.772

3

-INTRODUCTION

Ref. 1 describes the experiments made to investigate the nature of the boundary layer flow with partiaular reference to sweep instability on a ST/ept back half wing in flight. The wing which had a semi span of 8'6.5", a constant chord of 7'2" and a sweep back of 45 was mounted as a dorsal fin on the mid fuselage of an Avro Lancaster which enabled a Reynolds Number Range (based on 86" chord) of 6.4 x 10 - 13,7 x 10 to be obtained. The wing section was formed by two semi ellipses having a common minor axis of 13" S-t the maximum thickness, the fore and aft parts having semi major axes of 52" and 34" respectively.

The main conclusions of Ref. 1 may be sui-imarised as follows ;

(1) The use of the aircraft as a test vehicle was psrfectly satisfactory.

(2) The conditions for instability of the secondary flow are given by an equation of the type

iLSax ^ _^, ^20 < N < 160.

R 2 R 2

crit

(3) The measiired pressur's distributions were in close agreement with calculations based on the infinite sheared vdng and an equivalent source distribution for the zero incidence case (symmetrical section).

(4) Tuft observations showed that for both upper and lower surfaces within the incidence range 0 - 10 , three dimensional effects did not atihieve first order importance, thus permitting the use of strictly two dimensional techniques for boundary layer

4

-(5) No laminar flow was detected on either wing surface at Reynolds Numbers of 10,85 x 10 or above nor at incidences of 60 or above.

The reasons for using a dotible elliptiö dection were given in Ref.1

para, 2,1 and may be bfiefly ï-estated as follows

:-The three dimensional boundary layer instability phenomena

t

requires a study of the flow over and near to the leading edge of the wing at full scale Reynolds Numbers. Consequently if it is supposed that the flow in this region can be simulated by using a section with a foreshortened trailing edge representing a wing of much longer chord it is then possible to build a test wing with a mxoch larger distance between the leading edge and maximiJim thickness for the same actual wing area. Such a wing would enable much higher Reynolds Numbers (based on the distance from L.E, to max. thickness) to be

obtained for the same max. permissible loading than a wing of orthodox section. Thus the design of wing of Ref. 1 was based on this assijunption •

with an aotioal chord of 86" which, it was suggested, was representative

of an orthodox wing with a choid of 130", The latter dimension was referred to as the "effective" chora in Ref. 1.

In Ref. 1 paras. 2.1 and 2.3 the author uses the experimental rnd theoretical results to show that the above argument is erroneous as far as the simulation of flow conciitions is concerned. It is pointed dut that this does not invalidate the results but means that they may be compared to those on wings having similar pressure distributions.

In para. 2,1 it is also pointed out that no trace of wake instability • due to the bluff trailing edge was detected. It was the intention in

this note to show positively ;i\*iether or not wake instability was t encountered and in order to do so, a short V-section trailing edge

extension was fitted to the wing of Ref. 1 and a few checktests were made under the same conditions as those of Ref. 1.

5

-2 . EXPERIl^lENTAL EQUIPMENT

2.1. The Aircraft

The aircraft used for the tests of Ref. 1 (Avro Lancaster Mk,7 PA-474-) was used as the test vehicle for the present tests without further

modification,

2.2, The Test Wang

Tlie test wing of Ref. 1 was modified by the attachment of a trailing edge extension of light alloy skin stiffened by 7 light alloy ribs as described in Ref. 2. The above extension increased the actual chord of the wing by 24" from 86" to 110" as shown in Figs. 3 anti 4 and had a trailing edge angle of 12 . Five static pressure tapping holes flush with the skin surface at approximately mid semi-span were built into the extension,

The rea.son for not fairing the extension into the basic wing to give a smooth contouij was that the boundary layer was already turbulent well forward of the proposed wing extension jiancture and it was considered that the slight concavity in the wing contour would not affect the flow forward of the maximum thickness, beyond Virblch laminar flow had not been fietected,

The Boundary Layer Fence at the root was extended rearwards approximately 18".

Table 1 comparss the various geometric features of the wing with and without the trailing edge extension.

