Wind-tunnel investigation of NACA 66(215)-216, 66,1-212, and 65₁-212 airfoils with 0.20-airfoil-chord split flaps

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Material Information

Title:
Wind-tunnel investigation of NACA 66(215)-216, 66,1-212, and 65₁-212 airfoils with 0.20-airfoil-chord split flaps
Alternate Title:
NACA wartime reports
Physical Description:
7, 6 p. : ill. ; 28 cm.
Language:
English
Creator:
Fullmer, Felicien F
Langley Aeronautical Laboratory
United States -- National Advisory Committee for Aeronautics
Publisher:
Langley Memorial Aeronautical Laboratory
Place of Publication:
Langley Field, VA
Publication Date:

Subjects

Subjects / Keywords:
Flaps (Airplanes)   ( lcsh )
Aerofoils   ( lcsh )
Aerodynamics -- Research   ( lcsh )
Genre:
federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Summary:
Summary: An investigation was carried out in the NACA two-dimensional low-turbulence pressure tunnel of the NACA 66(215)-216, 66,1-212, and 65₁-212 airfoil sections equipped with split flaps having chords 20 percent of the airfoil chord. The purpose was to determine the maximum-lift characteristics of these low-drag airfoil sections with split flaps. All the present tests were made at a Reynolds number of approximately 6 x 10⁶ and a Mach number of about 0.15.
Bibliography:
Includes bibliographic references (p. 4).
Statement of Responsibility:
by Felicien F. Fullmer, Jr.
General Note:
"Report no. L-140."
General Note:
"Originally issued July 1944 as Confidential Bulletin L4G10."
General Note:
"Report date July 1944."
General Note:
"NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were previously held under a security status but are now unclassified. Some of these reports were not technically edited. All have been reproduced without change in order to expedite general distribution."

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003622093
oclc - 71380743
System ID:
AA00009397:00001


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Full Text
lC A L-o I4
CB No. LG1lO
F


NATIONAL ADVISORY COMMWE FOR AERONAUTICS


'll RTIll I P RT


WARTI'IKE REPORT


ORIGINALLY ISSUED
July 1944 as
Confidential Bulletin L4G1O

WIND-TUMNEL INVESTIGATION OF NACA 66(215)-216,
66,1-212, AND 651-212 AIRFOILS WITH
0.20-AIRFOIL-CHORD SPLIT FLAPS
By Felicien F. Fulmaer, Jr.

Langley Memorial Aeronautical Laboratory
Langley Field, Va.


WASHINGTON


NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of
advance research results to an authorized group requiring them for the war effort. They were pre-
viously held under a security status but are now unclassified. Some of these reports were not tech-
nically edited. All have been reproduced without change in order to expedite general distribution.


L 140


bOCUIMENTS DEPARTMENT





































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in 2011 with funding from
University of Florida, George A. Smathers Libraries with support from LYRASIS and the Sloan Foundation


http://www.archive.org/details/windtunnelinvesl0ang




I
', L1# '-


TACA CE Ioi. L4G10


CONPTDENTIAL


UATIOPAL KDVISOEY COr UITTEE FFOR AERONAUTICS


COIJTIDEiTIAL ST LETTIH


\iTI ,-TUNIr..EL T[iT"STIGATIOHN OF IACA 66(215)-216,

r6,1-212, 9AND 651-212 AIRFOILS WITH

O.20-AIRFOIL-CTH-ORD SPLIT FLAPS

By -;l. cien F. hl~l.-ner, Jr.


SUMMARY


.r if'e st ircat ion v:as carried o'lt in the !UTCA t'.o-
ldiiisionaal low-turbi.lence pressure tunnel of the
.ACA 66(215)-216, oC,1-212, and 651-212 airfoil sections
6qTuipped with split flaps having c'hord.s 20 percent of
the airf-il chord. The purpose was te determine ti-e
maxir:.urm-1ift chlracteri isics of these low-drasg airfoil
sections wvitn split' flaps. Ail the present tests :were
mad-e at Peynolds number of appromiiatel: o6 x l0' and
a ;ict-o nci-ber of about 0. 15

The nmaxi:num lift coefficients of these airfoils v:ith-
cat a--Jd J th fl.ap- :.re suirmmal'ized as follows:


Airfoil section


a'.'xa- imutrrl e c t i onl
lift coefficient
withoutut flapsiWith fla.ps


Flap deflection
(des)


HfAA 6 215)-216 1.56 2.61 70
iAJr. cI .,1-212 1..1 2.17 70
.N.. A _-212' l. 2.1 60



TIIR ODUCTITON


Extensive tests of split flaps and other types of
hi.ih-lift device used in conjunction with the older
conventional airfoils have been conducted in wind tunnels


CONFIDENTIAL









2 C 0~iT!D7I TAL I7ACA CB No. LLLC-10


and in flight. Because of the data available and
because of the simplicity of this device, the split flap
may conveniently be used as a basis for comparing the
_-axir::r-lift characteristics of various airfoil sections
e uipied with trailing-edge high-lift devices. The
present investigation was carried out in the tACA two-
dim3nsional low-turbulence pressure tunnel to s.jpply
inforr.stion on the m-.:r'imt.-l-lift and pitching-moment
characteristics of three low-drag airfoil sections with
split flaps.

