rEr H,

TECHNICAL NOT~S

N~TIONAL ADVISORY COMMITTEE FOR AERONAUTICS

1:10 . 763

WIND-TUNNEL INVESTIGATION OF TWO AIRFOILS WITH

25 - PERCENT-CEORD GWINN AND PLAIN FLAPS

By Milton B. Aces , Jr . Lan~ley Memorial Aeronautical Laboratory

\II 9.sh ingt on Ma;y- 19,10

https://ntrs.nasa.gov/search.jsp?R=19930081584 2018-05-25T13:03:57+00:00Z

· ,

NATIONAL ADVISORY COMMITTEE FOR AERONAU TIC S

TECHN I CAL NOTE NO . 763

WIND - TUNNEL INVESTIGATION OF TWO AIRFO I LS WITH

25 - PERCEN T-C HORD GWI NN AND PLAIN FLAP S

By Milton B. Am e s , Jr.

SUMMARY

Aerodynamic force te s ts of a n N. A . C . A . 23018 airfoil with a Gwinn flap hav i ng a chord 25 percent of the ov e r ­al l cho rd and of an N. A . C . A . 2 301 5 airfoil wit h a p l a in flap hav in g a 25 - per c ent chord were conducted in the N .A. C. A . 7 - by 10- foot wind tun ne l to determine the r e l a ­tive merits of th e Gwinn and the p l a i n fl aps .

The t ests indicat ed t hat , based on sp e~d - r ange ra­tios, the p l ain flap wa~ mo re effective t han the Gwinn f 1 a p • At sma 1 1 f 1 a p de fIe c t ion s , th e pIa i n -f 1 a p had the low er d r ag coefficients a t lift - co e ffici ent values l ess than 0.70 . For Lift coefficients greate r than 0 . 7 0 , how ­ever , the Gwinn flap at a ll downward f l ap deflections had the lower d r ag coefficients .

I N TRODUC 11 I ON

Improvement in a irp l ane pe rformance has depended somewhat on t he deve l opment and the use of h i gh - lift de ­vices . As an aid to ces i gne rs , the N. A . C. A. ,has conducted many exper i monta l investi g at ion s of vario us types of fl ap and has r epo rt ed the effects of these ~ifferent flaps on high or low drag at -h i gh lift ( g li de - la t h control) , low d r ag in t he cruisi ng - speed condition , and h i gh lift -and l ow drag at the t ake - off .

The p r esen t i nves ti ga tion , con d uct ed i n t he N. A . C. A . 7- by 10-foo t wind tunnel , wa s of a h i gh- lift device identified as the Gwinn flap and i nc luded fo r comparison the tests of a p l a in flap . The Gwinn fl ap i s essentially a flat p l ate mo u n t ed a t a po i nt ve r y nea r t he trai lin g edge of the wing . In it s neutra l pos iti on , the flap ex­tends past the trailing ed~e of the wing , t hereby re s ulti ng

2 N. A . C . A. Technicnl No t e No . 76 3

i n an i nc r eased ove r - a ll ch ord and wi ng a r ea . I n some r espe ct s t ho Gw i nn f l op i s s i mil a r to th e p l a i n fl ap . Both f l aps de flect upward or d own wa r d so t hat e ithe r of them may ~o u sed as a h i g h - lift dev ic e and a l so as an a il o r on .

The p ri nc i pal purposo o f t h i s i nvos ti gat i on was t o de t o r mi no tho r o l a ti vo m~ ri ts of t ho Gwi nn an d t ho ordi ­na r y plain f l aps as hi gh - li f t dov ic es . P r evious tes t s (r e f e r ence 1) hav e b een c ond u cted on a irfoils wi th p l a i n f l ops ; t he mode l s u sed , ho{ove r, we r e no t c ompa r ab l e wi th t he G1.>l i nn f l ap mod el used i n t h i s i nvGst i go t i on . The flode l s o f t ho a ir fo il s wit ~ the Gw i nn and the p l a i n f l aps u sed i n t hese ~ e , s t s had t he saoe ove r - a ll c h o r d , span , aspe ct r a ti o , and app r ox i mate Dax i ~uD t h ickness .

