Schneider, William C

English

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RESEARCH' ,MEMORANDUM 
 his material contam information affecting the National Defense of  the Unlted States within the meaaing 
of the espionage laws, n t l e  18, u.s.c., secs. 793 and 784, the transmission or revelation of which in  an^ 'i 
manner to an unauthorized person is prohibited by law. .. L ?,., . . ' :  
WASH.1 NGTON 
April.18, 1956 
https://ntrs.nasa.gov/search.jsp?R=19930089158 2020-06-17T08:35:01+00:00Z
I 
~ 
NACA N'I L56B14 -
NATIONAL ADVISORY COMMIT"J%E FOR AERONAUTICS 
RESEAR% MEMORANDUM 
EITECTS OF REYNOLDS I"BEB AND LEADING-EDGE SHAPE ON 
THE LOW-SPEED LONGITUDINAL STABILITY OF A 
6-PERCENT-THICK 45' SWEIrllBACK WING 
By Wiliam C. Schneider 
SUMMARY 
Force-test  data showing the  e f f ec t s  of changes i n  leading-edge 
radius (nose shape) on the low-speed long i tud ina l  s t ab i l i t y  of a 6-percent- 
th ick  45O sweptback wing of aspec t  ra t io  3 are  presented for  several  
ranges of Reynolds number and Mach nmiber. Data are  presented for  wings 
with NACA 0006-( 4.76)3 , NACA 0006-( 6.74)3 , and  modified NACA 0006-( 3.88)3 
airfoi l  sect ions with leading-edge radi i  of 0.0025, 0.0050J and 0.0025 
of  t he  loca l  wing chord p a r a l l e l  t o  t h e  p l a n e  of symmetry, respectively,  
r 
and f o r  Reynolds numbers up t o  6.6 x lo6 and Mach numbers up t o  0.20. 
This study i s  a cont inuat ion of  the s tudies  publ ished in  NACA RTlI ~ 5 5 ~ 0 6 ,  
NACA €331 L55H04 , and NACA RM L55L30. 
The resu l t s  ind ica ted  tha t ,  as t h e  a i r f o i l  s e c t i o n s  were thinned a t  
t h e  nose, separation effects occurred a t  lower values of l i f t  coeff ic ient .  
Increas ing  the  tes t  Reynolds number delayed somewhat the onset  of sepa- 
ra t ion  e f fec ts .  
INTRODUCTION 
As  pa r t  of a generalized study of t he  s t a l l i ng  o f  sweptback wings, 
a systerzatic investigation i s  being conducted t o  provide data on t h e  
e f f ec t s  of leading-edge radius on the  longi tudina l  s tab i l i ty  charac te r -  
i s t i c s  of a se r i e s  of wings with varying thickness ratio,  thickness dis- 
t r ibut ion,   cmher ,   aspect   ra t io ,  and sweep angle.  (See refs. 1, 2, 
and 3 . )  The present paper reports the results of tes ts  of a 4 5 O  sweptback 
wing with an aspect ratio of 3, and 6-percent- thick airfoi l  sect ions.  
The a i r f o i l   s e c t i o n  w a s  modified systematically t o  s tudy  the  e f fec ts  of 
leading-edge shape on the  s t a t i c  l ong i tud ina l  stability charac te r i s t ics .  
Tests were conducted a t  atmospheric pressure .and 'a t  33 pounds per square 
inch absolute pressure, which permitted Reynolds nurriber ranges from 
2 NACA RM L76B14 
1 . 3  x 10 to 4.1 x 10 (Mach  number from 0.06 to 0.20) and from 
2.3 x lo6 to 6.6 X lo6 (Mach  numbers  from 0.03 to 0.15), respectively. 
6 6 
SYMBOLS 
inflection  lift  coefficient  with  stable  ehange  in  pitching 
dcm moment, (- dCI; 
inflection  lift  coefficient  with  unstable  change  in  pitching 
maxFmum  lift  coefficient 
pitching-moment  coefficient  about 0.235, Pitching  moment/qSE 
wing  area,  sq ft 
mean  aerodynamic  chord, - Jb'*c2dy, ft 
s o  
density  of  air,  slugs/cu  ft 
free-stream  velocity,  ft/sec 
Reynolds  nmiber 
Mach  number 
3 
Ac/4 
A aspect   ra t io ,  b*/s 
sweepback of quarter-chord line, deg 
X chordwise  coordinate  of  airfoil  section, f t  
z vert ical   coordinate  of a i r fo i l   s ec t ion ,  f t  
MODEL 
The wings tes ted  dur ing  th i s  inves t i  a t ion  had the following geomet- 
r i c  cha rac t e r i s t i c s :  a spec t  r a t io ,  3; 45g sweepback of the  quarter-chord 
l i ne ;  t ape r  r a t io  of 0.5; and a maximum th ickness  ra t io  of 6 percent chord. 
