The Distribution of Pressure over the Surface of Wings of Small Aspect Ratio

KomaalMraat 10 - DWHT REPORT No. 52

2 6 MEI 1952

THE COLLEGE OF AERONAUTICS

' CRANFIELD

THE DISTRIBUTION OF PRESSURE OVER THE

SURFACE OF WINGS OF SMALL ASPECT RATIO

by

W. S. D. MARSHALL, Dip.Aero (Huil), A.F.R.Ae.S. of the Department of Aerodynamics.

This Report must not be reproduced without the permission of the Principal of the College of Aeronautics.

TECHNISCHE HOGESCHOOL

VLIEGTUIGBOUWKÜNDE Konaalstiaat 10 - DELFT

REPORT w. 52 2 6 MEI

I95J

F e b r u a r y . 1952

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

The Distribution of Pressure over the Siarface of Wings of Small Aspect Ratio

-by-W.S.D. Marshall, D.Ae.(Hull), A.F.R.Ae.S.

The distributions of pressure over v/ings of aspect ratio 1.5 and 0.5 have been measured for a range of incidence up to and Inoluding the stall at varioxis angles of yaw. This report

presents a detailed analysis of the results at two incidences coTrrespending to l/4 and 3/4 of the stalling incidences approxi-mately. Direct measurements of lift and pitching moment have also been made, and the results compared with the results of theory and previous experiments.

The analysis shows

that.-1. Regions of high suctions near the tips assiane greater importance as the aspect ratio is reduced. This tip suction rapidly increases in intensity with increase in incidence.

2. Apart from regions near the tips the spanwise distribution of load becomes more nearly elliptical vdth decrease in aspect ratio.

3. The effect of a positive sideslip is to skew the span-wise load grading ciirve and to produce a negative rolling moment. This effect is more pronounced at small aspect ratios.

A-. •..

MEP

* Much of the experimental work, upon which this note is based, was performed by Messrs. E.G. Havard, E.F. Lawlor, A.Lightbodj', and A.O. Ormerod in 194-8.

4. Comparison between the lift coeffióients obtained by direct meastirement and from the pressure distributions shows reasonable agreement and the variations of lift c\jrve slope with change in aspect ratio are in agreement v/ith the results of other workers. ' Fin'ther, a theoretical curve due to Wieghardt

shows close agreement with the present experimental values. (2)

5. The method developed by Flax and Lav/rence , based on a modified slender body theory, for estimating the position of the aerodynamic centre is found to be in reasonable agreement with the experimental results.

-3-LIST OF CONTENTS 1.O Introduction

Page No.

4

2.0 D e t a i l s of t e s t s 4

3.0 ResiiLts 5

4.0 Discussion 7 4.1. The Distribution of Pressure 7

4.2. Spanwise Load Grading 8

4.3. Tip Effects 8

4.4. Spanwise Variation in Position of

local Centre of Pressure 9 4.5. Aerodynamic Derivatives 9

4.5.1. Normal Force with Respect to

Sideslip 9 4.5.2. Pitching Moment due to Sideslip 9

4.6.

4.5.3. Rolling Moment due to Sideslip 10 Direct Measurements of Lift and Pitching

Moment

4.6.1. The Lift Curves

4.6.2. Pitching Moment Results

10

10

11

5.0 C onclusi ons

11

List of References

12

1.o Introduction

Recent developments have stimulated interest in the characteristics of wings of small aspect ratio. The tests described here were concerned with the pressure distributions, normal force coefficients and the effects of yaw on wings of aspect ratio 0.5 and I.5. The wings in this instance were

rectangular with constant symmetrical sections 12 per cent thick. Prom the pressure distributions the spanwise loading and the positions of the local centres of pressure have been determined as well as the derivatives of the normal force, pitching and rolling moments with rate of yaw.

The pressure distribution measurements were made in the No. 2 Yfind Tunnel at the College of Aeronautics diiring Jtine and Jvily of 1948, and later a few balance measurements of lift and pitching moments were made in the No. 1A Wind Tunnel.