APPARA TUS AND TEST S

Mode l s

Tw o a irf oi l mode l s we r e t es t e d : An 8 - l n c h - c ho r d N . A . C. A. 230 1 8 a i rfoi l se c t ion wi t h a Gw i nn fl ap hav i ng a cho r d 2 5 pe r cen t of the ove r - al l ch ord ( fi g . 1) and a 10 - i nch- c h or d N. A . C. A . 230 1 5 ai r fo il se ctio n wi th a p l a in f l ap havi n g a 25 - p e r cent cho r d ( fig . 2 ) . Ea c h mode l has a span of 60 i nches and an ove r - all chord o f 1 0 i n c hes . The maximum thickness of t h e mode l with t he Gw i nn fl ap , b ased on ove r - a l l cho r d , i s 1 4 . 4 pe r cen t of t he c ho r d and , of the mode l wi t h t he p l a i n f l ap , 1 5 pe r cen t . Bo t h a ir ­f o il s and t he pla i n f l ap a r e made of l am i na t ed b ee c h ; t he Gwinn fl ap i s made of a l uD i num . In o r de r to mo u n t t he Gw i nn fl ap , 5 pe r cent of the ove r - a ll cho r d was removed fr om the tr ai ~ in g edge of the H . A . C. A . 230 1 8 ai r fo i l se c ­t i on . The flaps are a r ran ge d to mo~e u p or down about t he ir r espect i ve h i n g e axes o r to l ock r i g id l y i n a g iven pos iti on . F l op def l ect i ons we r e me~sur ed wit h r espe ct t o ,t he a i rfoi l chord line , and a l l gaps between the ai r fo il s and t he f l aps we r e sealed wit h p l as ticine t o p r even t a ir l eakag e .

Wi nd Tu nne l and Bal~nce

The t es t s we r e nade i n t he N. A .C. A. 7 - by l a - f oo t t unne l , wh i ch has a c l osed th r oa t and r e t urn passage . The t unne l and the , r e gul a r 6 - c ompcnent bal a n ce are desc r ib od i n r e f e r en c es 2 and 3 .

, N.A.C.A. Technic n l Note No. 763 3

Tes ts

Test condition s . - ·Th e dynam ic pressure was maintained constant throughout the tests at 16.37 pounds per s~uare foot, correspondin g to an air speed of about 80 miles pe r hour at standard sea- level conditions . The average te s t Reynolds Number was 609,000 , based on the air speed an d t he 10-in ch a irfoil chord .

Test procedure. - Tare t es ts we re conducted to deter­mine the effects of the model - supportin g s trut on the li f t, the drag, and th e pitchi ng momen t s of the two airfoils and fl<1ps.

The ma i n portion of t he inves ti ga tion c onsisted i n de t e r minations of lift , drag , a nd pitching moments for flap deflections of _ 10° , _5° , _ 2°, 0 0 , 2 G , 5°, 10°, 1 5° , 30 0 , 45 0 , 60°, and 79° throughout a n a ng lo- of - attack r a n g e from -12° to beyond the stall for each of th o ai rf o ils .

RESULTS

Co ef fi c ients

Th e t es t result s are g iven in t he f o r m of standard absolute coefficients of lift , d rag , and pitching moment.

Om ( ) a . c. ° :::

where

CL lift

:::

~S

CD ::: drag

qS

pitching mom en t ab out aer odynacic c en t o r of airf o il wi th flap n eut r a l

~c S

L li f t :::

D d r ag

4 N. A . C . 1 . Te chnic a l Note No . 763

S win g area .

c ove r - all ch ord of win g a nd fl a p .

q dynam ic, pressure .

and a ang l e of a t tack .

6f fl ap def l ection ( d ownwar d def l e cti on i s positive ).

