Document
'T T.'; R No.
L5119
rd
2-51945
SE P
NATIONAL
',rOR AERONAUTICS
h' -1k
z^^
MEMORANDUM
REPORT
f or the
Air Technical Service Command, Army Air
' Forces
AERODYNAMIC CHARACTERISTICS OF FOUR REPUBLIC AIRFOIL
SECTIONS FROM TESTS
IN LANGLEY TWO-DIMENSIONAL
LOW-TURBULENCE TUNNELS
By Miit ' on M. Mein
Langley Memorial Aeronautibal Laboratory o Lafigley Field . , Va.
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A C A UIBRARY
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IVIR Y a . I, 5119 ,_
NATIONAL ADVISOR' COr;II r =TEE FOR AERONAUTICS
a REPORT
MEI"TC!I 3 ` 1 1 IND UM
for the
Air Technical. Service , Command, Army Air Forces
AERODYNAIPTIC' CHOP ^CTERTSTICS OF FOUR HEPUff.JC !JRFOIL
S7'CTIONS FROM TES'T'S IN LA.NGLTY T'ri10-DI^4rNSI01TAL
I-0'i1-TSRBULENCF; TLTNNTLS
By TAil ton NI. Klein
SUMMARY
Four airfoils sections, designed by the Republic
.nviation Corporation for the root and tip sections of the
XF-12'airplane, were tested in the Langley two-dimensional
low=turbulence'-,tunnels to obtain their aerodynamic charac-
teristics. Lift characteristics were obtained at Reynolds
numbers of ;,000,000, 6,000,000, 9,000,000, and 14,000,000,
characteristics were obtained. at Reynolds
whereas dra g
numbers of 3,000,000; 6,000,000, and 9,000,000. Pressure
distributions were obtained for one of the root sections
for several angles of attack at a Reynolds ,number of
2,600,000.
Comparison' of the root section that- appeared best
from the tests with the cori , esponding xTACA 65-series
section shows the Republic section has a higher maximum
lift and higher calculated critical speeds, but a higher
minimum drag. Ir_ addition, with standard roughness applied
to the leading edge, the maximum lift of the Republic air-
foil is loner then_ that of , the NACA airfoil.
Comp arison of the Republic tip section with the corre-
sponding NACA 65-series section shows the Republic airfoil
has a louver maximum lift ancl a higher minimum drag than
the NACA airfoil. The calculated critical speeds of the
Republic section are slightly higher than those of the
NACA section.
i
7 11 17
MR No. L5z19
I ITT RODUCT I'ON
At the reauest of the-Army Air Service Technical
Comrrrand, tests were made to obtain the_a..erodynarnic charac-
teristics of four airfoil sections designed by the
Republic Aviation Corporation for the XF-12 airiplane.
The airfoils, which had 'thickness ratios of 13 and 18 per-
cent, correspond.ing to the tip and root sections of 'the
wing, were designed to have moderate extents of laminar
flow. The , •lif..tand drag characteristics were obtained by
tests conducted in the Langley two-dimensional low-
turbulence pressure tunnel (TDT). one of the 18-percent
thick: airfoils was tested in the Langley two-dimensional
low-turbulence tunnel (ITT) to obtain pressure distribu-
tions at several angles of attack.
SYMBOLS
a fraction of chord for which mean line loading is
constant
cti section lift coefficient '(l/qc )
c Z maximum section lift-coefficient
max
c-d section drag coefficient ( d /q_e)
minimum section drag coefficient
cdrl-in
T —
h
p
e.: low-speed pressure coefficient .
S
q
Z lift per unit spar_
d drab per unit span
C,
ai.rfoi.l. chord
I
H '• -free-stream total pressure
local static pressure
p
free-^stre`am 'aynLimic pressur6.
q
Mcr critic21 14ach number
1AICLASSIR
4 t ^
1V1R L5119 N0 • 3 R Reynolds number
a o
section angle
of attack, degrees
MODELS AND TESTS
The models, submitted b y
the Republic Aviation
Corporation, were- of laminated wood construction and had
a chord length of 24 inches.
The models were designated
as follows:'
' R - 218 - 1 ;
R-4,4 0
-413-6; - 4,4 0 R-4,40-318-1;
R
In this designation the symbols have the -4,50-418-1.
following meaning;
R Renublic Aviation Corncration_'
of
first digit
family airfoil
next two digits
location of'maximum thloknesslof
airfoil in' percent ch-ord,..
next digit
design cL
next two digits
maximum thickness of airfoil
last digit
value'of a
for mean line of
a i rfo i l-
Ordinates f6r these'airfoils,
supplie i d by the manufactu
'ror, are presented in tables I through IV.