Dioring the preparation of this report, the results of (1)

some Swedish experiments came to hand. The work described in this reference covers much the same ground as the present experi-ments, and, where possible, comparison between the two sets of results is made.

2.0 Details of Tests '

The models were made of laminated mahogany having the symmetrical section shovm in Fig, 1. (The ordinates are given in Table l). The wing tips were half bodies of revolution. The chord of both wings was 15 in. and the spans were 22|-in and 7^in., exclusive of wing tip fairings. The values of aspect ratio

quoted apply strictly to the wings without tip fairings, the effective additional area and span due to the tip fairings having been neglected. The models were mounted from an overhead turn-table by a ccmbination of struts and wires, the arrangement of which can be seen in Figs. 2a and 2b.

Small bcre tubes of a pliable plastic material were inlaid into chordwise slots cut in the top and bottom surfaces of the wing from the leading edge of the wing back to 85 per cent of the chord. The tubes were faired over with beeswax, and the model french polished. The upper surface tubes extended around

the nose of the aerofoil and back along the lower surface to about 20 per cent of the chord. Lengths of rubber tubing trans-mitted the press\ares to a vertical multitube manometer.

The pressure orifices were formed by drilling holes in the tubes at a single chordwise position, and when the pressures at this position had been recorded these holes were sealed with plasticene and a fresh set of holes drilled at a different chord-wise station.

The chordvri.se positions of the orifices were as follows. -Upper Siorface yi/c,

0, . 0 1 , . 0 3 , .05, .07, .10, . 1 5 , . 2 5 , . 3 5 , . 4 5 ,

. 5 5 , . 6 5 , . 7 5 , . 8 5 .

Lower Siirface

x/c

.02, .05, .10, .15, .25, .35, .45, .55, ^^5y .75, .85.

The tunnel speed for all tests was 120 ft./sec,

corresponding to a Reynolds nimiber of 0.95 x 10 , and readings of pressure were taken at the following incidences (uncorrected).

Aspect Ratio 1.5

_10^

15° 20° 22° ( s t a l l e d )

Aspect Ratio 0.5

6° 18° 25° 28° (stalled)

The

angles of sideslip, denoted by " ^ , were 0 , 1 0 and 20 to starboard for both wings.

For the second series of tests each wing was suspended from the three component balance of the No. 1A tunnel and measure-ments of lift and pitching moment were taken over a range of nominal incidence from «-4 to 10 .

3.0. Results

The pressure coefficients were plotted against chord-wise positions for each of the wing attitudes tested, and smooth

ciirves were drawn through the experimental points. A selection of the resulting isobars is given in figs. 3 1 4

-The area enclosed by the curves giving the upper and lower surface pressure distributions along the chord for any given spanvri.se position represents the normal force per unit span acting on the aerofoil at that spanwise position. That is, the local normal force coefficient, Cj^ is given by

°NF = yi/c = 1.0 up x/c = 0 (C - C ) d /S> P Pn Vc/ ^upper -^lower ^ ' /where .,.

where x is the distance aft of the leading edge measin^d along the chord line.

On the wing of larger aspect ratio the lower surface pressures aft of the quarter chord point could not be measured at two of the sparavise positions, because the pressure holes would have been in the wake of the supporting struts. As there are no steep pressure gradients in this region, it was thought that little accuracy woxold be lost if the pressures there were obtained by interpolation from the pressures measured at the adjacent stations.

In the analysis it was necessary to draw complete span-wise load grading curves to obtain the total normal force

coefficient acting on the wing. This involved a certain amo\mt of extrapolation across the wing tip fairings. Since the normal force per unit span may be expected to be continuous and to

reduce to zero at the tips this was the qimntity extrapolated.

Normal force per xmit span can be represented by the non-dimensional coefficient C-„ x — . This coefficient has been plotted as the ordinate of the spanwise load grading curves in figs. l6 and 17. Over the parallel portion of the wing c n OJ, , so that an ordinate in this portion is simply the local normal force coefficient. In figs. 16 and 17 the ciirves are shovm slotted v»rhere the extrapolation is doubtfxjl.