The Chordwise rows of static pressure tappings with 13 tappings in each row were on the port side only. These were at spanwise positions of ^/a = 0.223, 0.497 aJ^d 0.772 measured from the boundary-layer fence. The chordwise positions of each tapping is given in

Table 2. As the wing section was symmetrical negative incidence resull^s may be taken as lower surface values. The same applies to the boundary

layer measurements which were made on the starboeird side only, at the same spanwise positions.

6

-Before the present flight tests were begun, it was noticed that the wing s'jrface had seriously deteriorated on both surfaces near the tip and had slightly deteriorated at about mid semi-span. Nearer the root, the wing surface v/as not affected. The above deterioration appeared to be due to the filling in the joint (approx. 5.65" from the leading edge) between the mahogany lea,ding edge member and the light alloy skin,

being forced out and lifting the cellul.ose finish. This was most probably due to temperature and humidity variations between the two

series of tests. The wing surface was refinished, and checked by a strip light as described in Ref. 1 para.4.3, but no measurements of the surface roughness or waviness could be made as all the work had to be carried out with the vri.ng n.ounted on the aircraft. Unfortunately the present series of tests Fhowed that the attempted restoration was not completely successful and that further deterioration took place during the tests,

2.3. Instrumentat ion.

The manometer, F.24 cairera i n s t a l l a t i o n and yawmeter vera used without a l t e r a t i o n as described i n Ref. 1 p a r a s . 4 . 4 and 4.5 except t h a t water (vdth a purple dye) was used as the manomebric f l u i d . For t h e l a s t f l i g h t of the s e r i e s a p i t o t - i n - v e n t i i r i was i n s t a l l e d on

t h e starboard side of the fuselage symmetrically opposite the standard a i r c r a f t p i t o t head on the port side and a 0 - i 2" D i f f e r e n t i a l

P r e s s u r e Gauge mounted on the p i l o t ' s combing v/as connected between the two p i t o t s .

2 . 4 . Boundary Layer Combs and T r a n s i t i o n I n d i c a t o r s These were e x a c t l y as described i n Ref. 1 p a r a . 4 . 6 .

2.5. Pressure Leads.

The pressure leads from the boundary layer combs and transition indicators differed from those used previously (see Ref. 1 para.4.7) in tv/o respects. The 10 tube P.V.C. "tube tape" was dispensed v/ith, thus eliminating several joints in each lead and the 3 ran O.D. neoprine

/

-tubing was taken direct from the boundary layer combs and T.I's into the fuselage and thence via a short length of 5" O.D. rubber tubing

to the manometer. As before, the leads were made long enough to traverse to the leaiüng edge but the excess v/as coiled and stored in the fuselage. To reduce tu a ndiiimum the interference of the pressure leads v/ith the aix-flow over the wing^ the tubing was taken as far aft as possible and thence down the cson-ja-vdty at the joint between the basic wing and the trailing edge extension^ over the aftmost point of the boiondary layer fence, and into the fuselage. It was fo\jnd convenient to form the 13 neoprine leads from each boundary layer comb into groups by sellotape straps at about 12" intervals along their lengths. The attachment of the boundary layei' combs, T.I's and the pressure leads to the wing was done entirely with sellotape and f;lnally the "leading- ecge" of the tubing running down the •vi±ng (spanwise) and faired over with sellota^pe.

3. DETAILS OF, TESTS

The static press'ore distribution over the basic wing previously measxared 'v/as from the leading edge on^-y to the maximum thickness, hence the static pressure over a large section of the Tving was in doubt. Consequently the speeu/'incideiice envelope of the test wing with the T.E. extension was more severely limited than for the tests of Ref. 1 see Fig, 6. However, this was of little consequence as it was possible to repeat all the conditions in which laminar flow had previously been detected.

Only four flights were possible before the aircraft was grounded for a major over-haul at v/bich point the present tests were terminated. On the first flight the pressure distribution was obtained and on

the subsequent three flights the boundary layer profiles and transition indicator readings v/ere obtained at 39", 26" and 13" from the leading edge at the same spanwise positions as those of Ref. 1.