The NACA 66(215)-216, the T.COA 66,1-212, and the
TACA 651-212 airfoils were equi-..Fed with split flaps
having chords 20 percent of the airfoil chords (0.20c).
Lift and pitching-moment data were obtained for each
airfoil for a range of flap deflection from 400 to 70
at a Reynolds number of aplroximately 6 x 106. Because
the tests were conducted at a Mach number of approxi-
matel 0.15, the results are believed to be free from
the effects of comoressibility.


APPARATUS AND !VT !ODS


T1Y. tests were made in the 7.,0A two-dimensional
low-turbulence pressure tunnel designated, TDT) by the
methods described in reference 1. All data have been
corrected for tunnel-wall effect. The ordinates for
the airfoils tested are ,.i-esented in tables I to III.

The 2-foot-chord models were constructed of
mahogany with chordwise laminations, anr the surfaces
were painted and sanded until aerodynni;.cally s rooth.
The split flaps were simulated by triangular blocks of
I.:lited mahogany attached to the lower surface of the
model. One face of the block was cut to the contour of
the flap portion of the airfoil lower surface. A
typical arr-ngement is shown in figure 1.


COT0 1 P 'TTAL


! L


---








.J;ACA C3 L'o. LJ410 COITFIDEiTIAL 3


RESULTS AND DISCUSSION [


T':.e section lift and pitching-moment characteristics
for th'-, F:ACA 6' 215)- 1, 6,-212, and 6 -212 airfoil
Ss ct i:': are presented in figures 2, 5, and I., r cs : -
tiel The lift and ,it chiing-.naoment characteri t i s
of' '. y i: lain airfoil are included for cc'ncfa._rison Wvit'h
.ie at ifoils with flans deflected.. .A comparison of the
3* -:.i.-.m lift coefficients of the three sections, tested
in 'Lth pr-esent investigation is given in fi-ur-e 5, with
sa_'iila data for the ;IACA 25012 airfoil from ref',er.nc., 2.
ZuioL shows the variation of the increme.nit of r.aximrum
section lift coefficient 7,ma with flap deflection
max
for the -arious airfol...

.Jr e;xa.:lin-ticn of figure 5 shcivs that higher maximiium
lifts :"ere obtained with the -lain U'ACA 051-212 airfoil
than 't:.th the plain -AC 6o,1-212 airfoil. Vehen the
flaps w.'ere defle-ted, however-, the mra:-:i.mu'n lift coeffi-
cients for both airfoi l were approxi:-.!tely equal. A
si ilar com iP:0ri son between the t'lwo ".CA -.,o-series
airfoils shows that considsrably i-;:iche ma;xirmumr lift
coefficients for all flap deflections were obttained with
the ic-rpercent-th.ick airfoil. The incrt i~e'nts cof .:a::imurii
lift coefficient for this airfoil sect on .. erle, on t-le
avera :-, 'j percent higher than the increments obiLain-'d
with the ACA 66 1-212 airfoil section. (See fig. .)
The increased mriaximum lift roefficients for the
NArCA 6.J(21)-216 airfoil ar- attributed to the greater
thickness and consequent increase in leadina,-ed_-e radius.
-:ui.-e r also shovvws that the ra:'imu-r. lift coeffi.-ients
cbtaiined with the plain iAC. 66(2 15)-216 airfoil at a
Re no dis number of 6 100 ere approximately the sane
as ;hose obtained from tects of the 'ACA 25012 airfoil
of re-frence 2 at an effective Reynolds n-umber
of 5. C 106. For most flap deflections tested, the
v."-..;.: of c x and Ac, (figs. 5 and o) obtained
'max "max
with the 1o-percent-thick low-drag airfoil were higher
than those obtained with the 12-percent-thick conventional
air'foi..


C OLTIDEl TI'AL








NACA CB No. LGl10


SI~TIEARY OF RESULTS


The maximum lift coefficients of three low-drag
airfoils without and with 0,20-airfoil-chord split flaps
obtained from tests at a Reynolds number of approxi-
mately 6 x 10 are as follows:


i M maximum section
Alift coefficient Flap deflection
Air'oil section !
i"ithout flaps With flaps deg)
NACA 66(215)-216 1.56 2.61 70
TNAA 66,1-212 1 2.17 70
'ACA 651-212 1 i.9 2.1560


Lanrlei MKemorial Aeronautical Laboratory
national Advisory Conmittee for Aeronautics
Langley Field, Va.









