The results presented in this note show the effect of twist and sweepback on the span load distribution over two monoplane wing models. The tests were made in the Atmospheric Wind Tunnel of the Langley Memorial Aeronautical Laboratory. The data are taken from the results of an investigation dealing primarily with lateral stability. As presented, they are suitable as an aid in the structural design of certain monoplane wings

Knight, Montgomery

Noyes, Richard W

English

NASA Technical Reports Server

TECHNI CAL NO TES 
RATIONAL ADVISORY COM"HTTEE FOR AERONAUTICS 
No . 346 
SPAN LOAD DISTRIBUTION ON TWO :.1O:·0PLA. E ViIl-G -WDELS 
AS AFFECTED BY TWIST AND Sv~ZP3ACK 
By Montgomery Knight and Richar d V'. :Toyes 
Langley Memorial Aeronaut ical Laboratory 
Washi ngton 
July, 1930 
https://ntrs.nasa.gov/search.jsp?R=19930081077 2020-03-17T04:20:53+00:00Z
NATIONAL ADVISORY CO~~ITTEE FOR AERONAUTICS. 
TECHNICAL NOTE NO. 346 . 
SPAN LOAD DISTRIBUTION ON TWO MONOPL ANE WING MODELS 
AS AFFECTED BY TWIST AND S~~EPBACK. 
By Montgomery Kni ght and Richard W. Noyes. 
Sum mar y 
The results presented in this no te show the effect of twi st 
and sweepback on the span load distribution over two monoplane 
wing models. The te sts were made in the Atmospheric Wind T~~el 
of t he Langley Memorial Aeronautical Laborato~y. The data are 
t aken from the results of an inve stigation dealing primarily wi t h 
lat eral stability. As presented, they are suitable as an a i d in 
t he structural design of certaim monoplane wings. 
I n t rod u c t ion 
In designing an airplane wing f or maximum structural safety 
and efficiency, a knowledge of the actual di stribution of the ai r 
loads along the span is e s sential. I f the wing is provided with 
e ither twist or swe epback, the distribution is no longer the same 
as f or a straight wing . However, comparatively little infor ma-
tion as to the changes thus produced has been published heretofore. 
This report has been prepared fo r the purpose of making avail-
able for design a limited amount of data on the aerodynamic char-
.. 
N. A.C.A. Technical Note No . 346 2 
acterist ics of two monopl ane wing models as affected by changes 
in twist and sweepback. The result s are taken from a serieS of 
pressure distribution tests made pr imarily for the study of later-
al stability. The tests were conducted in the Atmospheric Wind 
Tunnel of the Langley Memorial Aeronautical Laboratory (Reference 
1) • 
Models and Tests 
The shapes of the two semi-span models and the locations of 
the points at Which the pressures were measured are shown in Fig-
ure 1. One of the models , which was designated the N.A.C.A. 84 
wing, had the N.A.C.A. 84 airfoil profile from the root to the 
tip . The other, the N. A.C.A. 86 wing, had the N.A.C.A. 84 pro-
file at the root and the symmetrical N. A. C. A. -M2 profile at the 
tip. These profiles are shown in Figure 1, and their ordinates 
in per cent of chord axe given in Tables I and II. As may be seen 
from the figure , both wings had essentially rectangular plnn fo r ms 
except at the extreme tips wh ich were so designe~ that any tip 
cross section normal to the mean crunber line was 0.. semicircle 
whose diameter Was the wing thickne ss at that section. 
To permi t g iving the wings the desired amount of t'.-:ist, U 
speoial type of oonstruction WaS necessary. Each wing was made 
up of 3/16 inch mahogany laminations mounted parallel to the pl ane 
of symmetry and clamped together by two long internal bolts run-
ning spanwise. By loosening the bolts and rotating the lamina-
tions relative to each other through a small angle, any desired 
N.A.e.A. Technic~l Note No. 346 3 
geometric twist up to 150 Was obtainable. The IIsteps" on the 
sur f ac e of "the wing formed by the lruninations slipping past each 
other were f aired over with pl~sticine. 