All tests were perfomed at 10,000' at 90 kts., 110 kts., and 158 kts,, or maxrmum within the envelope with the test wing incidence varied by 2 increments each way. As before the above flight conditions

*"a Q 'T»

give Reynolds Number 0,88, 1.08 and 1.55 x 10 per ft.

As in Re-f", 1 the boundary layer combs were traversed chordwise at the spanwise positions (measured from the boundary layer fence) given by ^/a - 0,223, 0,497 and 0.-772. The transition indicators were traversed chordwise at '^/•r. = 0.086, 0,360, 0.635 ai^d 0.909.

For the last flight the wing was extensively covered with wool t\ifts and observations were made from the DoH. Dove G~ALVF as described in Ref. 1 para. 5.5.

It had been noted that in both series of tests (e.g. Ref, 1, Fig, 34), the total head, at low speeds^measur'od outside the boundary

layer on the wing was almost invariably higher than tliC free stream total head measured by the aircraft pitot system evea vrhen the static pressure error had been taken into accoiont. Hence the a/c pitot pressure was checked by the instrumentation mentioned in para, 2.3 and the system was found to have quite a large pitot pressure error at low speeds, as shown in Fig, 5a.

4. HffiSFJITATIGN OF RESULTS

The method chosen to present the results is one that enables a direct conparison to be made with those results of Ref. 1 which were obtained under the same conditions. With this end in view, the layout of the experimental results closely folloT/s that of Ref. 1 ; all f-all lines applying to bhe present results, a?l chained lines being the

corresponding results of Ref. 1 unless otherwise stated. As the present tests apply to the wing with a greater chord than that of Ref. 1, for clarity all chordwise positions are given in terms of the distance x from the leading edge in the streamwise direction in inches. Any distance x may be easily converted into the percenta.ge length of either chord by reference to Fig. 7.

Fig, 8 shows the mean chordwise pressure distribution from the leadijig edge to maximum thickness over the sT.tept back half wing as obtained from the surface tappings. The Cp shovm is the m.ean for three

9

-speeds for a = 0 - ~ 6 , two -speeds for cc = * 8 and the actual

C for oc = i 10°, at 90 kts, (R.N. = 0.88 x 10 per foot). For clarity the pressure distrmbuticns previously obtained (ref. 1 Fig. 25) are omitted. However, Fig, 9 conipares the chordwise loading A C on a percentage of chord basis, i.e. the results of Ref. 1, Pig, 26, are plotted against 39^0 and the present resiilts against V'x» ^ common scale being used for x/c and x/ ,

'^ ^ o ' c

Unfortunately no photographs of the tufts were obtained but Fig. 10 shows a sketch of the tuft pattern on the upper surface

at 10 incidence based on the description of an observer in the Dove.

The velocity profrles are compared in Pigs. 11 and 13 and transition fronts in Pig, 14.

5. DISqiSSIGN OF RESULTS

5,1. Static Pressure Distribution

When the results of Ref. 1 Fig. 25 are compared with those of ï'ig. 8 of the present note it is seen th/it, for all incidences except +6 and +10 , the static pressure coefficient is, generally speaking,

increased by about 0.02 over most of the chord by the addition of the trailing edge extension. This value is within the limits of accuracy of the experimental procedure as shown in Appendix 1 of Ref. 3. The slight decrease in static pressure at oc = +6 and oc = +10 is most likely due to the a/c having had a small angle of sideslip which would increase the incidence of the wing.

The chordwise loading graphs are more informative and when plotted against the distance x show that the loading is moved slightly afit wi.th the trailing edge extension. This is, of course, what might be expected considering the simple assumption that the pressure distribution on an aerofoil section at incidence oc is obtained from the sum of the pressures on the section at zero incidence plus a flat plate at

imcidence a . Y/hen plotted as in Fig. 9 it is shown that the trailing edge of the basic wing section of Ref, 1 is well defined; the effective

10

-trailing edge being, in fact, approximately 4^ c behind the actual trailing edge.