The local centre of pressure was found as the point on the chord line through which, for the section considered, the resialtant normal force acts. The position of this point was foiond graphically from the chordwise pressure distribution. The variations of the local centre of pressijre are shown plotted in figs. I8 and 19.

By integrating the spanvidse load grading curves the variation of normal force with angle of sideslip was obtained; and the variations of pitching and rolling moments with yaw vrere obtained by integration of curves giving the moments of the normal forces about the leading edge and the centre line respectively. The results are shown in figs. 20a,b and c.

The main correction due to tunnel constraint is a change in the mea.svred. angle of incidence at a constant wing lift

coefficient. This correction has been applied to the readings although some doubts exist as to its validity on acco\int of the large ratio of wing chord to tunnel diameter encountered in these tests. It is for this reason that, where possible, the overall normal force coefficient has been used as a parameter in the

-7-presentation of the res\ilts in preference to the angle of inci-dence.

4.0 Discussion

4.1. The Distribution of Pressure

The distributions of pressxire over the wings are shown by lines of constant pressure coefficient (C ) for the upper and lower surfaces in figs. 3 - 1 4 . A striking feature shown by these isobars for the larger incidences is the high suctions

occurring at the tips, near the trailing edge on the upper surface and, to a lesser extent, at about the mid-chord position on the

(l 2) lower s-urface. This phenomenon has been noted by other workers * and has been attributed to a spiral motion of the air fran bottom to top surface around the tip fairings related to the component of

(2)

flow normal to the plane of the wing.^ ' This spiral motion is said to result in a trailing vortex springing frem the wing sur-face which is separate in character, but which may become

indistinguishable from the normal trailing vortices of lifting line theory. Certainly, tuft observations reveal marked cross flows of the type described, but the lift distribution and the associated vortex flow are related effects resulting from some more fundamental cause, which must be sought by further and more detailed investiga-tions of the character of the boundary lajrer flow in the region of a vring tip.

Yawing the wing intensifies or diminishes these regions of suction according to whether the spanwise component of the flow due to yaw reinforces or opposes the inflow or outflow. This effect is clearly shown for 7^ = 20° (figs. 8 and 14), where the region of higher suction is confined to the leading tip on the upper svnrface and to the trailing tip on the lower surface.

In the case of both aspect ratios tested the regions of higher suction occurring on the top surfaces extend inboard of the tips for a distance approximately 15 per cent of the chord and it can be inferred that they are roughly of the same order of intensity at the same overall normal fcarce coefficient. This is consistent with the hypothesis that the tip effect is not a

characteristic solely of small aspect ratio wings and indeed something similar has been found 'to occur on a wing of aspect ratio 6. On this hypothesis, the.intensity and extent, expressed as a fraction of the wing chord, of the higher suction at the tips are approxiraately independent of aspect ratio but, as the aspect ratio decreases, these tip effects become relatively more important,

slnce they extend over a larger fraction of the wing surface.

As already noted the wings tested were rounded at the tips by a surface obtained by revolving the aerofoil profile

about a streainwise axis through the tip. This method of roimding produced a wing span which increased with distance from the

leading edge up to the location of maximum thickness of the section (in this case at 0.3 of the chord) and then decreased to the

leading edge span at the trailing edge. According to Jones, for thin wings of very low aspect ratio, the lift is zero over the entire portion of the wing over v/hich the span is decreasing if the trailing edge is sharp there and the Kutta-Joxikowski con-dition applies. With the tips roxonded as in these experiments the Kutta-Joiikowski condition is not applicable there, but then the simple slender body theory, on which Jones* theory is based, predicts negative contributions to the lift where the span is

(5) decreasing. This is in agreement with more exact theories,

which predict small or negative lifts in that region. Inspection of the isobars given in figs. 4 and 10 and the spanivise pressure distributions of fig. 15 shows that the present experimental

resTolts support this prediction; it vd.ll be seen that the suction at the tips is more or less cancelled out by the region of pressure over the central part of the span towards the trailing edge.