The swe epback adjustment WaS made by rotating the wing in 
its mounting bracket about an axis normal t o the midspan chord at 
its 50 per cent point and lying in the plane of symmetry. 
The semi-span wing models were used in conjunction with a 
s eparation plane as shown in Figure 2, which is a view of the tun-
nel se t -up looking upstream. Thi s r,1ethod has been used almost 
exclusively in making pressure di stribution tests in this wind 
tunnel Qnd is based on the assumption that the imaginary plane of 
symmetry of the wing can be replaced with an actual plane surface 
wi thout Qffecting the flow. It t hen becomes necessuxy to study 
only one-hQlf of the wing model. The results obtained by this 
method of tcsting have checked satisfactorily with the results of 
other methods, showing that the assumption is reasonably valid. 
The apparatus and test procedure in general are described 
in Reference 2, with the exception of the manometer. This instru-
ment was unusual in that it was designed to integrate automatic-
ally t he pressures over each test section so as to give the total 
sec tion load in a single reading. A detailed description of the 
principle of operation and of the design of the instrument will 
be published in a later report. 
Th e test program on each wing included the following varia-
bles : 
N.A.C.A. Technical Note No. 346 4 
1. Geometric twist - washin ~ 5°, 
washout = 0° 5° 10° 15° 1 1 1 , 
2. Sweep - sweepforward = 20° , 10°, 
° ° ° sweepback = 0 1 10 , 20 , 
tested at angles of attack of the root section ° = -9 , ° -6 , ° -3 1 
0° , 3° , 6° , 9° , 12° , 15° , 18° , 21° , 24° , 27° .. , 30°. 
During all tests the dynamic pre ssure WaS held constant at 
1.25 pounds per square foot, corresponding to an air speed of 
about 22 m. p . h., and a Reynolds Number of 160,000. 
Res u 1 t s 
The results are presented in absolute coefficient form in 
the following two groups of curves which show the effects of 
twist and sweepback on the two monoplane wing models: 
1. Semi-span load distribution (Figs. 3 to 6). 
2. Normal force versus angle of attack (Figs. ? to 10). 
In the first group three standard loading conditions used 
in structural analysis are represented, viz., nose dive, low an-
gle of attack, and high angle of attack. These conditions are 
taken as occurring at CNF = 0, 1/4 max. CNF and max. CNF , re-
spectively. The data from which these curves were plotted are 
given in Tables III and IV. 
To determine the magnitude of the section normal load per 
uni t span , N' , for any given set of conditions shown in Figures 
N.A.C.A. Technical Pote No. 346 5 
3 to 6, the following equation may be used: 
NI = 
where CNFI = absolute coefficient of section normal force, 
c = chord, 
q = dynamic pressure. 
The tot al normal force N, on a wing may be obt ained from 
the dat a in the second group of curves as fOllows: 
whe re 
N = q S CNF 
absolute coefficient of total 
nor mal force, 
S = total area of wing. 
In general, the data may be considered accurate to within 
±5 per cent . 
Langley Memorial Aeronautic al Laboratory, 
National Advisory Committee for Ael'onautics, 
Langley Field, Va., June 23, 1930. 
1. Reid, Elliott G. 
2. Re id , Elliott G. 
Ref ere n c e s 
standardizat ion Tests of N.A.C.A. No.1 
Wind Tunnel. Y.A.C.A. Technical Report 
No. 195 (1924). 
Pressure Di stribution Over Thick Tapered 
Airfoils, N. A.C.A. 81, U.S.A. 27-C Mod-
ified, and U.S.A. S5. N.A.C.A. Tech-
nical Report No. 229 (1926). 
N.A.e.A. Technical Note No . 346 6 
TABLE III 
N.A.C.A. 84 Section Normal Force Coefficients 
Wing Semi-span 
CNFI 
condition CNF sec. sec. sec. sec. sec. 