The method o.f,;- `.
testing was the same as that described in reference 1,..-f•or.
.
2-fqot
chord models'in the TDT.
Lift data were obtained 2t
Reynolds numbers .of 3;000,1000, 6' ) 000,000,
9,00b.,000 and 1LL,000,000 and. drag data •
t* 3,0.00,000,
6,000,000, and
Lift
9,000,'000. and drag with standard'..:,",-
roughness applied to the leading edge (reference
. were
obta i
ned at a Reynolds ,number of
6 x 10 6 . .In addition, s
r pressue di
tributions were obtane
d for'L.
the R-4,4o- 3 8
at a Reynolds number of
2.6 X 10- - ,. for angles of.att,ack
of -1.02 0
0 0 ', and' 2.03 0 .
-) The highest Mach number'
encountere.d during the tests was'less than 0.16;,there'fore
the results may be considered free,of'compressibility
effects.
Corrections for wind-tunnel. wall interference were
e made by the following equations wh re the primed quantities 7 cz, -Y 4, 41 F I E D IC UAS PF-11 L 511 L MP-M IMF No.
represent the angles of attack and. aerodynamic coeffi-
cients uncorrected for tunnel wall effects.
a 1.015- ao C d
0.0
34A), 1 c d (3)
c (0-985 - 0.0 34A) cz
.031.) S
(1 - 0 (4)
to be used
The value of the correction factor A
in the 'above equations is given in the following table:
mi o de 1 A
2 1
R-^-,40-41;^ 6
R- ) ,Lo-21 ^ 8-1
• 343
R-4, 0-W
L 3 01,
4 1 -
I
R- c 1." 0- 18-1
-395 1 '
of equations (1) through showsthat the
Examination,
(4)
' corrections are of the order of only a few percent. An
explanation of the tunnel wall corrections a pplied to
data obtained ,from the Langley.'t"No-dimensional tunnels is
given in the a p pendix of reference 1.
RESULTS AND DIscussibN
'air-
T he e lift -and d- .,--.g characteristics of the four
foils . 1 thro Pre'ssurb dis
are presented in figures ugh
4.
tributions for , the R-1 are presented in' 'f*1'g'.u res
,40 -318-1
5 4^,40
h
for several angles of attack. The - erc'ent
D
throug 7 13 :
has
section. ( R - designed for the tip section,
►
,40 -41 3-^.. 6 ),
a high maximum lift -arid a low minimum drag relative 'to
airfoils and thickness, and a range of -comparable shave
.
'
of 'C L .0f.0.2 for low-drag at a Reynolds number of 9 x 10 6
(f i g. it may there ore be considered. satisfactory.
f
1)
Although the R -418-1 airfoil - section,*which was one
-4,5 0
ection, has a Igne,d f or. th.,'.'r'oot 'section, of the , airfoils des, .
moderate maximum -lift 'and' a low minimum drag, it also has
shows an extremely rapid
a very narrow , low-drag range , , arid-*-
increase in drag for positive lifts outside the loir-drag, range (fig.
This airfoil. is therefore considered
4)-
MR 04
L509
unsati's'factory for -the root section. The two alternate
18 percent sections, R-4,40-218-1 (fig. 2) and,R-4;40-318-1
( fig. 1) have .minimu1-drags somewhat greater than. the.
However, .t hey. have . a
minimum, drag of the
-R-4,501418-1.
for low drag and show only a moderate
greater .range of 0 1
increase in drag for positive Lifts outside the low-drag
range. While, the two alternate sections . have . approximately
'the same dr.'ag characteristics, the maximum lift of the
318-1 was greater than that of the-R-4,40-218-1.
R- 4,40-
. the best
on this basis the R-4,24.0-318-1 appears to be
choice for the root section.