(See  f ig.  1.) Three a i r fo i l  s ec t ions  were tes ted :  NACA 0006-(4.76)3, 
NACA 0006-(6.74)3, and a modified NACA 0006-(3.88)3. (Coordinates f o r  t h e  
three  a i r fo i l  sections are l is ted i n  f i g .  2.)  The f irst  two sections 
were formed by modifying the standard NACA four -d ig i t   a i r fo i l   by   the   p ro-  
cedure outlined in reference 4 i n  such a manner that the leading-edge 
radii were 0.25  percent c  and  0.50 percent  c,  respectively. For t he  
third ai r foi l ,  the  s tandard procedure w a s  used t o  form the  NACA 0006-(3.88j3 
a i r f o i l  sect ion from 0.4 percent of the chord aft. The forward 0.4 per- 
cent w a s  formed by fa i r ing  the  a i r fo i l  wi th  the  lead ing-edge  rad ius  of 0.25 
percent  c,  instead of t h e  normal  0.167 percent c f o r   t h a t   a i r f o i l .  
The r e s u l t i n g   a i r f o i l  w i l l  be designated herein as an NACA 0006-(3.88)3 
modif ied airfoi l  sect ion.  
Wing %/4 Aspect r a t i o  
Taper 
r a t i o  
1 
0.5 3 45 3 
0.5 3 45 2 
0 - 5  3 45 
Leading- 
edge 
radius 
0 .0025~ 
0.0050~ 
0.0025~ 
Air fo i l  sec t ion  
NACA 0006- ( 4.76) 3 
NACA 0006- ( 6.74) 3 
NACA 0006-(3.88)3 (mod. ) 
TESTS AND CORRECTIONS 
Tests 
Tests were conducted at a tunnel pressure of 33 pounds per square 
inch absolute  for  all three  wings and at atmospheric pressure for the 
wings with NACA 0006-(6.74)3 and NACA 0006-(3.88)3 modif ied   a i r fo i l  
4 - NACA RM L56B14 
sections. The following Mach nuniber and Reynolds nmiber var ia t ions 
resul ted as the  models were tested through a large part  of the tunnel 
speed  range : 
Tunnel pressure, 
Mach nmfber Reynolds number lb/sq in.  abs 
Variation of Variation of 
14.7 0.064 1.3 x lo6 
(atmospheric) 0.20 4.1 x 106 
t o   t o  
2.3 x lo6 
t o  33 
0.05 
0.15 6.6 x 106 
to 
The model was supported on the  normal two-support system of the  
Langley 19-foot pressure tunnel. 
Corrections 
The pitching-moment data  and values of angle of a t tack  have been 
corrected for tunnel-wall  effects by the  method of reference 5 .  These 
corrections are 
Since primary interest  in these data was centered on the  var ia t ion  of 
the l i f t  and pitching moments rather than on their  absolute  values ,  tes ts  
to  detemdne the model-support t a r e  and interference effects  were not 
conducted. However, the  zero- l i f t  pitching-moment coeff ic ient  and the  
angle of a t tack at  zero l i f t  on a wing with no caaiber, twist, or dihedral 
were taken as an indicat ion of the  combined e f f ec t s  of model-support t a r e  
and interference, air-stream misalinement, and model asymmetry. These 
corrections were assumed t o   b e  independent of l i f t  c o e f f i c i e n t ,  and the  
l i f t  and pitching-moment curves were shifted by the average value of the  
zero- l i f t  p i tching moment and angle of a t tack as  a f i r s t -order  
approximat ion. 
RESULTS AND DISCUSSION 
The variations of lift and pitching-moment coefficient with angle 
of attack are presented in figures 3 t o  5 and the variations of pitching- 
xonent coefficient with l i f t   a r e  presented in figures 6 t o  8. 