A B C D E 
Straight 0 .000 +.010 .000 -.015 +.012 
1/4 max. .440 .384 .300 .230 .160 
max. 1.550 1.573 1 .473 1.230 1.396 
5° washin 0 -.030 +.027 +.050 -.005 +.005 
1/4 max. .400 ~ 425 .395 .244 .244 
max. 1.510 · 1.364 1.430 1.256 1.600 
5° washout 0 +.040 -.025 -.040 -.112 -~052 
1/4 max. .465 .380 .275 .125 .110 
max. 1.570 1.600 1.468 1.155 1.308 
10° washout 0 +.080 -.060 -.096 -.187 -.145 
1/4 max. .515 .310 .152 .045 .052 
max. 1.555 1.548 1.287 1.015 1.139 
° 15 washout 0 +.108 -.070 -.155 -.250 -.168 
1/4 max. .530 .265 .070 -.035 •. 000 
max. 1.570 1.533 1.270 .964 1.208 
20° sweep- 0 +.028 -.011 -.010 -.015 -.140 
forward 1/4 max .395 .354 .284 .250 .250 
max. 1 .340 1.613 1.570 1.461 2.552 
10° sweep 0 +.027 +.006 -.020 +.025 -.089 
forward 1/4 max. .430 . 380 .315 .246 .206 
max. 1.475 1.579 1.510 1.344 2.073 
10° sweepbaek 0 +.020 -.002 -.026 -.002 +.018 
1/4 max. .406 .380 .295 .212 .065 
max. 1.52.7 1.345 1.290 1.060 .630 
20° sweepback 0 +.010 +.001 -.002 -.002 .000 
1/4 max. .390 .350 .308 .207 .016 
max. 1.612 1.340 1.164 .897 .078 
N. A.C.A. Techn ical Note No . 346 7 
TABLE IV 
rZ . A. C. A. 86 Section Normal Force Coefficient 
CNF ' 
Wing Semi-span 
c ondition CNF sec . sec. sec. sec. sec. 
A B C D E 
Straight 0 .078 -. 080 -.130 -.094 -.089 
1/4 max. .400 . 225 .165 .100 .037 
max . 1 . 282 1 . 160 .995 .795 .913 
5° wash i n 0 . 020 -. 020 -.030 -.030 -.020 
1/4 max. . 335 . 268 .230 .150 .090 
max. 1 .175 1.058 .940 .845 1.025 
° . 115 -.180 -.160 -.140 5 washout 0 -. 060 
1/4 max. .430 . 240 .090 .030 -.012 
max. 1 . 292 1. 172 .935 .720 .730 
10° washout 0 . 150 -. 080 -.220 -.200 -.180 
1 /4 max. . 458 . 200 .024 -.030 -.045 
max. 1. 342 1. 170 .919 .661 .644 
° 15 washout 0 . 190 -. 140 -.338 -.300 -.322 
1/4 max. . 510 . 176 -.054 -.123 -.118 
max. 1 . 355 1.200 .900 .625 .555 
20° sweep 0 . 080 -. 038 -.108 -.087 -.194 
f orward 1/4 max. . 410 . 242 .185 .122 .185 
max . 1 . 190 1. 086 1.046 .912 1.770 
. . 
10° swe ep 0 . 050 -. 105 -.130 -.105 -.190 
f orward 1/4 max. . 385 . 200 .130 .090 .070 
max. 1 . 25J 1.110 .971 .880 1.290 
10° swecpback 0 . 0 47 -. 075 -.105 -.082 -.020 
1/4 max. . 354 . 235 .156 .090 .020 
max . 1.215 1. 070 .872 .670 .305 
20° sweepback 0 . 060 -. 060 -.112 -.094 -.020 
1/4 max. . 350 . 211 .140 .100 .030 
max . 1 . 255 . 996 .721 .590 .260 
N.A.C.A. Technical Note No.346 
....... a 
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Fig .1, 
Continued. 
a= N.A.C.A. 84 Tip 
b= N.A.C.A. 86 Tip 
Fig.l Orifice 
locations 
and profile or-
din~tes for 
K . P U. A. 84 and 
K." A. C . A. 86 
se:-Ji Epan 
wing r.1odels. 