4.2. Sparori.se Load Grading (figs. I6 and 17)

Apart from the tip effects the loading distributions show an increasing tendency towards the elliptical distribution vri-th reduction of aspect ratio, in agreement with theory. Ai remarked above, however, the tip effects become more dominant

with decrease of aspect ratio. The curves for both If = 10 and 20 show that yaw tends to skevvr the load grading so that the lift is increased at the leading tip.

4.3. Tip Effects

At zero ynw and particularly at the large incidence the high suction region on the upper surface near the tips results in a local maxhavm or peak in the load grading curve.

(1)

Holme ^ '^also noted this effect. With the wing yawed this peaking is more pronounced at the leading tip, whilst at the trailing tip it tends to disappear. These changes in loading at the tips are far more marked at high than at low incidences.

TECHNISCHE HOGESCHOOL

VLIEGTUIGBOUWKÜNDE Kanaalstraat 10 - DELFT

-9'

4 . 4 . Spanvrise Variation i n P o s i t i o n of Local Centre of Pressiare

(Figs. 18 and 19)

The method of estimating centres of pi^'essure is less accurate than that of estimating the normal force coefficient, C.j^ ,

because.-a) frictional drag effects vrere neglected

and b) pressures near the trailing edge were obtained by extrapolation, the rearmost reading of pressure having been made at 0.85 of the chord. These pressxxres have no little bearing on the centre of pressure position.

Nevertheless, the overall trends revealed by the data are of interest.

It appears that at zero yav;' and small angles of inci-dence the centre of pressure tends to move forward as the tip is approached. With increase in incidence, however, the suction at the tips near the trailing edge resxjits in a rapid rearward movement of the local centre of pressure there. It will be noticed that the cxarve for aspect ratio 0.5, C ^ = 0.11, ']t- = 0 lacks symmetry, this lack of symmetry is a measure of the relia-bility of the deduced positions of the local centres of pressxire in an extreme case where the accxaracy can be expected to be least.

With increase in the angle of yaw, the centre of pressxire on the leading tip moves back, and moves forward on the trailing tip. In general, it can be seen that the centre of pressxjre moves rearward with increase in incidence,

4.5. Aerodynamic Derivatives (Fig. 20)

4. S.l. Normal Force with Respect to Sideslip (z ) (Pig.20A) For both aspect ratios the normal force coefficient

G-m remains constant vrith change of yaw at the smaller incidence. yaw.

At the larger incidence there is a slight increase in C^^ with

4.5.2. Pitching Moment due to Sideslip (m ) (Fig. 20B) The method for estimating m f rem the pressxare

distributions is less accurate than that for Jt and z , since, in this case, an accxjrate determination depends on an intimate

knowledge of the spanwise pressxire distribution near the tips. The resxalts indicate that the pitching moment coeffic-ient C becanes less negative for the smaller values of yaw and increases again with fxirther increase in the angle of yaw. This may be attributed to the fact that for small angles of yaw the rate of bxiild up of sxiction at the leading tip is less than the corresponding rate of decrease at the trailing tip. At the larger angles of yaw, however, this effect is reversed.

4.5.3. Rolling Moment due to Sideslip (i ) (Pig.20C) The skewing of the spanwise load grading cxjrves produced by the sideslip caxises a considerable rolling manent. The rolling moment derivatives obtained in the present experiment are compared

(6)

in the following table with a semi-empirical law quoted by Levacic; based, however, on data relating to vrings of larger aspect ratio than those considered here.