The ef fec ts  of var ia t ion   in   a i r fo i l   sec t ion  on the  pitching-moment 
character is t ics  of the wings are shown in  f igure  9, whereas the  e f fec ts  
of changes i n  Reynolds nuTdber a re  i l lus t ra ted  in  f igure  10. Mach  number 
e f f ec t s  fo r  two of the wings are shown in  f igure  ll. The variations 
of c 
figures 12, 13, and 1 4  for  the  wings with the various airfoil sections. 
L in fuJ  h n f S  > and C& 
with Reynolds number are shown i n  
The quantit ies C and C are used to designate the lift 
L i d u  Linf 
coefficient beyond which there i s  a marked change in   t he  pitching-moment 
characterist ics,  C denoting an unstable change and C 
Linf ,  L i d ,  
denoting a s table  change. In   f igure 13, two values of C are shown 
Linf, 
f o r  the data obtained at 33 pounds per square inch absolute, since the 
force data show the pitching moment t o  have two successive stable shifts. 
Changing the nose shape of the 6-percent-thick adrfoil had consid- 
erable effect  on the pitching-moment character is t ics  of the wings. (See 
f ig .  9.) As the  a i r fo i l  sec t ions  of the  wing  were thinned at the nose, 
the separation effects occurred at  lower values of l i f t  coefficient, as 
evidenced by the lower values of . A comparison of figures 12 CLinfU 
and 1 4  shows that, the  values of C f o r   t he  two  wings with  leading- 
%nfS 
edge radius of 0 .0025~ are  almost identic& through the Reynolds number 
range. The unstable  break C i s  not similar for  the  two wings, i n  
Linfu 
tha t   the   e f fec t  of thinning the section behind the nose was a lessening 
of the Reynolds number effects, thereby lowering the value to a marked 
degree at  the higher Reynolds numbers. 
Increasing the Reynolds nmiber at a constant Mach  number generally 
increased  both C and C fo r   t he  wing with an NACA 0006-(6.74)3 
airfoi l  sect ion.  (See f i g .  lO(a).) For the  wing with NACA 0006-(3.88)3 
modified a i r fo i l  sec t ions ,  the  e f fec t  of Reynolds number was  somewhat 
less aarked for the unstable break than for C . (See f ig .  10(b).) 
L i d u   L i d s  
Linf, 
" 
The ef fec ts  of an independent change i n  Mach number a re  shown i n  
f i g u r e s  l l ( a )  and l l ( b ) .  L i t t l e  e f f e c t  i s  noted fo r  e i the r  wing. The 
6 NACA RM L56B14 
most s ign i f icant  change at the higher Mach nmibers i s  the elimination of 
t he  s l i gh t  s t ab le  shift in  the  p i tch ing  moment which occurred at low 
l i f t  coef f ic ien ts  for  the  wing with the lazger radius.  (See figs 11 
and 13. ) 
The e f f e c t s  of leading-edge radius and Reynolds number on the  max- 
imum l i f t  coeff ic ient  can be seen t o   b e  small for  these  wings. (See 
f i g s .  12, 13, and 14.) 
Langley Aeronautical Laboratory, 
National Advisory Cormnittee for Aeronautics, 
Langley Field,  Va., January 31, 1956. 
REFERENCES 
1. Foster, Gerald V., and  Schneider, Wiliam C.: Effects of Leading-Edge 
Radius on the Longitudinal Stabil i ty of Two 4 5 O  Sweptback Wings As 
Influenced by Reynolds Numbers Up t o  8.20 X loG and Mach Numbers Up 
t o  0.303. NACA RM L55F06, 1955. 
2. Foster, Gerald V . :  Effect  of Leading-Edge Radius on the Longitudinal 
S t ab i l i t y  of Two 45' Sweptback Wings Incorporating Le ding-Edge cam- 
ber  As Influenced by Reynolds Numbers Up t o  8.00 x 10 and Mach Num- 
bers  Up t o  0.290. NACA RM L55H04, 1955. 
2 
3. Schneider, Wiliam C.:  Effects  of Leading-Edge Radius on the  Longi- 
tud ina l  S tab i l i ty  Charac te r i s t ics  of Two 60° Sweptback Wings A t  
High Reynolds Numbers. NACA RM L55L30, 1956. 
4. Stack, John, and Von Doenhoff, Albert E.: Tests of 16 Related Air- 
f o i l s  a t  High Speeds. NACA Rep.  492, 1934. 
