Effect of compressibility on the pressures and forces acting on a modified NACA 65,3-019 airfoil having a 0.20-chord flap

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Title:
Effect of compressibility on the pressures and forces acting on a modified NACA 65,3-019 airfoil having a 0.20-chord flap
Alternate Title:
NACA wartime reports
Physical Description:
12, 69 p. : ; 28 cm.
Language:
English
Creator:
Lindsey, W. 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:
Aerofoils   ( lcsh )
Compressibility   ( lcsh )
Aerodynamics -- Research   ( lcsh )
Genre:
federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Summary:
Summary: An investigation has been conducted in the Langley rectangular high-speed tunnel to determine the effect of compressibility on the pressure distribution for a modified NACA 65,3-019 airfoil having a 0.20-chord flap. The investigation was made for an angle-of-attack range extending from -2° to 12° at flap deflections from 0° to -12°. Test data were obtained for Mach numbers from 0.28 to approximately 0.74. The results show that the effectiveness of the trailing-edge-type control surface rapidly decreased and approached zero as the Mach number increased above critical value.
Bibliography:
Includes bibliographic references (p. 11).
Statement of Responsibility:
by W.F. Lindsey.
General Note:
"Report no. L-76."
General Note:
"Originally issued January 1946 as Advance Confidential Report L5G31a."
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 - 003638549
oclc - 71835200
sobekcm - AA00006244_00001
System ID:
AA00006244:00001

Full Text

(AcA L-7&


ACE No. L5G31a


NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS


WARTIME REPORT
ORIGINALLY ISSUED
January 1946 as
Advance Confidential Report L5G31a

EFFECT OF COMPERESSBIITY ON THE PRESSURES AND FORCES
ACTING ON A MODIFIED ACA 65,3-019 AIRFOIL
HAVING A 0.20-CHORD FLAP
By W. F. Lindsey


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.


DOCUMENTS DEPARTMENT


/


L 76






































Digitized by the Internet Archive
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/effectofcompress001a







ITACA ACR No. L5G31a


NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS


ADVANCE CONFIDENTIAL REPORT

EFFECT OF COMPRESSIBILITY ON THE PRESSURES AND FORCES

HCTI:!G ON A 'ODIFIZD NACA 65,5-"019 jTRFOIL

HAVING A 0.20-CHORD FLAP

By W. F. Lindsey


SUMMARY


An investigation has been conducted in the Langley
rectangular high-speed tunnel to determine the effect of
comnressibility on the pressure distribution for a modi-
fied NACA 65,5-019 airfoil having a 0.20-chord flap. The
investigation was made for an angle-of-attack range
extending from -20 to 12 at flap deflections from 00
to -120. Test data were obtained for Mach numbers
from 0.28 to approximately 0.74.

The results show that the effectiveness of the
trailing-edge-type control surface rapidly decreased and
approached zero as the Mach number increased above the
critical value.


INTRODUCTION


The available information on the aerodynamic charac-
teristics of airfoils, with and without flaps, at low
speeds is quite extensive and previous investigations
have shown the general effects of compressibility on air-
foils without flaps. For airfoils with flaps, however,
the available information at high speeds is limited.

The earlier investigations at high speeds demon-
strated the limitations of the theoretical methods in
extrapolating low-speed data to high speeds (reference 1).
In addition, the investigations illustrated the inappli-
cability of the theoretical methods in the supercritical-
speed range, in which pressures and forces change radi-
cally. These radical changes in pressures and forces are
the adverse effects of compressibility, a knowledge of








NACA ACR No. L5G51a


which is necessary in the design of high-speed airplanes.
Because of the inadequacy of the theoretical method in
applications to the supercritical region, recourse to
exr-c-riment is necessary to determine the adverse effects
of compressibility.

Data at high speeds were required in connection with
the design of a specific airplane; accordingly, an inves-
tigation was conducted in the Langley rectangular high-
speed tunnel to determine the effect of compressibility
on the noessures and forces acting on a modified NIACA 65,3-019
airfoil having a 0.20-chord flap. Pressure-distribution
measurements were made at Mach numbers between 0.26
and 0.74 for angles of attack from -20 to 120 and flap
deflections from 00 to -120.