Nominal Aspect Ratio High Incidence - t /c.|^ Low Incidence - -*>/C^

Levacic

[ret. 6)

1.5

0.25 0.10 0.25

0.5

0.40

0.46

0.72

Some measxire of agreement is obtained with Levacic's formxjla for the wings of higher aspect ratio, but for the smaller aspect ratio his formxala considerably over-estimates Jf

4. 6. Direct Measxarement of Lift and Pitching Moment 4.6.1. The Lift Cxjrves

Fig. 21 shows the lift coefficient obtained by balance measxirement plotted against incidence corrected for the effect

of txmnel constraint. Also shown in fig. 2 are the resxilts obtained from the pressxare distributions. The agreement is thought to be satisfactory, bearing in mind that some

extra-polation v/as necessary in determining the normal force coefficients, The variation in lift cxirve slope at zero incidence

with change in aspect ratio is shown in fig. 22. The resxalts of /this ...

-11-(l 7) this experiment show fair agreement vri.th those of other workers. * ' Calcxilations ' ' 'of this variation with aspect ratio based on different theories are also shown in fig. 22, and it is seen that the three cxorves are almost identical. If choice mxost be made, the cxjrve due to Wieghardt (ref. 8), which assxmes an elliptical load distribution, woxild seem to offer the best agreement with the present experimental valxies.

4. 6. 2. Pitching Moment Resxilts

These resxjlts are presented in fig. 23 in the form of - -'L °^^ves. For the balance measxirements pitching moments C - C,

m I

xvere measxired about an axis through the quarter chord point. The position of the aerodynamic centre has been determined by measxare-ment of the slope of the cxarves at C^ = 0,

Pig. 24 shovra a cxarve derived by the method of Flax and Lawrence (ref. 2) based on a slender body theory, where, however, the tips are considered separately from the rectangular portions of the wings, and the resulting loadings are added. The agree-ment between this cxjrve and the measxH"ed resxilts is satisfactory.

5. Conclusions

The main conclxisions can be sxjmmarised as follows.-a) Tip effects in the form of regions of high suctions on the upper sxjrface near the trailing edge and on the lower sxarface near the mid chord position become relatively more important as the aspect ratio is reduced. The upper sxjrface suction region rapidly increases in intensity as the incidence is increased. b) Apart from regions near the tips, the spanwise

distribu-tion of load becanes more nearly elliptical as the aspect ratio is decreased.

c) The effect of sideslip is to skew the spanwise load grading curve so that a positive sideslip produces a negative rolling moment. The region of high sxjction on the upper sxjrface at the tip becomes intensified by yaw, whilst that at the trailing tip becomes reduced. The effects are more pronoxmed with

reduction in aspect ratio.

d) Comparison between tlie lift coefficients obtained by /direct ...

direct measxjrement and fron the pressxjre distributions shows reasonable agreement, and the variations of lift cxjrve slope with change in aspect ratio are in agreement with the resxilts of other

(l 7)

workers. ^ ' '' Fxjrther, a theoretical cxjrve for this variation (8)

due to Wieghardt shows close agreement with the present experi-mental values.

(2)

e) The method developed by Flax and Lawrence, ' based on a modified slender body theory, for estimating the position of the aerodynamic centre is foxmd to be in reasonable agreement with the experimental resxjlts.

LIST OF REFERENCES No. 1. 2.

3.

4.

5.

6.

7.

8.

9.

Author

Holme, Olaf A.M.

F l a x , A.H. and Lavyrence, H.R. Knight, M. and Looser, 0. Jones, R.T. Lawrence, H.R. Levacic, I. Scholz, N. Wieghardt, K. Weissinger, J. T i t l e , e t c . Measxjrements of t h e Pressxjre D i s t r i b u -t i o n on Rec-tangxjlar ^Vings of D i f f e r e n -t Aspect R a t i o . P . P . A . R e p o r t No. 37, "1950 The Aerodynamics of Low Aspect R a t i o

Wings and ?fing Body C a n b i n a t i o n s .

T h i r d Anglo-American Conference, 195'1 Pressxjre D i s t r i b u t i o n over a R e c t a n g u l a r Monoplane Vfing Model up t o 9 0 ° Angle of ' A t t a c k .

N.A.C.A. R e p o r t No. 288. 1929 P r o p e r t i e s of Low Aspect R a t i o P o i n t e d

¥/ings a t Speeds Below and Above t h e Speed of Soxjnd.