5. Sivel ls ,  James C., and Salmi, Rachel M.: Jet-Boundary Corrections 
f o r  Completc and Semispan Swept Wings i n  Closed Circular wind 
Tunnels. NACA TN 2454, 1951. 
1- 53.13 - i 
. . .  
," 23.61 
Figure 1.- Details of wings. A l l  dimensions in inches  unless  otherwise 
noted. 
NACA 0006-( -7613 1.e .r. = 0.0025~ 
N A C A  0006-(k.&)3  1.e.r. = 0.0050~ 
NACA 0006- ( j .  ) j  Mod. 1.8.r. = 0.0025C - - - 
_”” 
- 
- ”.- L_ ”-  
”----- 
leading-edge r a d i i  
0 I I I I I I I I 
0 .02 .04 
Ordinates of a i r f o i l s  
./4 
NACA a ir fo i :  
I z  
.0125 
.0250 
-0750 
.0500 
.loo0 
.1500 
.2000 
0 
.0081 
.011 
-0168 
.0201 
.0230 
,026 
.0300 
.0292 
.0271 
.023 6 
.0191 
0074 
.0136 
.OOOb 
.02d 
0 
.0103 
.O1 9 
.0220 
e02L.3 
.0y 
.02p 
.02 9 
.0300 
.0292 
,0271 
.0191 
,0236 
0 
,0066 
.0100 
..0190 
so151 
.0220 
..026 3 
.0286 
.0292 
,0300 
0271 
.@23 6 
.01g1 
.0136 
.OO& 
-0006 
i 
! 
Figure 2.- Coordinates of airfoil sections. 
LO 
.8 
.6 
.4 
CL 
.2 
0 
- 9  
.L 
-4 0 4  8 /2 /6 20 24 28 
0 
0 0  0 0 0 0 0 4  8 /2 I6 
0 0 n v h Q L 7  
a;; deg 
(a) CL against a. 
Figure 3 . -  Variation of lift  and  pitching-moment  coefficients  with  angle 
of  attack for a 4 5 O  sweptback  wing  with  aspect  ratio  of 3 and NACA 
0006- (4.76 ) 3  airfoil  sections.  Leading-edge  radius, 0.0025; tunnel 
pressure of 33 pounds  per  square  inch,  absolute. 
20 24 28 
10 
R Y 
NACA RM L56B14 
0 -.04 
-. 04 
-4 0 4 8 /2 /6 20 24 28 32 
E, 
(b) C, against a. 
Figure 3. -  Concluded. 
1.0 
.8 
.6 
.4 
I CL .2 0 
-. 2 
-4 0 4 8 /2 1'6 20 24 28 32 
n u o  0 0 0  0 0 4 8 /2 /6 20 24 28 
0 0 n v 0 d 
(a,  CL  against a; tunnel  pressure,  atmospheric. 
Figure 4.- Variation  of  lift  and  pitching-moment  coefficients  with  angle 
of  attack  of  a 45' sweptback  wing  with  aspect  ratio  of 3 and NACA 
0006- (6.74 ) 3  airfoil  sections.  Leading-edge  radius, 0.0050. 
.8 
.6 
0 
-. 2 - 4 0 4  8 /2 I6 20 24 28 32 
u o o o  0 0 0 0  0 4  8 /2 /6 20 24 
0 0 n v [1 A D Q 
G de4 
(b ) CL  against a; tunnel  pressure, 33 pounds per square inch,  absolute. 
Figure 4. - Continued. 
R Y 
00 
0 1.36~14 0.065 
.0b0 
-093 n 2.32 .112 
8 :kg 
.04 
0 0  
-. 04 
-. 08 
-. 04 
- 4  0 4 8 /2 16 20 24 28 32 
dep 
(C) C, against a; tunnel  pressure,  atmospheric. 
Figure 4.- Continued. 
NACA RM L56B14 
.04 
0 0 -.04 
0 0 -.08 
-. 04 
-4 0 4 8 /2 /6 20  428 32 
(d) C, against a; tunnel  pressure, 33 pounds  per  square  inch,  absolute. 
Figme 4.- Concluded. 
I. 0 
.8 
.6 
-4 
CL 
.2 
0 
-.2 
-4 0 4 8 f2 16 20 24 28 32 
0 0 0 0 0 0 0  4 8 12 It5 20 24 28 
0 0 n v h A 
G dep' 
(a I CL against a; tunnel  pressure,  atmspheric. 