5.2. Tult Observations

Although the sketch of the tuft pattern shown in Fig. 10 is only qualitative as suggested in para, 4, particularly in the region of the trailing edge extension, it is seen that the flow is approximately chordv/ise in the region in v/hich boundary layer measurements were made. Hence the use of the boundajry layer combs which strictly are for use

in tvro dimensional flow only, is not invalidated.

5.3. Boundary Layer Measurements

It is immediately apparent from the velocity profiles. Pigs. 11 - 13, that at the tip and mid semi-span stations transition occurred further forward in the present tests than in those of Ref. 1 due to the reasons given in para. 2.2. At the root station, transition appears, in general, to be slightly delayed by the addition of the trailing edge extension. The agreement between the turbulent profiles of Ref. 1 and the present note is in accordance with the observations that, on a flat plate, the thickness of the turiDulent boundary layer is dependent on the distance from the stagnation line and not on the point of transition. The above agreement also shows that experimental conditions caii be consistently reproduced to a high degree of accuracy with the present technique.

From an examination of the velocity profiles it will be seen that, near the surface, the heights at which the measured velocity ratios have been plotted do not correspond with the nominal tube heights given in Ref, 1 Fig. 13. The reason for this is that in general it VIELS not found possible to draw a smooth set of profiles using the nominal tube heights and consequently they were measured on each comb at each chordwise position,

-. n

.-5.4-. Shape Parameter and Boundary Layer Transition

As only three chordwise positions of the siorface total head distribution were availrble, it was not possible to obtain a complete picture of the bo\mdary layer transitions by the same method as in Ref. 1, This method used the total head rise shown by a creeping

surface pitot when passing from a laminar boundary layer to a turbulent one, although in some cases this total head rise Viras not well defined.

However, by plotting the Shape Parameter H against x for the velocity profiles of Ref. 1 clearly defined curves were obtained and taking the end of the transition region to be the point where H attained a steady value (approximately 1,5, the value fbr a turbulent boundary layer) it was possible to redetermine the transition points for the original wing with more certainty. These points which as in Ref. 1, were taken to be the end of the transition region, were in

general in good agreement with those of Ref, 1, Pig, 35 which were mainly determined from total head measurements. Then, assianing that the

curves of H ~ x for each incidence and speed would be similar, over a small region, for the wing with or without the extension, the above curves were given a small chordwise displacement to pass through the points for the wxng with the extension, and hence the transition points for the win^ with the extension were found,

Using a similar technique for the total head measurements of the transition indicators, the transition fronts for R = 0,38 x 10 /ft, shown in Pig. 14 were determined. For consistency the transition fronts shown for the original wing are those redetermined by the shape parameter method. No fronts for R = 1 , 0 8 x l 0 /ft are shown as

transition appeared in g:sn.er,il to be anead of x = 13", "the most forward position at which measurements were made vd.th the trailing edge extension,

12

-6, CONCLUSIONS

Y/hile the tests described herein were not conclusive due to the short time available and to surface deterioration towards the wing tip, there is sufficient evidence to suggest that the trailing edge of the swept back half wing of doubl;e elliptic section, described in Ref, 1, was in fact fairly W3ll defined and was not likely to have given rise to waka instability.

It is thbrefore suggested that the results of Ref, 1 may well be representative of similar wings with orthodox sections,

In general transition movements over the forward part of the wing do not appeal' to be unduly influenced by the aliape of the fairing aft of the maximum thickness to the trailing edge,

REPEREI^'CES

1, Burrows, F.M,

2, Department of Aircraft Design. C. of A.

3. Burrows, P,M.

A Theoretical and Experimental Study of the Boundary Layer Plow on a 45 Swept Back Wing,

College of Aeronautics Report No,109.

Supplement to Addendum to Type Record Avro Lancaster PA.474. College of Aeronautics, 1956.

Characteristics of the flow field over the mid upper fuselage of Lancaster PA.24-74-. SYMBOLS c X

y

R

X

Semi span of test wing (B.L.Pence to tip)

ohord of original wing

chord of wing with T.E. extension

distance along chord from L.E.

distance spanwise from B.L. Fence

Reynolds Number

Secondary Flow Reynolds Number

14

-TABLE 1

Comparison of t h e Geometric P a r t i e t J l a r s of t h e wing, w i t h o u t and w i t h t h e t r a i l i n g edge e x t e n s i o n .