APP.RRJ-TUS .rD TESTS


The tests were conducted in the Langley rectangular
high-speed tunnel, which is an induction-type tunnel
without return passages and has an 18-inch by 4-inch test
section. The variation in the Iach number in the test
section along the tunnel axis without a model installed
in the tunnel is 0.4 percent of the stream Mach number.
In a plane normal to ~-e tunnel axis the variation is
t0.8 percent of the test-section Yi.zh number at a stream
Mach number of 0.60. The direction of the air flow
appears to be misalined by -0.10 with a possible variation
of 0.10. I1o correction for misalinement has been made
for the data presented herein.

The model comelately spanned the test section along
the ,.-inch dimension and was supported by large circular
end plates which were fitted into the tunnel walls in
such a way as to rotate with the miodtkl and to retain
continuity of the surface of the tunnel walls. The
juncture between the model and the end plate was sealed.

The profile of the 5-inch-chord nodel having a
0.20-chord flap differed from the modified IACA 65,3-019
airfoil section in that the profile from approximately
the 81-percent station to the trailing edge was fonned
by straight lines having an included angle between 200
and 210. r'orty pressure orifices were installed in the
modol surface in tw.o chordwise rows inch from and on
4.


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NACA ACR No. L5G31a


either side of the model center line. The model profile
and pressure-orifice locations are shown in figure 1.
The airfoil ordinates are given in table I.

Pressure-distribution measurements were made for a
range of Mach numbers from 0.28 to approximately 0.7L at
angles of attack from -20 to 120 and flap deflections
from 00 to -120 (up). Additional pressure-distribution
measurements were made for positive (down) deflections of
the flap at angles of attack of 20 and 40. These tests
were supplemented by sohlieren photographs of the flow in
the supercritical speed region for a few of the lower
angle-of-attack configurations. These photographs show
density gradients in the flow by changes in light
intensity. (For details, see reference 1.)


SYMBOLS


M Mach number

Mcr Mach number at which sonic velocity was obtained
locally within the flow field (as at the model
surface)

q dynamic pressure

Po free-stream static pressure

p local static pressure (as at model surface)

a angle of attack

P pressure coefficient Po1

Pcr critical pressure coefficient, corresponding to
0.528 spo
local Mach number of 1. 0(-5

H total pressure

PU upper-balance-chamber pressure coefficient

PL lower-balance-chamber pressure coefficient

Cn section normal-force coefficient


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NACA ACR No. L5G31a


Cmrc/ section pitching-moment coefficient of normal
force about quarter-chord location

Cnf flap normal-force coefficient

ch flap hinge-moment coefficient (determined by the
pressure distribution from flap hinge axis to
trailing edge)

6 flap deflection measured with reference to the
model chord; negative deflection is up


RESULTS AND DISCUSSION

Tunnel-Wall Effects


The results of these tests have not been corrected
for constriction or tunnel-wall effects. The most
important constriction effect on these data in the super-
critical region is the change in the I.:ach numbers given
herein to higher effective stream Mach numbers. (See
reference 2.) The difference between the two Mach numbers
increases rapidly as the maximum tunnel speed is approached.
It is further shown in reference 2 that very near or at
the maximum speed attainable for a given mndel-tunnel
configuration large gradients in velocity occurred at the
walls. The maximum-speed test -oints given herein for
each angular confipu.ration are therefore considered to be
of questionable value.


Pressure Distribution

The pressure distributions albng the chord of the
model are presented in figures 2 to $1, inclusive. Each
figure shovs the effect of c:mrpressibility on the pressure
distributionn for a -riven anullr configuration. The
effect of angle of attack and flao deflection can be
obtained from a c-::nparison of the various figures.

3ubcritical region.- The figures for the lower angles
of attack and 3nmall normal-force coefficients show that
increases in M"ach number in the subcritical ranges are
acco,)panied by increases in the maximum negative pressure
coefficient, s'hich is in agreement with theory. In the
high angle-cf-attack and large normal-force-coefficient


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NACA ACR No. L5G31a


range the change in the maximum negative pressure coef-
ficient with Mach number is approximately zero and not in
agreement with theory, probably as a result of the exist-
ence of separated flow.