N.A.C.A. R e p o r t No. 835. 1946 The L i f t D i s t r i b u t i o n on Low Aspect

R a t i o Wingsat Subsonic Speeds.

I . A . S . P r e p r i n t No. 313. Janxiary, 1951 R o l l i n g Moment dxje t o S i d e s l i p , P a r t I I . The E f f e c t of Svreepback and Planform. . R.A.E, R e p o r t No. Aero. 2092,Nov, 1945 Kraft-xmd Druchverteilxongsmessen an

T r a g f l a c h e n k l e i n e r Streckxjng. Porschxjng axjf dem G e b e i t e des Ingeniexxrwessens, P a r t B, V o l . 1 6

No. 3, pp. 8 5 - 9 1 . 1949-50 Chordwise Load D i s t r i b u t i o n of a

Simple Rectangxilar Wing.

N.A.C.A. Tech.Memo N0.963. Oct. 1940 The L i f t D i s t r i b u t i o n of Sweptback Wings. N.A.C.A. Tech.Mem. No. 1120. 1947.

-13-TABLE I

Ordinates of the Aerofoil Profile

x / c

.450

.500

.550

.600

i

.650

.700

.750

.800

.850

.900

.950

1.000

+ y / c

.055

.051

.047

.042

.038

.032

[

'^^"^

.022

.016

.011

.006

.002

x / c 0

.013

.025

.050

.075

.100

.150

.200

.250

.300

.350

.400

± y/c

0

.018

.024

.033

.040

.045

.051

.057

.059

.060

.059

.058

1

- ^ ^ . . ^ ^ \ CHORD — - ' ^J^ORWARD SECTION.-SEMI ELLIPSE i-CENTRE

SECTION.-FAIRING BETWEEN FORWARD AND REAR SECTION

tREAR SECTION. : ;» FLAT SURFACES GIVING A TRAILING EDGE ANGLE

OF 12-2!'

SYMMETRICAL SECTION, MAXIMUM THICKNESS CHORD RATIO « . | 2 , AT .3c F R O M THE LEADING EDGE

FIG 2.A.

_{FIG 2 . a}

30 m O as O O m O m O •n > m :o O z > c H O

WING OF ASPECT RATIO 1-5

WING OF ASPECT RATIO -5

lO

>

'a»

N

01

lOO PORT STARBOARD UPPER SURFACE , + l-O PORT LOWER SURFACE C^P COVERALL)» 017» "Vp °" O*

ISOBARS A,R. a I • 5.

STARBOARD

COLLEGE OF AERONAUTICS REPORT No. 52.

FIG. 4.

PORT

»-roo

PORT UPPER SURFACE LOWER SURFACE. C N F C O V E R A L L ^ * 0 . 5 2 , ^ - O " STARBOARD STARBOARD

ISOBARS A . R . - 1-5

PORT °/o CHORD

ro

2 0 • 4 0 - 6 0 8 0 100 PORT UPPER SURFACE WIND DIRECTION STARBOARD STARBOARD LOWER SURFACE C N F C O V E R A L L AT Ap =0**") = o * I7, i p = 10=

ISOBARS — A.R. = 1-5.

COLLEGE OF AERONAUTICS REPORT No. 52.

FIG. 6.

»/o CHORD - 2 0 - 4 0 - 6 0 - 8 0 1-100 PORT - 3 0 WIND DIRECTION. - 2 . 5 - 2 0 STARBOARD UPPER SURFACE. «/o CHORD

r O

- 2 0 4 0 - 6 0 - 8 0 t o o PORT _STARBOARD LOWER SURFACE.

C N F (OVERALL AT X^ =sO**ï as O-52,%^! s : lO**

REPORT No. 52. o/o CHORD 'O WIND DIRECTION. - 0 - 9 - ! 0

r

- 2 0 - 4 0 •60 -BO I-IOO PORT STARBOARD •-IOO PORT UPPER SURFACE LOWER SURFACE C N F COVERALL y = 0 ° ) = 0 - 1 7 , Ap =20? + 0 - S STARBOARD

COLLEGE OF AERONAUTICS REPORT No. 52,

FIG. 8.