All coefficient s we r e ot tain ed d ir e c t ly fr o ~ th e bal­ance and ~ofer t o t he wind ( or tunno l) axes .

~ o rr c ct cd test r es ult s .- The da t a we r e correc t ed fo r tunne l ef f e cts to a spec t r atio 6 i n fre e a ir . T~o s t andard j e t - b ou nda ry corr e cti ons we r e ap p lie d . (S ec rofere n ce 4 .) I n addi tio n , correcti ons were applied for t he e ffects of t h e s u pp orting strut on t he ae r odynam ic c o e ffici en ts of t he models as indicated b y t he t a r e t es t s .

Th e cor r e ct ed t es t re sults are presen t ed in figur es 3 t o 6 as p l o t s of lift , d r ag , and p i t c h i n g - momen t c oeff i­c i en t s an d lift-drag r a ti o aga i nst angle of a ttack .

DISCUSSION

The lift an d the d r ag c oeff icien ts for t he ai rf o il with the Gw i nn fl ap are p l o tt ed aga i ns t ang l e of a tt ack for t he diffe r ent f l ap defl e ctions i n fi gure 3 . The p itchin g -moment c oef fici ent s and the li f t - d r ag r a tio s fo r th e same c ond iti ons a re p l ott ed i n fi gure 4 . Figure 3 i rid i ca t es tha t CL varied r egula rl y with f l ap def l ection

except w~en t he fl ap was deflected 30 0 • The irr egul a r curve f o r the 30 0 flap deflection ma J be a c ha racteri s tic of t he ai r fo il - and - flap co mbinat i on , or i t may be a ttrib­uted t o scale ef fect . The caxiDum va l ue of CL occur red a t 6 f = 60 0 ( fi g . 3(b) ), an d th e max imum va~ue of LID was 1 8 . 6 a t 6f = 0 0 ( f i g . 4 ( a )). Th e v a ri a ti on o f GD ( , ) wit h fl ap def l e ction i'lnS un i for r.:l ; t he upwar d

n . c . 0

defle cti ons o f the flap t ended to g i ve st a llin g Dom en ts , and the downward deflect i ons gave d i v in g DO Den t s .

N.A. C . A. ,Techn i cul Note No. 76 3 5

P l a i n F l ap

The li ft and the d r ag c oeff i c i en t s f or t he a i rf oil wi th the p l n i n f l ap a r e p lott e d aga i ns t angl e of a t ta c k for the d i f f e r en t f l a p de fl ect io ns i n f igur e 5 . The p i tching- momen t co ef fi c i e nt s a nd t he li f t - d r ag r at i os are p l otted in f i gur e 6 . As i n the c ase of the Gyr,i nn flap , the CL c urve f o r t he p l a i n f l a p de fl ec t ed 30° i s i r r eg­ular, which may be at tri buted t o t he fl ow c ha r a ct e r istics and the s c a l e of t he t es t s . The max i muTIi va l ue of 01 - ' occurr ed a t 5 f = 60 0 ( f i g . 5(b )) , and the max i mu m va l ue of LID was 19 . 4 a t Of = 0 0 (f ig . 6 ( a) .

Compa ri son of Gwi nn and P l a i n F l aps

Enve l ope po l ar c urves of CD p lo tted a g ainst 0L for the airfo il s-w i th the Gw i nn and t he p l a i n f l aps a t the d iffe r en t v a l ues of o~ a r e g i ven i n f i gure 7 . These curves i nd i cate t ha t, fo r ~a l ues of C1 fr om - 0 .1 5 t o 0 . 66 (wh i ch c ov e r s the h i gh- spee d and the c ru i s i ng- speed ranges ) the pla i n f l ap had the l ower d r ag . The mi n i mum value of CD fo r t he a ir fo il wi th t he p l a i n f l ap , 0 . 0095 , occurred a t Of = 0 0 and , for t he ci i rfo il wi t h the Gw i nn f l ap , the mi n i mum va l ue was 0 . 0 1 07 a t Of = -2° . (See also figs . 3 ( a ) and 5 ( a ) . ) On l y a s li gh t d i ffe r ence exis t ed be­tween tho drag characteristics of the Gw i nn flap at Of = - 2 0 and ' Of = 0° . The maximum li f t coeffic i en t 01 _ max of the Gw i nn fl ap a t Of = 60 0 wns 2 . 03 and , of the plain flap at the sa~e f l ap def l ect i op , was 2 . 00 . ( Sec fig . 8.)