A comparison of part (a) of figures 2 to 51 for a
constant angle of attack and various flap deflections
shows that the increment in load produced by a deflection
of the flap is distributed approximately uniformly along
the chord as could be expected from low-speed tests. It
can be seen, however, at this M.ach number (approx. 0.45)
that the maximr.ul relative change in loading on the main
part of the wing occurs near the leading edge. This
change has an appreciable effect in increasing or relieving
the pressure peaks that occur near the leading edge for
some angle-of-attack conditions. (See part (a) of figs. 2
to 5 and 18 to 21.)

Supercritical region.- Although it has been shown
that in tne suocr- tical Mach number range the action of
the flap in changing the loading along the chord was
similar to that shown ty the low-speed Lests, in the
supercritical range, when the region of supersonic flow
is relatively large, the loading in and ahead of the
supersonic region is a function only of angle of attack.
The extent of flow affected by the flap is limited to the
region of subsonic flow behind the supersonic region and
to the flow over the flap itself. (See parts (d) and (e)
of figs. 6 to 9.) The chordwise influence of flap deflec-
tion on the flow over the main part of the wing could be
expected to be limited to the subsonic flow region ahead
of the flap, since pressures are propagated at the speed
of sound. This effect of compressibility in producing a
marked change in the flow over the forward portion of the
airfoil at supercritical Mach numbers is comparable to
that which occurs for cambered airfoils, as evidenced by
the change in the angle of zero lift. The similarity can
be more clearly seen when the cambered airfoil is con-
sidered to be a multiple flapped airfoil.

Comparison of the pressure distributions for the
flap deflected and neutral shows that, for this model,
the chordwise extent of the influence of the flap
decreases gradually as the speed is increased above the
critical speed.

If the fundamental aspects of the flow are con-
sidered, deflections of the flap could produce changes


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NACA ACR No. L5G31a


in the pressures in the subsonic flow region, thereby
chirr,.-nwi the shock location. This effect, however, is
not al',arent in these results because of separated flow
and the resulting absence of the usual discontinuities in
the pressure-distribution diagrams which indicate the
shock location.


Schlieren Photograchs

3chlieren photographs of the flow in the higher Mach
number range are shown in figures 52 to 37, inclusive,
for the model at angles of attack of 0 and Oo with
various flap deflections. (Note co:rpressions are white
on fig. 32, black on figs. 35 to 37.) These photographs
show that for angles of attack from 00 to 4 the flow
separates at approx.nimtely the 0.60-chord location for
Mach numbers near the critical values, and with increasing
Mach number the separation point moves forward. The
forward movement of the separation point is accompanied
by a rearward movement of the shock along the sdoaration
boundary.

A comparison of the schlieren photographs with the
corresponding pressure-distribution diagrams shows that
the existence of separated flow has a serious effect on
the pressure distribution in the vicinity of compression
shocks. (See, for example, figs. 9(d) and 55(d).) It
will be noted that for the condition of a well-defined
shock and separated flow the pressure-distribution diagram
indicates smooth compression. In addition, the location
of the pressure corresponding to the critical pressure
coefficient generally occurs from 5 to 10 percent of the
chord downstream from the shock location. This phenomenon
is probably the result of the existence of large static
pressure gradients in the separated-flow reblon between
the boundary and the model surface where pressures were
measured. The extent of the serrated flow would be
reduced for an airfoil havlrin a smaller thickness-to-
chord ratio.


Critical MHch ?Tumiber

The variation of the critical Mach number of each
surface with flap deflection for constant angles of attack
is presented in figure 38. The ltrge decrease in critical
' ch number that occurs at angles of attack between 6


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NACA ACR :o. L351 la CO'IFIDENTIAL


and 8 for the upper surface is a result of a rapid
increase in the magnitude of the negative pressure coef-
ficients near the leading edge. The difference between
the critical Nach numbers for the upper and lower surfaces
for corresponding conditions can be attributed to air-
flow misalinement. The highest critical Mach number for
this airfoil is approximately 0.65 and is obtained for
angular configurations corresponding to a normal force of
approximately zero.


Compressibility Effects on Force and Moment Coefficients

The normal-force, moment, flap normal-force, and
hinge-moment coefficients obtained from integration of
pressure-distribution diagrams are presented in figures 39
and 40. Each figure shows the variation in the coeffi-
cient with Mach number at a constant angle of attack for
each flap deflection. These figures are cross-plotted in
figure 41 to show the variation in Lhe coefficients at a
constant Mach number.