«/o CHORD r-O - 2 0 - 4 0 - 6 0 - 8 0 •-IOO PORT o/oCHORD •O - 2 0 • 4 0 >60 .BO L l O O PORT WIND DIRECTION -2-5 Ur^PER SURFACE. WIND DIRECTION. LOWER SURFACE. CNF COVERALL ,'.T Ip = 0 ) = 0 ' 5 2 , Tp ^^ 20° - l - O - 0 ' 5 STARBOARD STARBOARD

PORT STARB'O LOWER SURFACE.

C^P COVERALL)

2 0 • CHORD 4 0 6 0 -8 0 " 1 0 0

-= O-ll.

t

/ 7 /

1_IIL—'-•••iJf,i»_f '1JP-'. ""^ - ^ ^ . , ^ 0 2 2 ^, — O-l • + 0 - I / y \

= 0 ?

PORT UPPER SURFACE.

ISOBARS. — A.R.= 0-S.

STARB'D

z

o m O m O -n • m 10 O

z

> c Ó

31

o

2 0

-X

CHORD 4 O' 6 0

-

ao-l O O J

Cj^p(OVERALL) - O-41 - ^ ^ O*^

- 2 - 5 - 2 ' 0 - .5 id y y •Xi _tn O 30 H 7 O «Jl y O O r-r m O m Ü •n > m ao O 2 > C H O in PORT STARB'D UPPER SURFACE

ISOBARS

Ai?.= 0-5.

23

STARB'D LOWER SURFACE 2 0 H Its CHORD 4 0 ' 6 0 ' 8 0 -! O O J

Cj^P (OVERALL AT 1p = 0'*) - O-ll lj; = IÖ

PORT

ISOBARS

STARBO UPPER SURFACE

--AK

=

0-5.

m •v Ü :n - t T: o tn w O r-m O rn (J -n > rn 5D O

z

> c • H O y i •jn

WIND DIRECTION O T 2 0 -'o CHOPO 4 0 6 0 8 0 -I 0 O - "

C,^p (OVERALL AT - l | ; « 0 ) = 0 - 4 l .

WIND DIRECTION m -a O X} -i

z

o

n

O m Q m O -n > m O z > c O C/) l7> W PORT STARS D UPPER SURFACE

V

== 10 ISOBAR S — A.R.=O • S.

+ 0 - 8 S + Ü - 5 0 PORT STARB'D LOWER SURFACE CHORD 4 0 -6 C ' 8 0 ' I O O - '

Cf^jp

(OVERALL

AT T/J=0**) =*0-ll

t • - 0 - 7 - 0 - 6 Ü -*. -< z o O l to r-rn O m O -n > m :n O 7-> C H O CO

2)

Y **20° ISOBARS — A.R. == OS

WIND DIRECTION PORT STARB'D LOWER SURFACE. 2 0 H % CHORD 4 0 -6 0 ' 8 0 -lOOJ

Cj^P (OVERALL AT'^jZ-O®) = 0 - 4 I

WIND DIRECTION PORT 30 m O 3D -I Z O hi

a

o

t- T-m m O > m O z > c

o

11 STARSO UPPER SURFACE.

Ij/ — 2 0 ISOBARS—A.R.= 0-5.

- l O - 8

t

—6 — 4 - • 2

—o

y • O — ' 0-8c >'95c . • ^

y

•6 -8 l-O FRACTION OF SEMISPAN

ASPECT RATIO 1-5 C,gp ( O V E R A L L ) = O - 5 2 "ip = O*"

- l - O - • 8 ^ — •6 - • 4 — 2 • i • ^

1

T ^

i * * " " " " " ^ -O—o - X — o •8c •95c ^ '

o/

X

A

3 ^

cy^

^ / ^ ^ ^

t

•6 -8 1-0 FRACTION OF SEMISPAN

ASPECT RATIO O-S C ^ p ( O V E R A L L ) - 0 - 4 1 Tp

SPANWISE DISTRIBUTION OF PRESSURE CUDSE TO THE

TRAILING EDGE ON THE UPPER SURFACE.

lOO PORT 6 0 8 0 STARBOARD lOO » O «2 O O r X3 m -* O m ^ O ••^ > m Xi O z > c O

DISTANCE FROM <|__ '>/o OF WING SEMISPAN.