Th e effec t of flap deflec ti on on CD at diffe r en t values of 01 i s shown i n f i gu r e 9 . The pla i n f l ap has lower drag coefficient s for va l ues of 01 t h r ough 0 . 70 i n the flap - deflect i on ran g e o f - 100 to 50 . At C1 = 1 . 00 , the Gwinn flap had t he l ower drag coeff i cient .

Fur t her conparisons a r e g i ven i n t he fol l owin g ' table for tlO cond iti ons of f l aps ne u tr~l and deflected 60 0 • The comparison includes £he ' i ncrecont of naximuD lift coeff i­cient due to f l ap defle cti on 6C L ' , the speeQ-range

D 8 X

ratio and ' the glid~ path , indicated by

LID

6 ,N.A, O.A . Tec hn ic a l Not~ No . 763

-------T--------------------,-----------------------------------

ToL 11/D a t 1 ° I max I

F ap 1 ,--- ' OT

=1 max I °D . , .wmax I O l n

I - I

I Gw 1. n n 1. 1 9 I 112 7 . 0

Pl a i n 1. 09 11 5 11. 9

( a )O D a t 01 = 0 . 20 . o in

2 .'03 0 .84 190

2 . 00 . 9 1 I 21 0

CD . [1 1n

( a )

1/D a t

°1 171 ax

I 145

1 ,_66---1.1

___ _

5 .1

4- . 8

From t his tabl e it is observed t hat , a lt hough t he Gwi nn fl ap both neutral or def l ected 60 0 has s li gh tly h i ghe r values of 01 ' t he incr emon t of CL c aused

Dax max b y de flecti ng' t he f l ap i s g r eater fo r t he p l a i n fla p . The conpar i son also i nd i ca t es t ha t the p l a in flap has the hi gher s.pe8d - ran ge ratio w,he t he r t he fl ap i s n eutr a l or de -flect ed 60 0 and whether CD ' i s taken a t 8 1' = 0 0 or

O l n

at C1 = 0 . 20 . A c ompa ri son of the val ues of LID a t '

C 1 max

shows t hat a s t eepe r g liding angl o c oul d b e ob t a i ned

wi-th th e pla in fl ap deflected a t t he snDe de fl ec tion.

60 0 t han with t he Gwinn fl ap

CONCLUDING REMARKS

The r esult s of t h is i nvesti g nt i on of an N. A. O. A. 230 18 airfo il with a Gw i nn flap hav i n g a c ho rd 25 pe rce n t of t he ove r- a ll chord and of an B . A . C . A . 230 1 5 a irfoil wit h a p l a in flap hav i ng a 25 - pe rc en t chord i nd ic a t ed that , fro m a con ­sideration of speed - r ange r a tio, the p lain flap was more ef fecti ve than the Gwinn fl ap .

A cODpnri son of t he two t ypes of fl ap a t soall f l ap def l ec ti ons showed t ha t, for va l ues of the lift coefficien t of 0 . 70 or l ess , t he p l a i n f l ap had t he l owe r d r a~ c oeff i-

N. A. C.A. Techn i cal Note No . 763

cient. At nIl downward fl ap de fl e cti ons , however , the Gwinn flap had t he l ower d r ag coeffici ent at lift coef ­ficients g r eate r t han 0 . 70 .

Lan g ley MeDor i a l Aeronautical Lab ora t o r y . Na ti ona l Advisory Comm itt ee for Aeronautics,

Lan g l ey F i el~t Va ., Ap ril 4 , 1 940 .