Figure 5.- Variation  of  lift  and  pitching-moment  coefficients  with  angle 
of  attack  of a 45' sweptback  wing  with  aspect  ratio  of 3 and NACA 
0006- (3.88)3 modified  airfoil  sections.  Leading-edge  radius, 0.0025. 
LO 
.8 
.6' 
.4 
CL 
.2 
0 
-. 2 
-4 0 4 8 12 I6 20 24 28 32 
0 0 0 0 0 0 0  0 4  8 I2 I6 20 2'4 
0 0 n v 0 A 17 
a, deg 
(b) CL against a; tunnel  pressure, 33 pounds per  square inch, absolute. 
Figure 5.- Continued. 
R M 
0 
0 
0 
1.6i 1.3 0.064 
1 .90 -079 
n 2.32 .112 .090 
V 2-69 .131 
2.97 
A 4.05 -137 *193 
-04 
0 0  
CI 0 -.04 
0 0 -.08 
A 0  
cm 
0 0  
0 0  
0 0  
-.04 
( C  ) C, against a; tunnel pressure,  atmspheric. 
Figure 5.- Continued. 
I I1 1111 I I I 
0 0  
00 
NACA RM L56B14 
R Y 
.u4 
0 0  
-.04 
-.08 
-. 04 
( a>  Cm against a; tunnel  pressure, 33 pounds  per  square  inch,  absolute. 
Figure 5. - Concluded. 
LO 
.8 
.6 
I 4 CL .2 
0 
-. 2 
0 -.04 
I7 
Figure 6.- Variation  of  pitching-moment  coefficient  with  lift  coefficient 
of a 45' sweptback  wing  with  aspect ra io of 3 and NACA 0006-(4.76)3 
airfoil  sections.  Leading-edge  radius, 0.0023; tunnel  pressure  of 
33 pounds  per square inch,  absolute. 
1.0 
.8 
.6 
.4 
.2 
0 
-. 2 
(a)  Tunnel  pressure,  atmospheric. 
Q 
? n 
V 
0 
A 
,04 0 -,04 -.08 
0 0 0 0 0 0 0 - . 0 4  
13 0 n v L 11 
cm 
R Y 
1.36~106 0.0 5 
1.65 00 Q 0 
1.93 093 
2.32 8112 
2 069 
-1 7 
8201 t ;:; “t’ 
Figure 7.- Variation  of  pitching-moment  coefficient  with  lift  coefficient 
of a 43’ sweptback  wing  with  aspect  ratio  of 3 and NACA 0006-(6.74 ) 3  
airfoil  sections.  Leading-edge  radius, 0.0050. 
10 
.8 
.6 
.4 
CL 
.2 
0 
-. 2 
.04 0 -.04 -.OB 
0 0 0 0 0 0 0 0 0 
c, 0 0 n v h Ll U 0 
(b)  Tunnel  pressure, 33 pounds  per  square inch, absolute. 
Figure 7.- Concluded. 
R M 
o 2.32 x106 0.050 
02 064 
.087 
.092 
b 5.5 .122 
A 
CI 6.64 
5.08 .io8 
a 5’:7oa 
I. 0 
.8 
.6 
.4 
CL 
1 .2 0 
-. 2 
.04 0 -.04 
n 
W 0 0 0 0 0 0 -.04 
0 0 n G7 [1 Ll 
cm 
(a ) Tunnel  pres sure, atmospheric. 
Figure 8.- Variation of pitching-moment  coefficient  with  lift  coefficient 
of  a 45' sweptback  wing  with  aspect  ratio of 3 and NACA 0006-(3.88 ) 3  
modified  airfoil  sections.  Leading-edge  radius, 0.0025. 
I. 0 
.8 
.6 
.4 
CL 
.2 
0 
-. 2 
.m 0 704 w.08 
0 0 0 0 0 0 0 0 -.04 
0 0 n v @. d I7 
cm 
(b) Tunnel  pressure, 33 pounds  per  square  inch, absolute. 
Figure 8.- Concluded. 
t 
Iu w 
A i r f o i l  section 
NACA 0006-(4.76)3 1.e .r. = 0 . 0 0 2 5 ~  
- " - NACA 0006- ( 6  -74 ) 3 1 .e .r. = 0.0050~ 
- "" NACA 0006-(3.88)3 Mod. l . e . r .  = 0 . 0 0 2 5 ~  
f.0 
.8 
.6 
CL 
.4 
I 
I 
2 
0 
, #08 .04 0 
(a)  Tunnel  pressure, 33 pounds  per  square  inch,  absolute. 
