NACA-RM-L7B18 · Two-Dimensional Wind-Tunnel Investigation of Modified NACA 65(sub 112)-111 Airfoil with 35-Percent-Chord Slotted Flap to Determine Pitching-Moment Characteristics and Effects of Roughness

NASA (NTRS) · 1947

Open the PDFPublic domain · NASA (NTRS)Technical Reports

Overview

An investigation has been made in the Langley two-dimensional low-turbulence pressure tunnel to develop the optimum configuration of a 0.35-chord slotted flap on an NACA 65(sub 1120)-111 airfoil section modified by removing the trailing-edge cusp. The section pitching-moment characteristics and the…

Pages: 16

Key points

The investigation focused on the pitching-moment characteristics and lift characteristics of a modified NACA 65 (112)-111 airfoil with a 35-percent-chord slotted flap.
Tests were conducted at Reynolds numbers ranging from 2.4 x 10^6 to 25.0 x 10^6 to determine the effects of flap deflection and leading-edge roughness.
The section pitching-moment coefficient was approximately 0.10 higher than that of the NACA 65-210 airfoil with a 0.250-chord slotted flap.
At a flap deflection of 20 degrees, two values of the section lift coefficient were obtained due to inconsistent stalling of the flap.
The decrement in the maximum section lift coefficient caused by leading-edge roughness was similar for both the flap-deflected and flap-retracted conditions.

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Frequently asked questions

What was the primary focus of the investigation?

The investigation focused on determining the pitching-moment characteristics and lift characteristics of a modified NACA 65 (112)-111 airfoil with a 35-percent-chord slotted flap.

What Reynolds numbers were tested during the investigation?

Tests were conducted at Reynolds numbers ranging from 2.4 x 10^6 to 25.0 x 10^6.

How did the section pitching-moment coefficient compare to the NACA 65-210 airfoil?

The section pitching-moment coefficient was approximately 0.10 higher than that of the NACA 65-210 airfoil with a 0.250-chord slotted flap.

What effect did flap deflection have on lift coefficients?

At a flap deflection of 20 degrees, two values of the section lift coefficient were obtained due to inconsistent stalling of the flap.

What was the impact of leading-edge roughness on lift coefficients?

The decrement in the maximum section lift coefficient caused by leading-edge roughness was similar for both the flap-deflected and flap-retracted conditions.

Summary written by AI from the source document. Check the original before relying on it. See which parts of Sprinkle are written by AI.

Document

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,RESEARCH MEMORANDUM

NATIONAL ADVISORY COMMITTEE

I t I The L-esults of the first phase of t h i B investigation,which cwered the developaent of t h e optimum f l q ? confimwation at Reynolds number of 2.4 X lo6, have been i,elIo;.ted? in imference 2.

The second phase of the investigation cove:~aC the deve1opQen.t; of the Optimm flap configuration at high ReJmolGs nan3ers and t e s t a o f the airfoil with the fla5 retracted md tho condition with t e

optimum flcg configuration at Reynolds nUmbe2s r;~ t o 25.0 X 10 2

I The r e s u l t s of those tests have been repwteC- i n reference 3.

This paper presents the results of t e s t s miie t o iletemine the section pitching-moment c h ~ a c t e ~ i s t i c s , t h e li?t characteristics for a develoyed f l a p path, and the effects 02 leadiqyecige roughness on t h e s e c t i o n l f f t chasacte:-ietics. These @ore- mentioned t e s t s conclude& the investigation.

NACA RM No, Lml8 , section & a g coefficient 'd section 1 3 . 1 3 coefficient c l .

'section pitching-mcanent coefficient Cmcr4 R Reynolds number * .