1. Jacobs, Eastman N., Abbott, Ira H., and Davidson,
Milton: Preliminary Low-Drag-Airfoil and Flap
Data from Tests at Large Re-nolds i'urb-ers and
Low Turbulence, and Supplement. FACA ACR,
.rch 1942.

2. '"enzinger, Carl J., and Harris, Thomas A.:
"Vind-Tunnel Investigation of N.A.C.A. 25012,
25021, and 23050 Airfoils with Various Sizes of
Split Flap. NACA Rep. :ro. 668, 1939.


CCO!FIDE:TIAL


CONFIDENTIAL






Jr.CA C:- To. LG1El0


TALE I.- iACA 66(215)-216 ATRFOIL

[Stations and ordinates arl gEiven
in percent of airfoil choi-di

i0per s r face Lovwer s's face

cion Ordin. t e S a t i I Or. i nate


U


.1-.01
S. 4
Cj 7
,J ,-4

l' ',



1_. CO6

r ,.725
,%0. 00C
'.95


'",. 067C
70.0;
75.087

*-.C75
:',0. C
DU. 055


0
1.2?0



8 4. -
S. iii0
1 r. -




3.20'
7.1' 6
d. ',L'O




S 092
.. ,:.O
8. 875




42
1 .. 61 7


3.T95
2.10$
; j -.J


0


1. 7

.10
7 .1) o


20.1 00
;i lo!


.5.0. 6
So.omb
k .111

,L.. 9
c, 0'2



79. '15
84.925
2, 9n
9o.972
100.000


0
-1l.10
-1.54

-2. ;.






-O .L22

-0 00.



-5.002

-o. 07~0

-2. 09


0


L.E. radius: 1.575
Slope of radius through L.E.: O. C;4


'ATTIC-AL ADVISORY
CO',-ITTT3.. FOR AE'RO-NAUTICS


CO PF IDENTI IAL


CONFIDE LTIAL


0





-i


U paper s-.u-face


;S tion I Ordinate


u 0
2 .-7
i 1: i. 150
i. 1.1 05
I 1.157 1.W7





6"12
F:, 2.7 .
l 7.579 J.l1

S. }-97


.. ~ 6.816


7.005
7.093
7.075
6. 36

7 507
/.195


3.75)
2.770
1.7b0
.792


L.E. radius: 0.S39
Slope of radius through L.zL',


o. cl3'


IATTIO;AL ADVISORY
C -'lJTTEB Pl ? AON C-AUTI CS


CO'rFIDE.TIAL


oo. 001


70. 0o6
75.066
?~o.n^ ,
tc. C :l

I,7" k C ;


---


*CT C, O o. L4GO1.0 COi.'IDCITIAL 6


TABLE IT.- ITACA 66,1-212 AIRFOTL

[s itions and: ordinates are riven
n -. f .' n "-a I, 7' 1 .n-.I ,.i


Lov.er surface

Stat on Ordin: t


0 0
.576
k.-4 --1. 010
1.53L -1.25
2.605 -1.61i
5.116 -2.165
7. 21 -2'. 5
10.122 I -2.965
1,.116 -3.553
20.105 6 .9'. .
25.01 -4 .522
50.075 -4.576
L,.C57 -. 756
0.c0. -4.956

50. 000 i -4. 09
:..%3l -.749
5 o I -4. L: 5)



1.9175 -. 167

100.40 -1
100L. JO 0


---'----






NACA C3 No. L4G10


TABLE III.- NACA 651-212 AIRFOIL

[Stations and ordinates are given
in percent of airfoil chord]


Upper surface


.970
1.176
1.491
2.058
2.919
.593
.162
5.073
5.770
6.300
6.687
6.9142
7.068
7. 044
6.860
6.507
6. o14
5.411
4.715
5.9546
5.140
2. 02
1.465
.672


I





















I
I


Lower surface


Ordinate


1.46
2.609
5.122
7.627
10.127
15.121
20.110
25.094
50. 077
35.o058
40.059
45.019
50. 000



74 .9 47
q.9 h8
0. o5,5

89.967
900.0009
100.000


0.952


Slope of radius through L.E.:


0
-.870
-1.056
-1.277
-1.6o6
-2.287
-2.745
-3.128
- .727
-4.176

-Lk.a26
-4.510
-4 .551




-2.771
-1.5)82
-L.026
-4.654

-3.17
-I .872

-2.771
-2.164
-1.548
-.Q^6

-. 046
0


o.o84


NATIONAL ADVISORY
CO0'ITTL': FOR AEROTAUJTICS


CONFIDENTIAL


.1423

1.154

7.573
19.675

',..66
1 .67Q

19.961
2C).Q2,
'. 142

50.000
55.017
60.032
65. OIL
70.050
75 -0~5
9o.052
90.0533
95.017
100,000


L.E. radius:


- --1` -- 'i


; _,___ __ __


CONFIDENTIAL










NACA CB No. L4G10 Fig. 1




HO
-













Si C


c00
o I
Cl 0





kO H-I




Id
S0
a)



rd

r-1


00








.,4
'I e




-3.t














2.8




2.h




2.0











0




S .8
C,
1.6




S1.2
o
4J


S.s
m-


-21,


Fig. 2


-16 -8 o 8 16 24l

Section anrle of attack, ao


S I I CONIDENTIAL
-.b 0 ., .8 1.2 1.6 2.0 2.1

Section lift coefficient, CL

Fl ure 2..- Section lift and pitcling-moment characteristics for an NACA 66(215)-216 airfoil
with a 0.20c split flap: Reynolds number, R, 6 x 106. Teats, TDT 247, 568, 571.


NACA CB No. L4G10







NACA CE No. L4G10 Fig. 3



.8 --- I [ I
CONFlDlNTIAL

i ~ ~ __ __ ---- -- -- -- -- -- -- -- -- -- -- --



f

1.6









S1.2 _
1 .2 --- -- -- --- -- --- -- -- --- -- --
















E .2/







0








0---- --0 ---



-.2
-2Jo -16 -8 o 8 l6
Section angle of attack, ao



E

___~ ___~ ___~ ___~ ___ __ __ NAlO0AL ADVISORY
c COIIMITTEE FOP AEfluNAIJrICS

5 0----------------------------------- --------------













C'


-.L 0 .to
.ectlon


.8 1.2
lift coeffciernt., c


1.6 2.0 2.h 2.8


Pi.lure .- ctr.lon lift ari 'rpi..:nri --omert .r.racteriestIcs for an NiACA t,i6,-212 alrfoli
witr, 0.20c sc.li I' lap; R', 1 .' Tesr. TDI L '2 .70, ,', 602.







NACA CB No. L4G10 Fig. 4



-.8 I I
CONFIDENTIAL





fI
__ 6f --- -- --- -- -- --- -- --- -- --- --

(dee)
.0 0 (plain airf.ll
+ 60
x to
0 70
S1.6




1, 1.2




o .



,' ______ __ ____















o NAOf'AL ADVIl:ORY















SIO I I I I JFIDET4L
P oir .- Sec__on lt nd nn-mn nCOMMII.NEE f AERCINAutI CS


0e
U 0-




_. -- ---- --_._-- -- -- -. .. -- -- -- -








J4 -0 .A 1.2 1.6 2.0 2.8
Section lift coefflcM Int, c&

Pljre. 4 .- Section lift and rltcnnri,-m, mEnt cnr er&teirlst1cI for r NACA 51-ei; airfoil
-lrh a 0.20C 3.11't flSr; P, .'. x i10 Tes[t T"DI .'.T .69, .''I,






NACA CB No. L4G10


JA2r-~


4 -'- + -t


/


/
- /


_NJACA 66(215)-216
R = 6 x 106 (approx.)

1ACA 23012
*Effective R = 3.5 x 10
(from reference 2)
lACA 66,1-212
R = 6 x 106 (approx.)

INACA 651-212
R 6 x 106 (approx.)


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS









CONFIDENTIAL

20 40 60 80
Flap deflection, 6f deg


Figure 5.- Effect of flap defiectior on maximum section lift
coefficient for the various airfoil sections.


CONFIDENTIAL
_R I I I I-


2.0



1.6.



1.2



.8



.4


r 1 1 I


' '


Fig. 5


V. 1111






NACA CB No. L4G10


CONFIDENTIAL
I


U________


.6


NACA 66(215)-216
- = 6 x 106 (approx.)



MACA 23012 6
-Effective R = 3.5 x 10
(from reference 2)

__i


-- iACA 66 1-212
R = 6 x lu6 (approx.)



IIACA 651-212
S= 6 x 100 (approx.)


0
NATIONAL ADVISORY
S / COMMITTEE FOR AERONAUTICS
S .2 /

' ----

CONFIDENTIAL

0 20 40 60 80
Flap deflection, 6f deg

FPL.ure 5.- Effect of flap deflectionL orL the increment
of max:iau section lift coefficient for the various
airfoil-flap arrangements.


Fig. 6


-







UNIVERSITY OF FLORIDA

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UNIVERSITY OF FLORIDA
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