Semi-Span

(Prom B.L, P e n c e t o T i p ) Chord

Wing A r e a of Half Wing (Outboard of B.L.Pence) Aspect R a t i o T a p e r R a t i o Max. T h i c k n e s s Thickiiess/Chord R a t i o D i s t a n c e t o Max.Thickness from L . E . a l o n g c h o i d 52" 52" D i s t a n c e t o Max. T h i c k n e s s a s % of chord 6 0 . 5 4 7 . 2 5 Without E x t e n s i o n ( a s i n R e f . l ) 8 7 . 5 " 86" 5 2 . 2 5 7 s q . f t . 2.035 1 13" 0,151 With E x t e n s i o n ( a s i n p r e s e n t n 8 7 . 5 " 110" 6 6 . 8 4 s q . f t 1.591 1 1 3 " 0 . 1 1 8

15

-TABLE 2

The p o s i t i o n of s t a t i c p r e s s u r e t a p p i n g s i n t n e swept back h a l f Vvdng. Spanwise p o s i t i o n s : ^/s = 0 . 2 2 3 , 0.497 & 0 . 7 7 2

fo

Hole No. 1 1 a l b 1o 2

3

4

5

6

7

8 9 10 11 12 13 D i s t a n c e from L . E . a l o n g c h o r d i n c h e s 0 0 . 3 2 5 0 . 6 5 0 0.975 1.3 2,6 5 . 2 7 , 8 1 0 . 4 13 1 9 . 5 26 3 2 , 5 39 4 5 . 5 5 0 . 6 X %

c

0 0

0,378

0,756

1-135

1.512

3.03

6,04

9.06

12.10

15.12

22.70

30.25

37.80

45.40

52,9

58,8

0

0.29e

^.591

0.887

1.183

2.365

4.73

7.095

0.45

11.83

17.72

23.65

29.55

33.45

41.3

46

Holes numbered l a , b and c were i n the neoprine tubing l e t fnto the

leading edge, see Ref. 1, para, 4 . 3 . Each hole was used in turn and

then f i l l e d with beeswax.

rtCHNlSCHE HO^..^:>-^;-^-^

VLIEGTUIGBOUWKUNI^^ Kanaalstraat 10 - I>E"T

8WKPT BACK WlUg,

FIG. I . ARRANGEMENT OF AIRCRAFT fc WING.

WUOT'» » I O « 5 L I P INDICBTOH.

6-5"

96"

THREE ROWS OF STATIC PRESSURE TAPPINGS. (PORT SIDE ONLY) AND LINES OF BOUNDARY LAYER COMB TRAVERSE, (STBD SIDE ONLY)

BOUNDARY LAYER FENCE. \

TRANSITION INDICATOR TRAVERSES SHOWN THUS

•yur\ CI I D C V - » O T I I \

FWD. SUPPORT

PLATE 86 REAR SUPPORT PLATE.

FIG. 3. PLANFORM OF SWEPT BACK HALF WING

BO««ARY L/(fER FENCE MAX

THCKNESS

A _{BOUNDARY LAITER FENCE}

EXTENSION

BO lOO I20 I40 I60 ISO 2 0 0 V„ KNOTS

FIG. 5a PRESSURE ERROR CORRECTION CURVES FOR

LANCASTER PA 474 12 to AIRCRAFT g FUSELAGE INCDENCE 4 DEGREES \ \ \ \ \ \ . " -- 2

SO lOO I20 MO I60 I80 2 0 0 I.A.S. KNOTS

FIG. 5b. VARIATION OF AIRCRAFT FUSELAGE INCIDENCE WITH SPEED AT 44,OOOIb A.UW.

lO ALLOWABLE TEST WING 6 INCIDENCE DEGREES \ \ iWITH • \EXTE

v

FRAILIM MSKDN S EDGE ^ ^ - ^ \ \ \ o \ ^IGINAL V \ \ WING ^ ^ \ \ \ \ N S V

6 0 lOO I20 I40 I60 I80 2 0 0 220 240 260 280 3 0 0 E, A S KNOTS.