X> Y horizontal and vertical ,positions, -emectively,,of center of the flag leadFng-eQe radius with respect t o u2ue:- l i p of s l o t in yercent C, l o s i t i v e forward of ana belaw s l o t u p , respctive-, (fig. I> The 2-foot chord model tested b t h i e i n n s t i g a t i o n was a modified NACA 65 (n2>-U1 airfoil s e c t i o a m t h a 0.35~ s l o t t e d The airfoil section had beenmodizied bF .-amoving the ' f l a p .

traling-edge cusp esd 829 therefore aimilcc: t c m NACA 65

( 1 1 2 ) -

a i r f o i l section(refezence 4). Ordinates POT the ?lain - f o i l I section and the slotted flap a m gi-ren i n %&lea I and 11, respectfvely.Figure I is a sketch of the aizcoil and flap and also shows the reference pointe d e f i n i q the f l a g position.

A flap path was develowd by deflecting the fir? through a circular path fn such a m a n n e r t h a t a t a flag CePlection or" 35O the center of the flap leading-eQe racXus rraa located.

1.98-prcentc behind and 3.21-percent c belmr t h es l o t . l i p .

This flap deflection and position had been f o m d t o be the optimum .

at high R e p o l d s ntrmbere (reference 3) %e location of the I pivot point about which the flag w w deflected is s h o w in figure 2. The model w a s constructea of almin-m U o y and cnmpletely spanned the 3-foot-wide tunnel test section.

Test8 were made i n the Larigley two-dimensional low-turbulence ~ ~ S E I U ~ - Q tunnel to determine t h e m a l e e f f e c t s on the section pitchfqymcanent chmacterietics of the ai;Toil Eection with the f l a p retracted and slot sealed f o r Reynol6-s nwnbezs ranging frcan

3 . o x 10 to 9.0 X 106. L i f t meaawements 7m-e made at a

Beynolda nmber of 9.0 >: 10 6 t o determine the section-lift chur-actez.ist9cs of the m o d a l f o r flap deflections u2 t o 3 5 ' .

' The effects of standard roughn.ee8 (reference 5 ) on.the l i f t chazacteristics of the 0ptirm.m flap conflaxration and on the lift Ch€1~aCteri&iC8 of the model w5'Lk the f l a : ) i d X W t e d and s l o t sealed. were detei-mined f o r R Reynolds nmbei- 03 6 . 0 X lo6.

The effects of standard roughneaa on the section d;'q character- iatfcs of the model wikh the f l a p retracted s l ? & el& sealed were also dete;Plined for a Regnolde number of 6.0 :: 106. The t e s t methods and the m e t h o b ueed i n c o x r e c t i r g t h e t e s t data t o free-alr conditions =*e discussed i n reference 5. The magnitude of the corrections used In correcting the te3t data to free-air conditions waa of the order of a few serceak., The maximum free-atmam Mach numba:L- attained during of t h e . tests did not exceed 0.17.

PitchingMcmmnt Cha-cacteristica I The eection pitchfng-mmentcharacteilisticB I Q ~ the a i r f o i l section with the f l a p deflected a r e presented i n ffgure It.

Includedwith the section pitching-mcment data, lregented i n .

figure 4, a r e the eectiion l i f t chmacteL-istlce o b t a h e d from reference 3- For the flap aeflected ConZiCuration %lie elope of the pitching-mament cume becomes posi,tive8 at t x q l e s df attack f ' r a n about 2O t o s l i g h t l y above the stall, but BLL>therincreases NACA BM No. ~ 7 ~ 1 8 i n the section of attack cause a reversel of the slope and theslope of the pitching-moment curve becomes necative. Tlhe , value of the section pitchinglncanent coefficient is approximately 0 . 1 higher than that obtained for the N A C R . 65-210 a i r f o i l section with a 0.250~ slotted. flag (slotted flq] 1, 2efer.ence I ) and approximateb- 0.05 lower than that obtained 20s the mACA 65-210 airfoil section with a 0- 3l2c double 81okteiL fla:, (reference I) .