The Republic'sections have ,.pressure distributions
somewhat similar to the NACA 65 series. i comparisod has
therefore been .made of the aerodynamic characteristics..of.
the R-4,40-413-:6 and R-4,40-318-1 with those of the
corresponding NhCA 65-series airfoil sections in the
following table:
or
Airfoil
cZmax^ _cdmin c^max
}. 10 6 R - 9 x 10 standard roughess,
` section 'R:_
lot
R = 6 :X
R .6 ^ 1..18
- 4,40-4 1 1 - 1 59 0-0047
1 .0039 1.30.:., ^NACA 65 -13 K.63
4,4o-318-1 1.59 .0049 1.13 _
^R- I kNACA 65 3 -318 1.53 ,0042 "127"
I
+ I I
-__ __ - J
- - .-
The value s r for -the NACA sections have been .obtained ,by.: interpolation ' of the values given in reference 1. 'to the , ., proper camber and.thickness.
In addition, a com
p
arison of the theoretical critical.
speed curves of the R-4140-413-.6 and R-4,40-318-1 and the
corresponding NACA -
65-series sections has been made in
figures 8 and 9, where critical Mach number is plotted
against low-speed lift coefficient. The critical speeds
for the Republic sections were obtained by calculating
their theoretical pressure distributions for variouslift
coefficients; the critical. Mach number corresponding to,
the maximum value of S in each pressure distribution was
then obtained from the plot of critical Peach number against
low-speed pressure coefficient as given in reference 1.
The critical-s p eed curves for the NACA sections were
obtained by interpolation of the critical-speed data for
the NACA 65 series as given in reference 1. Critical-speed
C UNGLASPE-^UF01- t
lo
11%1144 111 WITS
5TJ9
6 R ITO L
values for the • obtained from the e7perimental
R-4,40-318-1
pressure, distributions f igs. 6, and 7) have been
'( 5,
s
potted on f igure ood agree y r.ent with the theoretical
9. G , values - mav be observed.
Comparison of the charactris^tics of the Republic
s
and section shows that, for a thickness ratio of
NACA
13 percen.t, , the airfoil hc .s a higher maximum lift
NACA
with
and without standard roughness and a lower minimum
drae,
than the Republic section. The crit i cal speed
values are approximately the, same except at higher lifts
where the critical speeds'oP the NACA airfoil are lower
than those of the Republ i c airfoil. For, a thickness ratio
of 18
percent, 'the . Republic section has higher critical
speeds and maximum lift but a. somewhat lower maximum- lift
with standard roughness* and. a higher minimum drag.-These
results indicate that the Republic R- and '
,0-318-1
s ae equally .suitable for -318 airfoil section' r
NACA
65 3
use as the root section of the XF-12 airplane; however,
the a pp ears o be more suitable tr:an the 't
NACA
65 p -413
Republic R-4 for thetip section.
.6
r ,4o-413
CONCLUDING REMARKS
.6
1. The Republic 13-percent airfoil
R-4.4o-413-
has a lower maximum lift and a higher minimum drag than the
NACA 65 section the calculated critical speeds, 651.-413;
of the Republic section are approximately the same as'
those of the
NACA section except at the higher lifts where
the Republic section has slightly higher critical speeds
than the
NACA section.
2. The Republic 18 - p ercent airfoil has
R-4,40-318-1
a higher maximum lift and hi gher calculated crit'ibal
speeds than the
-6 NACA section -318, but a'somewhat '
5 a
i
^,Aj'ith stan and rougl-inp ss,applied to
higher m nimum drag
j
,..{ a ti-:_ ^ Ste* , iz -..
y
the leading edge the maximum lift of the Republic section
is lower than that of the NACA section.
Langley, Memorial Aeronautical Laboratory
National Advisory Committee' for Aeronautics
Langley Field, Va.
Milton M. Klein
Physicist , Approved:
'r/
Clinton H. Dearborn
Chief of Full-Scale Research Division
FMJ
SSIFIED
t F
Y,R No• L5I19
rte
A '
' RETER LACES
and. Stivers, Ira H., von Doenhoff, Albert E.
1. Abbott,
Data. IdACA ACR
Summary of Airfoil
Lours S., Jr.
No 1,5005 :,
1-945-1-
f NYC MR No. L5I19 ,.