0-9 0-8 0 - 7 0-6 2C. Co 0-5 0 - 4 0-3 0 2 0 - 9 0-8 0-7 0-6 0-5 X C 0-4 0-3 0 2 10 2 0 3 0 4 0 5 0 6 0 7 0 8 0 9 0 lOO MO X " FROM LEADING EDGE

(Kv

l x

-1 r

r ^-^ '

X

r^ 4 ° •s a • Z-B-2° ^ * • O 2 ^t——^ 0 o-— o-— - " ^ ' ^

j^r5==

o 3 DISTANCE BEHIND ' 8—

1^ — r

O 4 L.E. - IN. • - n

—r:t===

" — " " • * - — o 5 -o O 6

FIG. 8a. MEAN STATIC PRESSURE DISTRIBUTION

R = o 88 X 10* — I 55 X l o ' / F T

/^ I \ \ \ \ \ \ \

kV

\ ^

V

\ ^

W

^ ^ ^ ^ ^ ^ ">i. ^ ^ ^ ^ " ~ ~ — ~ ? ^ o^ „ 4° 1^1 B-cP__<^ < ^ O 2 o X — " ' • o 3 DISTANCE BEHIND - ^ "-* 1 J u , » « -o 4 L. E. — IN. . . * • " O • ^ " + ' o 5 -o-X -*--** p 6C

FIG. 8 b . M E A N STATIC PRESSURE DISTRIBUTION

B = O BB « l O * — I SS X 1 0 * ; F T 4- = o 4 97

- 1 8 -1-6 -1-4 - 1 - 2 - 0 - 6 - 0 4 - 0 - 2 • O 2 O 4 0 - 6 2 0 3 0 DISTANCE BEHIND L E. — IN.

6 0

FIG. 8c. MEAN STATIC PRESSURE DISTRIBUTION

R - O - 8 8 K O * — I 55 Xio'/FT ' O 772

% C H O W L

T- - 0-223 f = 0-497

R=Oe8x l o ' - l-SSxlO^/FT

FIG. 9. MEAN CHORDWISE LOADINGS.

FIG. IQ TUFT OBSERVATIONS UPPER SURFACE oC= IO° \ \

r\

'\\\ ' \\ \ * ' \ ^ *^ V \ \ ^ ^ "^^^

c ^

is

) -^ -ft' « «. - s ' ^ «-^ » "~~-lO JO K) « o K) W X CHOno -J - o 772

o -^s: ( ; " ^ : iï ; • ^ o V ^ \

1

tï-V

^ ^

'1

^ •o CM II >-< O u N { 0 to LU J u. O O O _ l -^ ?—-^'^ zr \ ^ V

'1

~"—^**^

^\1

'^ \ «ï N V "ol -^ ^

S;

\ ? | • - ^ \ i %\ [ O m <J)

8 ° _ _ , jl^^ _ J ! _ = : 2° o° _ • ^ j r ^ 'V"^ i i ^ - - ^

-A

\

A

Y i \ ₁ 7 1 / / fl - >y \k < / ƒ / ƒ / / / j \ y / \ y / f / / y \ y ƒ '^""""'^ A : > ^ 0-6 0-8 ORIGINAL WING WITH TRAILING EDGE EXTENSION 1

^''^y

6V^" 1 if / / / / / / / y /

J ^

y o* „^^ / / 1 if 1 / / / / 1 1 1 l\ 11 ' \ 1 '>\ ij I i

' 7

'' ' i / / 0-2 0-4 0 8 l O u 3:=i3 X = 2 6 X=39 (b;) R = l - 0 8 x l O ' / F T O ( - 0 ° - 8°

y o" -2» - 4 ° -4° - 8 ° ^'^ ^^^ ^.' ƒ / y / /y' ) y/ y / y/ 0 6 0-8 y" 0 3 0 2 O

__S'^

f f / y

^y.

~^1-^ _ : £ - ^ ^-£^^ II /

y

1 1 1 1 1 1 1 1 ij 1 Ij 1

j/l

n

1

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