The maximum section lif’t coefficient vas, however, qproxfmately- 0 . 1 higher than that obtained f o r the NAGA 65-210 aii.foil with the slotted f h p and apTroximateu 0.1 lmrel- t h m t h a t o b t a h e d fa- the XACA 65-210 a i r f o i l with the aouble slotted Tlap (reference 1) .

Lfft Characteristic6 for Intermediate F 1 q Deflections The section l i f t characteristics of the model with the flag b f l e c t e d up to 350 f o r a Reynolds number o~ 9 . o :: 106 a r e px-esented i n figure 5. Measurements of ,the l i f t chax-acterfstics for flap deflections -eater than 350 weLw not m a b because, as discussed in reference 3, FncreaBingl the f l q deZloction beyond gave no increase in the m a x j z ~ m flection E P G coefficient at a Reynolds number of 9.0 X lo6, A t a fla., cleflection of 2 0 ° and a t section angles of attack higher than abouk -4O, two values of the section l i f t coefficient Irere obtained et ecch m e of attack, f

&though the values of the rmxhmsm section lirt ooefficients were

nearly the same. R e r u n s of these teste shad!. that the condition giving the lower l i f t coefficients was the moi’e stable of the two.

T u f t studies at a flap deflection of 20° indicated that the irregular behavior of the lift coefficients wa6 amociated wLth p a r t i a l stalling of the ffa2 cawed by the :-elztively poor slot shape for this flag deflection.

Effects of Le&%-Edge Eo-eas The section lift md drag chmactei*ist;icS at a R e y n o l b number of 6.0 X 10 fcc the airfoil section with standard rou&ness are pzasented i n figui-e 6. Lncluded with thedata presented i n figure 6 are the section chazactez-istica f o r the smooth condftion which wa-8 Obt&hl€d fram i-el”o;-ence 2. The decrement in the mRxirm?m section l i f t coefIicie&, caused by leading-edge roughness f o r the flag-defleckeL conation is approximately the same as that f o r the fla-.,-zet::mAml condition.

The decrement i n maxirmrm. section l i f t coeEicfen”; i s approxinately the same as that obtained f o r the NACA 65-210 a i ~ ~ f o i l section with the 0.2% slotted flag designated as elotted f l a p I i n reference 1. The minlmwn section drag coefPic2en-h f o r the smooth condition and the condition with leaCLi&-erlge :o@mees a r e the same as those estimated f x g n consideration oi: Sata f o r ~imllar airfoil sections p e s o n t e d in T i g v ” e 8 Sl&&, 511I-7, and S148 of reference 5.

The results of tests of a modified MACk & j -111 a i r f o i l

section with a 0.35-shord s l o t t e d flaz in⁣a.i;e L+ t , le following

conclusions: 1 . The section pitching-moment coefficient va8. approximately 0.10 Mgher . than that obtainea f w the MAC.!! 5”j-210 a i r f o i l with a L).250-choi-d slotked flap ,and a22roxinmtiely 0.05 1mmr than that obtained f o r the NACA 65-210 .airfoil sectLon v i t h a ’ 0.312-chord doukle alotted. flap.

2 . At a flap deflection of -20° and sect-lon angles of attack higher than about -kO, two vPlues oi? the section lift I coefficient at each a n g l e of attack were obtainsi? because of inconsistent stalling of the flap althoLgh tho neximxa section I lift coefficient l’m8,ined near* the same.

b ?. The decrement in the ‘maximum aec’;fon 1iP-t;. coefficient obt&ed by qqlying standard roughnees to the. ai;*ioil with the flap deflected 3 5 O w a ~ ag$roxlma’cekf the mane qs t h a t obtained f o r t h e airfoil wfth the T l q retyacted.

I bw NACA RM Wo. ~ 7 ~ 1 8 7 I I I I I NACA RM No. L7B18 S t r t l o n s and ordinatesgiven -011 o h o g ‘ - h peroent I I ORDIHA’ES BOK 0.35-CHOI(D FLAP I : Upper surfaeo fairs l n t o p l a i n a l r f o l l a e o t l o n