TABLE I.- ORDINATES FOR THE TABLE II.- ORDINATES FOR THE REPUBLIC R-4,40-413 -.6 AIRFOIL SECTION REPUBLIC R-4,40-218 -1 AIRFOIL SECTION' • • and ordinates given Stations and ordinates given Stations airfoil chord In percent of in percent of airfoil chord
^
Surface Lower Surface Surface Lower Surface Upper U pper • Station Ordinate Station Ordinate Station Ordinate Ordinate Station 0.50 1. 08 0.50 0.709 0.50 1.600 0.50 1.220 1.510 .75 1. 67 .888 •75 •75 1.9954 •75 2.09 1.25 1.175 1.25 2.495 1.25 1. 70 1.25 2.810 2.50 2.92 2.50 1.646 2.50 46o 2.50 2.231 5.00 4:770 5.00 5.00 4.120 5.00 33 935 7.50 5.019 7.50 2.609 55 740 7.50 7.5 0 4:765 10.00 10.00 2.869 10.00 6.520 10.00 55.390 5 . 771 q. 25 15.00 6: g 30 15.00 3.238 15.00 15.00 6.325' 20.00 8.210 20.00 6. 80 20.00 818 20.00 3. t^ 5 9 25.00 3.606 25.00 9.242 25.00 7.30 2:467 25.00 30.00 9.69a 30.00 7.7 3o.00 8.9 3 8 30.00 to 3.65h4 35.00 9.9965 35.00 7.885 35.00 9.247 35.00 3.656 40.00 10.070. 40.00 40.00 40.00 3.6o 9.399 7-225 45.00 10.010 45.00 7.d3o 9.399 45.00 50.00 242 50.00 99.805 .600 50.00 550.00 3.3 3 4 2 5.00' :922 55.00 3.140 777 60.00 2.896 . 20.00 bo:00 6755 2 0.00 8.4.29 8.910 65.00 7.736 65.00 2.653 65.00 8.230 65.00 6:130 2.886 70.00 70.00 2.398 7 0.00 7 395 770.00 -385 75.00 2.101 7 75.00 55 879 4: 91 80.00 80.00 1.765 80.00 b^35 80.00 :68 35 5 11^ 3.638 85.00 1.402 85.00 85 Co 90.00 6 90.00 1.002 2.880 90.00 7 90.00 2.4 1:70 .00 1.500 95.00 88 100 .00 1.039 1 00.00 •0 100.00 0 100.00 0 L. E. radius't 1.143 L. E radius: 2.250 Slope: 0.21297 Slope: 0.07807 l y TABLE III.- ORDINATES FOR THE TABLE IV.- ORDINATES FOR TAE 8 - 1 AIRFOIL SECTION REPUBLIC R-4,50-418-1 AIRFOIL SECTION REPUBLIC R-4- ' 40-31 Stations and ordinates given Stations and ordinates given in percent of • air£oil chord in percent of airfoil chord Upper Surface Lower Surface Upper Surface Lower Surface_ Ordinate Ordinate Station Ordinate Station Ordinate Station Station 1.759 0.50 1.119 0.50 1.545 2 0.50 0.928 .50 .?5 2.084 .75 1112 1:25 2 : 398 1.25 1.25 2.609 1.25 1.885 1 2:700 2.50 2.50 2.267 2.50 33 5 5 2.50 354 5.00 5.00 768 .713 3.170 5. 00 5 9 9 27 5.00 .9 80 5 7.5 0 772, 7.50 0 1.00 2:8213 10.00 5.103 10.00 4:31 15.00 8.089 15.00 55 73 7.850 5:8 15.00 15:00 20.00 20.00 9.023 20.00 6:5692 8.834 20.00 5.57 25.00 25.00 9.630 25.00 25.00 9.707 6.986 55 30.00 10.183 30.00 7.248 30.00 10.239 30.00 6:267 35.00 10.679 35.00 35.00 10:482 35.00 6.L82 7.3799 40.00 40.00 10.960 40.00 6. 33 40.00 10.609 7.392 7.281 11.159 6.763 45.00 10.569 45.00 45. 45.00 50.00 11.223 50.00 6.801 50.00 lo.365 50.00 7.652 555.00 11.139 555.00 6.765 b 25-0 0 9 20.00 20.00 0.00 10.882 99 6-5k 65.00 65.00 6.1 65.00 8:742 65.00 8.625 10-370 0.00 99 5.00 6.8 8669 775.00 44:2129 7 8.4953 775:00 75:00 080 80.00 7.138 80.00 3.822 80.00 80.00 55 77333 3.286 8 00 85.00 2.4199 85.00 5.514 85.00 2.65 4:4 4 0 0 3 7066 90.00 90.00. 3. 1 90.00 1.5534 6 1:50 5.