N.A. C.A. Technical Note No.346 Continuation of Fig.1 
5.45% 
5.45% 9.28% 12.3Q5b 9.287t I 1.30~. ~~. 4h2.~4.2~4.8~ ~4.23~\ + ~ -+ ~I 1.3~ 
I I": , I i I 
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< 2'8. 33%~ , ' 
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<: .83.97%------->1 
, 
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~------------------- - laO .. "'1 \U---~ 
N.A.C. A. 84 Profile 
Table I. 
Profile 
ordin[l,tes 
Stations N.A. C.A. 84 
in % 
chord Upper Lower 
0 2.50 2.50 
1.,25 4 .85 .95 
2.50 6.05 . ·11 
5.00 7.78 .10 
7.50 9.03 .02 
8 .50 --- 0 
10.00 10 .00 0 
15 .00 11 . 50 0 
20.00 12.71 0 
25.00 13.51 0 
30 .00 14.00 0 
35.00 14.18 0 
40 .00 1 ,1 .11 0 
50 .00 13.50 0 
60 .00 12.31 0 
70.00 10.32 0 
80~.GQ 7.71 0 
90 .00 4 .39 0 
95 .00 2.41 0 
100.00 .30 0 
N. A.C . A. Technical Note No.346 Continuation of Fig.l 
5.45% 9.28% 12.30% 12.30% 9.28% $.45~ 
1.3~ - p+ ! 1'- ! *-l.4.2~~4.8~~4.2~ I -r-11.J~ 1.30% 
I I ' I I 
1.3~ _ I'C- ~ I 
6. 75% 16.03~ J' I 
-213.33%- .. _~.J 42.56~~ .1 ~-----57.44%~--?>, I ' 
71.67% i I 
I<E:-~------- 63.97% '> 
93.25~ --~~ I 
'L----------------98.70% "l 
K:---------------- 100~::lbll 
N.A.C.A. 86 Profile 
Root section N. A. C.A. 84 
Tir section N.A.C .A. Y2 
Table I I 
Profile ordinate s 1 
Stations N.A.C.A . 84 N.A.C.A. M2 
in % 
-
chord Upper Lower Upper Lowe~ 
0 d. bO 2 .bO 0 0 
1.25 4.85 .95 1.30 -1.30 
2.50 6.05 .41 1.74 -1.74 
5.00 7.78 .10 2.33 -2.33 
7.50 9.03 .02 2.74 -2.74 
8.50 --- 0 --- ---
10.00 10 :00 0 3.05 -3.05 
15~OO 11. 50 0 3.49 -3.49 
20.00 12.71 0 3.78 -3.78 
25.00 13 .51 0 
--- ---
30.00 14 . 00 0 4.03 -4.03 
35.00 14 .18 0 --- ---
40.00 14 . 11 0 4.00 -4.00 
50.00 13.50 0 3.74 -3.74 
60.00 12 . 31 0 3.30 -3.30 
70.00 10 .32 0 2.71 -2.71 
80.00 7.71 0 1.99 -1.99 
90.00 4.39 0 1.15 -1.15 
95.00 2. 41 0 .69 -.69 
100.00 .30 0 .20 -.20 
N.A.e.A . Technioal Note No.34S Fig.3 
Fig.a Wind tunne~ set-up of twisted wing. 
I .A .O.A. Technical Note No.346 
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N.A.C.A. Technical Note No.346 
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N.AJC.A. Technical Note No.346 Fig.5 
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Span Load Distribution on Two Monoplanes Wing Models as Affected by Twist and Sweepback

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