REFEREN"CES

1. Wenz i nger , Carl J.: Wind - Tunn el Inv es t igation of Ordinary an d Sp lit F l aps on Airfo il s of Diffe r en t Profile . T. R . No . 554 , N,.A . C. A ., 1 936 .

2 . Harr i s , Thooas A.: The 7 by 1 0 Foot Win d Tunnel of the Nati onal Advisory Committee for Aoronautics . T . R . No . 412 , N. A.C . A., 1 931.

7

3. We n zing er , Ca r l J., and Harris, Th omas A.: Wind - Tun ­ne l Investi g ation of an N. A.C . A . 23012 Airfoil wit h Various Arrange ments of Sl otted Flaps . T. R . No . 664 , ILA . C. A., 1 939 .

4 • 1~ e n z in g e r , Car 1 J ., an d. Am 0 s , Mil ton J3 ., Jr .: Wi n (1 -Tunn ol In v estigation of Rectan gular and Tape r ed N. A . C. A . 23 0 12 Win g s with Plain Ailerons and Full ­Sp a n Sp lit Flap s . T. N. No . 661 , N. A . C.A ., 1 938 .

N. A.C . A. Tec hnical Note No. 763 F1gs. 1 ,2

~------------------.BOc --~--------------~

- .0057c

1<---------/0 in. = /Oo.OOC ---------->1

Figure 1 . - Gwinn flap on the rectangular N.A . C.A . 23018 airfoil .

. 0146 c -11 ~ ,cf4 - ---=*-~,' -------- ----------."-(a.c)o

--

k--------IO in. = IOO.OOc ---------->1

Figure 2.- Plain f lap on the rectangular N.A.C . A. 23015 airfoil.

N. A.C. A. Techn i cal Note No . 763 Fi g . 3

I.

J I, ,I \ Of-tlf , deg Of ' deg I .L.! ~' , -10------ 0 /5- - - --- ---6

I-- -5-- - --<1 30----0 l¥ l/ - 2 ------ - - ---v 4S-----v 81-- 0 0 60 0 ~ L t I- 2--10. 79 -- -------iO>.

5--- -[>. lL / V ,,- .., 61- /0- -- - - --0

lL 1/ I-a to I pI/, ~ / A a:" ".." U.

4 0--0 /1' k~ .P"t] :a..o L , ,

!

I /' ~ .... f<:'~ -l>... /Y / lP ,~

/ 2

J/ II/: ~ [;".., l~l'> ld" fa' .I .L ~// V, '" 1/ fdA

~ "/ iP' ,/ , "

, 0 V j ~/ i. F" 10--<>- I/;~ / lL' 10"0" r>--c , ,

/.

IV :l/ ' / V IV ~ , ' 8

,,'; V, I , ~I ~ / ,

~/ / VI ~! ~ ~ V If , 6

" ,-/ If! / / V ,1 ~ , ,

/ )1/ V If/ If' , , 4 8

~: / VI,' ,

,/ 1/ If I } " , ,/ ;' rl 1// ,I / I

, " 2

11,/ rio II' P' / l? / ,~ ,

/ v) ,;5 V' / I .;{ lL I

/ // ~, / / 1Y' 1/ p

/ V rl ' ~' <I' It V // 1// :' 1/ / ~,

..L /' Is II; ,/ J ~/ l? II }'

1/ ) (/' J~ lP' ~ IP

o

-.2

- .4

44

40

36

32

V) " 1/ 1/ ¥ V v LL " /' / ,/ Ii II p

( I

-.5

,Q' • ,~ ,/ v y/ } I." ~ 0-

V ,1 Ij:~ 1I'.f.. v ,/ ,P ," -.8

~rir tl ~' ,(" 1,/ / I

0 Ii IL ifz .L.~ L 1/ ;/ p

? ! //.. r/ i «/ .r! ,A

-I.