0 -.04 ~ 0 8  
Figure 9.- Effect of changes  in  nose  shape  on  pitching-moment  character- 
istics of a 45O sweptback  wing  of  aspect  ratio 3 .  
A i r f o i l  sect ion 
NACA 0006- ( 6 .  1.e.r. = 0 . 0 0 5 0 ~  
-" NACA 0 0 0 6 - ( 3 .  Mod. 1.e.r. = 0 . 0 0 2 5 ~  
l.0 
-8 
.6 
.2 
0 
.08 - 0 4  0 0 
cm 
(b ) Tunnel  pressure,  atmospheric. 
Figure 9. - Concluded. 
0 -.04 -.08 
M = 0.06 M = 0.11 M = 0.15 
l.0 
.8 
.6 
.4 
.2 
0 
.08 .04 0 0 0 -.04 -.08 
( a )  Airfoil section, NACA 0006- (6.74 ) 3 ;  leading-edge radius, 0.0050~. 
Figure 10.- Effect of  Reynolds number changes on pitching-moment charac- 
t e r i s t i c s  of a 4 5 O  sweptback wing of aspect  ra t io  3 .  
H 
1.0 
.8 
-6 
.4 
.2 
0 
M = 0.06 M = 0.11 M = 0.15 
.08 .04 0 0 0 704 -.08 
(b ) Airfoil  section, NACA 0006- (3.88 )3  modified;  leading-edge 
radius, 0.0025~.  
Figure 10.- Concluded. 
.8 
.6 
I 
CL 
.4 
2 
0 
R = 2.3 x 10 6 R = 3.0 X 106 R = 4.1 x 10 6 
.08 .04 0 0 0 -.04 -.08 
(a)  Airfoil  section,  NACA 0006-(6.74)3; leading-edge  radius, 0.0050~. 
Figure 11.- Effect of Mach  number  changes on pitching-moment  characteris- 
tics of a 4 5 O  sweptback  wing  of  aspect  ratio 3 .  
I 
R = 2.3 x 10 R = 4.1 x 10 R = 3.0 X 10 6 6 6 
I 
1.0 
.8 
.6 
c, 
.4 
.2 
0 
.08 .04 0 0 0 704 "08 
(b) Airfoil  section,  NACA 0006-(3.88 ) 3  modified;  leading-edge 
radius, O.OO25c. 
Figure 11.- Concluded. 
NACA .RM L56Bl4 
0 Tunnel Pressure atmospheric. 
0 Tunnel pressure 33 pounds per square inch, absolu te .  
1.0 
.8 
.6 
.4 
.2 
0 
f 
/ 2 3 4 5 6 8 /OX/06 
R 
Figure 12.- Variation  of  inflection  lift  coefficients  and  maximum  lift 
coefficients  with  Reynolds  number  for  a 45’ sweptback  wing  of  aspect 
ratio 3 with NACA 0006- (4.76 )3  airfoil  sections.  Leading-edge 
radius, 0.0025~. 
0 Tunnel pressure atmospheric. 
Tunnel pressure 33  pounds per square inch, absolu te .  
-8 
.6 
CL 
.4 
.2 
0 
I 
Figure 13.-  Variation  of  inflection  lift  coefficients  and  maximum  lift 
coefficients  with  Reynolds  number for a 45O sweptback  wing of aspect 
ratio 3 with NACA 0006-(6.74)3 airfoil  sections.  Leading-edge 
radius,  O.OO5Oc. 
NACA RM L56B14 
0 Tunnel pressure atmospheric 
Tunnel pressure 33 pounds per square inch, absolute. 
1.0 
-8 
.6 
GL 
.4 
.2 
0 
/ 2 3 4 5 6 8 /Ox /06  
R 
Figure 14.- Variation of inflection  lift  coefficients  and  maximum  lift 
coefficients  with  Reynolds  number for a 45’ sweptback  wing  of  aspect 
ratio 3 with NACA 0006-(3.88)3 mdified airfoil  sections.  Leading- 
edge  radius, 0.0025~.  
”- 
NACA - Langley Field, Va. 


Effects of Reynolds number and leading-edge shape on the low-speed longitudinal stability of a 6-percent-thick 45 degree sweptback wing

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