Normal-force coefficients.-Figure 79 shows that, with
the flan at 00, the effect of compressibility on the
normal-force coefficient at subcritical i'.ach numbers is
in accord with previous experimental and theoretical
results. The variation for the other flap deflections at
a fixed angle of attack, however, appears to follow, to
some extent, both in direction and magnitude the variation
for the condition of the flap at 0, as indicated by a
lesser divergence of the curves than would have been
expected from theoretical estimations. (See parts (c),
(d), and (e), fig. 59.) A variation of this type shows
dcn
that the effect of compressibility on n is small in
d6
the subcritical Mach number range.

In the supercritical Mach number range above the
value at which the peak normal-force coefficient occurs,
the convergence of the curves for the various flap
deflections at a given angle of attack indicates a rapidly
dcn
decreasing with increasing Mach number. The effect
d5
dcn
of compressibility on -, presented in figure 42, is
66
in accord with the variations indicated in figure 59.


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INACA ACR :io. L5G31a


In the sunercritical !ach ran-e for arngles of attack
not rreater than 60 it can be seen in figure 39 that the
oeak normal-force coefficient for a given angular configu-
ration occurred at a 1..ach number of approximately 0.65,
wh-ich indicates that this airfoil section should be
restricted to designs v.herein the maximum I.:ach number
does not exceed 0.65.

A co.:narison of the various parts of figure 41 in
which is 6iven the variation of the normal-force coeffi-
cient with angle of attack shows that as the Miach number
dcn
increases increases and reaches a mraxi.nu value at
da
a Mach number of 0.65 as could have been expected from
the preceding discussion.

A further ex.'ination of fi,.ure l1 sho.vs that
deflections of the flap produce a change in the angle of
den
zero normal force, but have no a:rnreciable effect on -d
da
for the more linear arts of the curves. The effect of
dc
coo-i ressibility on n for the more linear parts of
doL
these curves is presented in figure 42.
da
The flap effectiveness -, obtained from the ratio
c d ,-,
of -n to (fig. .2) and )r:escnted in figure 35,
d6 da
decreased with increased Mlach number and rapidly approached
zero as the Mach number was increased above the critical
value.

',r.ent coefficients.- The variation of the mr.o'ent
coefficient .ith .:achi, number in the subcritical range as
shown in figure 39 is small. In the s,-percri.. cal region
the moment coefficients generally increase with increasing
Mach number. This increase is followed by a rapi.i
decrease, which occurs at a Mach number above the value
at which the decrease in normal-force coefficient occurs.
The nrrcssure-distribution diagrams show that the decrease
in nornal-force results fr.-m a general decrease in the
rmagnitu*.ie of the loading, anwl at hih.-r .:ach numbers the
continued diecr.ase in loa'in.g is accor.ptanied by a change
in distribution which results in a -Jecreased moment coef-
ficient.


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NACA ACR No. L5G31a


At the supercritical Mach numbers at which the normal
forces for the various flap deflections tend to be approxi-
mately the same, the direction of the change in the moment
coefficieints is similar to the direction of the change
in On. A variation of similar magnitude, however, could
not be expected in these coefficients because of the large
effect that a small change in load at the rear of the model
has on the moment coefficient.

A comparison of the effects of angle of attack and
flap deflections on the variation in moment coefficient
with normal-force coefficient can be seen in figure 41.

For constant flap deflection there is a small positive
increase in moment coefficient with an increase in normal-
force coefficient. This slope remains approximately con-
stant for Mach numbers to approximately 0.65. For Mach
numbers above 0.65, the variation depends on the flap
deflection.

At a constant angle of attack between -20 and 60,
large changes in moment coefficient occur in a negative
direction with an increase in the normal-force coeffi-
cient. These slopes, which are approximately constant to
a Mach number of 0.6, increase as the 1ach number is
further increased to 0.7.

Flap normal force.- The flao normal-force coeffi-
cients are presented in figure L0. In the supercritical
region, the variations in the coefficients are large and
irregular, probably being influenced by the effect of flow
separation.

Hinge-moment coefficients.- The variation in the
hinge-moment coefficients with "ach number, also shown in
figure 0, are very similar t- the variation in flap
normal force.

At Mach numbers from 0.01 to 0.02 below the maximum
test value, a flap deflection range is indicated in which
the flap tends to become or is overbalanced. At the same
Mach number and for the same angular configuration, in
figure 39, no change occurs in the value of normal-force
coefficient, and the control would therefore be unresisting
and ineffective. Although the Miach number is near che
maximum test value, for which the data are of questionable
value, the possibility of this condition of the control
should not be overlooked.