/6

Ii Ifl , L fr/ p/ If' !

V /''11/ ,,/ ~ ,ff

/

/2

/ / W i/ lL' "..,. p/ ~

A" / /~ ij ,/, 4/ /~' If , 08

c " I'-~I\ p" ~ ~ .,'. 10"/0 y :.-n- ~........-/

/

~~ --'c ~ ---" 0'" .--::: ~ ff:;', /1>-/ - ~-

04

I-i ". ~. - -- (al (b)

4 8 /2 /6 -8 -4 o 4 8 12 /6 20 -12 8 4 a o Angle of attock, d I deg

(a) Small flap def l ections . (b) Large flap deflecti ons.

Figure 3 .- Lift and drag coefficients of the rectangular N. A. C. A. 23018 a i rfoi l with 0 . 25c Gw i nn flap.

N. A. C. A. Technical Note No . 763 Fig . 4

.5.-.-.-.--.-,-,-.-,-I-'-'-'-'--'-'-'-'--''-'-'--'-'-'-'--'-'-'-'--'-'-'-'--~ ~~~-+~~-+~-+~~-+~~~~~~~~4-~~4-~~

.41--+-+--+--+--+-+--+--+ 6f ' de) -f--ll--t-+--t-+t--t--t-+-+--+-t--l---t 6f ~ dJ9 -t-+_+__+-t--l -10------ 0 15---------A

I--+-+--+--+--+-+--+--I--t -5--. -----.:I .30- - --{J -+--I~ - 2------- -- --v 45- - -- -v -+--I~ .31--+-+--+--+--+-+--+--I--t 0 0 60 0

~ 1--+-+--+--+--+--+-1---1-- 2 ----A 79-- - -- ----i>. -+--1--1 ~ 5 - - - -1>-

cJ . 21--~_+_-I--+-_+__+__+--I--t 10 - -- --- - - -0 +-+-+-++-+-+-+-+--+--jf-t-+-+-++-+--+-i--H

I

(a ) Sma l l fl ap def l ections. (b) Large flap defl ections.

Figure 4.- Pitch i ng-moment coeff i ci ents and l i ft-d rag rat ios of the r ectangular N. A. C. A. 23018 airfoil wi th 0 . 25c Gwinn fl ap .

N.A.C.A. Technical Note No. 763 Fig. 5

20 J I J I, ~~ 1-0,., deg I I. o,.,deql I I ,

f/ ~ -10------0 15--- ------/1 - -5-----.l 30----0 "

1.8 -2---- -------v 45-----'1' L /

- 0 0 60 0 I/" II' IP. - 2----6 79-- -----t. /

5----t. _1~ ~ P ~~ - /0-------0

/.;5 / ? ~i<-o ,r / 1/ &f' A. ... A.

1.6

1.4

;:/ ''0 -D-u ~/ '<> I>- rt-, / / d

V /' ~ "..."..~ ,til l)f Ii , ,>'

1.2

p-' l/ P/ .,7tJ "'"v-i'>-': II', 9' 1/ II , , ,

1/ .I I/; , l/' ,....,- "\, / V I / , rf ,

1.0

/ I, V: / fY o-<T 10" ">-0-~ ,tI / )f fl ,

/ If ~, I /, ff V ,8 " .8

,~ / VI I ~. II // / // ,(' , ,

V / /; P'; / / 1/ , , , p.--

" / :/, I , .( II / i'"

/ V ~// I / / , p

48

/ I/i ,./ ,/ /' / , "

V I: V, i :; , ,

,.J '" , /

44

(I' V ,

;/ / '

V 1/ ,v' /' , , ~/ j;17 I p

I

o 40

// I V' ,I / / .?

~ , t<

V V, I I' 1/ V ) -,2 36

~' / ,IS' Y I y'

V 'i , l/ / /' , /' / /"

-,4 .32

1/; ,'lV / "'"

V v'" r V,~ / /- ,/ 1/ ,if

f '" -.6

~ / /~ ,l- V / l/ lP' d

,I-,

LP lL 1/ )1/ ,

(!