0 00 5 1.6 3 9 95.0 95.00 •b7? 9 100.00 .017 100.00 .017 100.00 100.00 0 0 L. E. radius: 0. 48 L. E. radius: 1.620 Slope: 0.254 Slope: 0.1689 '$ NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS ^k+A'L' •.
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^' MR No. L5I19 MR,No. L5119 2.2 2.0 1.8 O Upper surface Lower surface Ei ca 1.6 -P C.
1.4 ID 2.2 m C.
P.
1.0 .8 .6 .4 I L .2 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 80 go 100 0 10 20 30 60 40 50 70 Percent chord airfoil section; Pressure distribution for the Republic Figure B-4,40-318-1 5-- 2.6 x 106.
— 1.02°; R, LTT test 379; a o I Teti to ; MR No. L5119 "ED U i-,SiS i 2.0 1.6 Q W 1.2 1.0 r4 .6 .2 0 10 20 80 100 30 40 60 go Percent chord Figure 6, Pressure distribution for the Republic R 1 airfoil section; 4,40-318-1 x -106.
R, LTT 2.6 test 379 0 0; cL o MR No. L5I19 2.2 G^f^'LL 2.0 1.8 O Upper surface Lower surface 1.6 m 1.4 H U H 4-1 0 1.2 m W m a 1.0 .8 .6 .4 .2 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 0 10 20 30 40 50 6o 70 90 100 8o Percent chord R-4,40-318-1 airfoil section; Figure 7.- Pressure distribution for,the Republic R, 2.6 x 106.
LTT test 2.03° ,; 379; a o MR No.
L5119 -j- F-4 r7.; 7: :---t 7 lir: 1; 7 i- ;r 77r 7 :r- :.I..
r-:: =r .
7 17 "1 7: 4 ... ... ... .... . - _ _ --r : : v^ 1 - :i: : :I iL :;7:. .7- -ii: 7 1 ri i —=4= a: 4 7— =---7 ri Li Zrl— ii i iii --t i. r- :4 Zil: r:_7 rI -7 rrii i 1 i: 2AAL ii1Z iiii. ... ... . .... .....
..... .... ........ rr - i: 17. tri i., :T- I.: . - T . .... ....
rr- Z .
7 7: -
...
Rep ublic R-4o4o-413—.6 7-iz: :: *-...,:-
-NACA (reference l)-
r 651-413 . .... ..... ..
r: Hi: 1: -a - .. .... .... ...
.... .......
.. .... .....
i:,: .... .... .... ...
t::: t!:;:: 17 7-' ja r: it t':' P .... .... .... ........ ....
% C3 Z., ; -.7 . .... .... ...
14 ; ...
.6 .. .... .... ....
... ...
Cd .... ... .... .... .
. .. . .... .... :!; := .,I . .... ... .... ... .... ..
.... ... ... .... . .... ..
rl ... ... . . .... .... .. .... .... .
.. ... .... ...
3- ... .... ....
.. . . . . ... ... .... .... ..
...
.... .... ...
.... .. . ....
% .. ... .... .
....
... .... ... ... ..
NATIONAL ADVISORY AERONAUTICS FOR COMMITTEE 777 77—
F
777-
Em^A L
........... ... . ... ... ...
I -E Low-sneed section lift c. coefficient.
Variation of critical Machpupber wit4 low speed section
Figure 8.-
and the NACA
lift coefficient for the Republic
R-4,40-413,-- b 651-413
airfoil sections.
MR No. L 5 I 19 _ :
ens sue- -
L ....... _ ......................::::::- :.. _.__ ._ Republic ' R-4,40-318-1 ::.. -::
- -- _ - --- NACA 65 3 -318 (reference 1)
:....._ .._..t..._....,.._....._ ....._.. .....
/:=:: (experimental) o Republic - _ - _- - - f - - - - - _- - — _ w _ :i ....
777 N c ?i T i U r.: F .; U ii i; :t ... ; :`: .......
:: ci:::;::::::::: :i: NATIONAL ADVISORY i: i .iii
ii' `: : ii: :::::i' i
::':::::::::COGS E FOR AERONAUTICS .... .. ........ . ...:.. .:: ::;:: MITTE : ' :.: >.: — :`:.^' iii iii ii i• _. ........ ..
, ^
Low-speed section lift coefficient, cZ
Figure 9.- Variation of critical Mach nusaber with love speed ' section
653-318
lift coefficient for the Republic R-4 1 ,40-3 18-1 and the NACA airfoil sections.