REPORT No. «

O

n

<

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COLLEGE OF AERONAUTICS REPORT No. 52

FIGS. iaA.&B.

FIG. I8.A.

2 0 t 2 0 DISTANCE FROM 4.— % OF WING SEMI-SPAN.

4 0 60 STARBOARD

FIG. I8.B.

6O 4 0 2 0 (E_ 20 4 0 6 0

PORT STARBOARD DISTANCE FROM ^ — °/o OF WING SEMI-SPAN.

POSITION OF LOCAL CENTRE OF PRESSURE.

A P . « l - 5 .

REPORT No. 52. 6 0

FIG. I9.A.

1

•< - "t é

1

^ ^

-T"

< = I00 =^20«

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1

STARE

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1

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=S5!

^ ^ ^ " • - ^

.-c

^ ^ C^P ( O V E R A L L AT 1|> = 0 ) = '11 ^

r

8 0 6 0 4 0 PORT 10 ^ DISTANCE FROM t ' 2 0 4 0 6 0 STARBOARD OF WING SEMISPAN. 8 0

FIG. I9.B.

>

1

*^, \

F=

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-0-8

.oV = ^ u y . p 0 -J a u. -4 a 0

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B _ 1 ^ • • " * j — ii "~—2 , — . — - ' ^

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/ / ^ ^ / / % i / C D A l 1 A T ^ l j — ^ ^ ^ . > l l ^ C l i p ^ OVER ALL AT u; x Q ) = ««H —

f

\ 6 0 4 0 PORT 2 0 ^ 2 0

DISTANCE FROM (^ — "Vo OF WING SEMISPAN.

4 0 6 0 STARBOARD.

8 0

POSITION OF LOCAL CENTRE OF PRESSURE

A.R. s= O S .

COLLEGE OF AERONAUTICS REPORT No. 52. FIGS. 20.A.B&C. FIG. 20. A. 0-2 G f •ift ' '<^ I Cjyjp"-52 • 0 — "©—C^p«.4l •Eh" "13"-C|^p='il J > ARs»|'5. AR=0-5. • » * -1 FIG. 20, B. - 0 - 2 C M ABOUT LEI - O ' ' ^ MA A AR a»(.5 C)^]pn<52 - ^ 5 ) . - . i 0 . . A R = O - 5 Cj^F = "*' FIG. 2 a C. - 0 . 0 6

H

- 0 0 4 - 0 - 0 2 -4^ C, AR=i|-5 •17 J 3 — 0 > - S F - M J - E h

1

lO 2 0 ANGLE OF SIDESLIP "^'» TO STARB'D

EFFECT O F SIDESLIP O N NORMAL FORCE, PITCHING AND ROLLING MOMENTS.

FIG. 21.

COMPARISON OF RESULTS FROM DIRECT MEASUREMENT AND

PRESSURE DISTRIBUTIONS.

FIG. 22.

A BALANCE MEASUREMENTS. V PRESSURE MEASUREMENTS. + HOLME ( R E F I.) O SCHOLZ (REF. 7.) WIEGHARDT (REF. 8.)

. _ pLAX AND LAWRENCE (REF 2) WEISSINGER, ( R E H 9.)

H^^\<.

1 2 3 4

NOMINAL ASPECT

COLLEGE OF AERONAUTICS

REPORT No. 52.

FIG. 23.

C ^ - C L CURVES OBTAINED BY DIRECT MEASUREMENT

FIG. 24.

A PRESENT EXPERIMENT

O SCHOLZ EXPERIMENT |5 «»/o THICK WING - THEORETICAL RESULT (REF 2 )