-.8

~ ,Il V ,/ -T' il-

I' l1'f l-{ ./ ,

? -/0

/J / IJ /* / }'

P -.ft '/1

[Iv / 16 d , 16

1,/ i / 4 ' ,<r' :>-

/,0 , ¥

/ /' "'j I'~I / rf" ;/ , /2

" rf/ / ~,,;; ~,6 'I f:>:::. V lP/

~l\ \: p/ Y ~ fP/ /'f{ >---' ~-;r" /~ p' -

08

, cr/ 1'<" ~ ",,', ~_if ~~ i'-.... ..- j"/

~ ~ ;:-:--. > >..-, p--' ~ "'....--~ r:' ;'-', ;---- - - -- - --P- -04

- -- (a) (b)

-/2 -8 -4 0 4 8 /2 16 20 /2 -8 -4 0 4 8 12 16 o

Angle of attack, d • deq

(a) Small flap deflections . (b ) Large flap deflections.

Figure 5.- Lift and drag coefficients of the rectangular N.A.C.A . 23015 airfoil with 0.25c plain flap.

N.A.C.A. Technical Note No . 763 Fig. 6

.5 I I I I

1-6f ,deq Of. deq

.4 -/0------ 0 /5 -- -- -- - --0 r--S _ ____ il 30--~ -2-----------"<1 45-----"<1

I-- 0 0 60 0 .3 r- 2----0 79-----__

5----~ t I-- IO -------~D l!. ,

.2

- . / I - < ---< ~ ~-+..o- -¢-:>- - po-10-- 10-0.

~I~ .- - :>-- -- d--~ !->-~, v

~ - t---' -~ -- "'- -- It>-.,. 1<>-0..

r -- """" ~. ~ '." Iz>....n "-,, ~I<> - - -F>- j<o-- """f'>-t. ~...,

10--_n

, -.I ;>--1 ... >--p-'-~- 1->-- .<>-- /.0.-- 1<>-- fA- 0-."

--\<>-- ..... r.>

h -- -- --- -)- . I"" -.2 t- ~ :::.-' r;, ~ Jt....o. _ k.- '1<" -- --

-.3

20 F7 ~ t,::,,,,

16 /i! : >I ~ );)('< ~o,

/Jf. 7: P-.. 'f. ~ ~' ~ t<\, bo- "-. I.( .r i' '-~, , ~"\'Q --, it>.. I --12 / f / t '1"- "'" ~~ l\ - "-: , - 4._

/ I~ : I 1 ! ; Llrb.~ ~~ , -11.."",\

8 ,

II I , J b.~ ~. , p- - cr P- fa- "0_ fo- -<:>;: 1> ,

I

---- '" ; , : ro "&~ ,'/ ....- \<>-- - -fi'= .... ~ ~ ~ -....... ), , 4

/ I I ~ ~ tr- fir" ~ro , I ~ : I I I i , :

/ 0 ~ I : I I

f

f 1/ t f Ii ,

f , , -4

/ III , II 1 , 1/ I I , T ,

I~ / I 1/ , 1/ ,

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(al Small flap deflections . (b) Large flap deflections.

Figure 6. - Pitching-moment coefficients and lift-drag ratios of the rectangular N. A.C . A. 23015 airfoil with 0. 25c plain flap.

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Figur e 8 .- Effect of f l ap def l ecti on on maxi mum lif t coef ­fi cient. The 0. 25c Gwinn flap on t he N. A. C. A. 23018 airf oil ann t he 0 . 25c plain fl ap on the N. A. C. A. 23015 airfoil.

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Figur e 9 . - Effect of flap 1eflecti on on drag coeff ic i ents at di f ferent lift coeffi cient s . The 0 . 25c Gwi nn flap on the N. A. C.A. 23018 a irfoil and the 0. 25c pla in f l ap on t he N. A. C. A. 23015 ai r fo i i.

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