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;NACA ACR N:. L5331a


The hinCe-n-ment coefficients are presented in
fiTui'e 41 in the sa-ne ,anner as the moment coefficients.
It can he seen that for a constant flap deflection the
'magnitude an.- direction of change of hinge-r.oment coeffi-
cient with increase in normial-force coefficient depends
on the absolute value of the flap deflection. The slope
for a given flap deflection increases positively with
increases in Mach number to 0.675; further increases in
.ach number are accompanied by increases in the slope in
the negative direction.

At a Mach number of 0.40 the changes in hinge moment
with flap deflection at a constant angle of attack are
generally uniform over the flap deflection range. At a
"ach number of 0.60 and above in the negative normal-
force-coefficient rsnge the changes in hinge-moment c.ef-
ficient for flzo deflections between 0 and -o0 is very
small cc.rpared with the changes at larger reflections.
This effect is pos.i'ly a result of the reversals in the
load over the flap produced by the very thick boundary
layer or separated flow. (See pressure-distribution
ria.rams and schlieren photogra-.-hs.)

Balance-chamber Dressure-coefficient differential
(L TP)j- Tfe difference oetv.een the pressure coeffi-
cients for the lower-surface and the upper-st'rface balance
chambers is .resented in figure +.. T'-is figure shows
the effect of ,ach number, flap deflection, and angle of
attack: on the pressure-coefficient differential.

The effect of cornressibilit.,- on the cressure-
coefficient differential is generally small at ;latch
numbers below 0.65. This small effect could be expected
when the .anui'tudes of the indivi.sual pressure coeffi-
cients are small and are measured at a station in rear of
the position of the maximum negative pressu-.e cc-fficient.
The decrease in the magnitude of the p'"e.tsu2e coff2icient
at the i.igh :.:ach numbers is roriir.aril a result of the
.-ffects of flow s-raration.

The balanced hin.ge-moment coefficient for this model
can be obtained by acdin:' the hinje-mor.ent o-ocfficient of
fi.:ur 0 to the product of the balance-chamiber prssure-
co.fficient differtntial (PL PTJ) and a constant, the
constant depincidng on the lun.-th of the balance tab. A
tricf cyr..parison of fi:;r,:s L' an3 L' indicates that at


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NACA ACR No. L5G1la


dch
lo, speeds the effect of the balance would be to reduce
d6
as compared with the unbalanced condition. With increasing
Mach number the effect of the balance decreases for the
lower angle-of-attack range.


CONCLUDING R E AIRK


The results of the investigation on the modified
NACA 65,5-019 airfoil having a 0.20-chord flap indicate
that the effectiveness of a trailing-edge control surface
of small chord rapidly decreases and approaches zero as
the I.ach number increases above the critical value.


Langley Merimn-ial Aeronautical Laboratory
National Advisory Commrittee for Aeronautics
Langley Field, Va.









.UTI






REFERENCES


1. Stack, John, Lindsey, vV. F,, and Littell, Robert E.:
The Compressibility Burble and the Effect of Com-
pressibility on Pressures and Forces Acting on an
Airfoil. NACA Rep. IT. 646, 1938.

2. Byrne, Robert W.: Experimental Constriction Effects
in High-Speed Wind-Tunnels. NACA ACR No. L4L07a,
19.


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NACA ACR No. L5G31a


TABLE I.- BASIC SECTION ORDI'INaTES FOR 1'CDIFIED

iiNCA 65,5-019 AIRFOIL SECTIDI


iSuations and


ordinates are in percent chord]


I Ordinate


Stati on


0
.5
1
2

8
10
io
18
22
26
50

ii
9.5
2,
)46
50
54
58
62
66
70
74.
78
82
86
90
94
98
100
Sr..E r diu111


Upper Lower
surface surface

0 0
1.108 -1.10i
1.921 -1.921
2.59, -2.598
.620 -6.620
-.457 .4537
5.127 -5.127
5.725 -5.725
6.715 -6.715
7.525 -7.525
El.192 -1.192
I 8.721 -8.721
9.113 -9.115
S 9.71 -9-571
9490 -9.490
9.o00 -,. 00
-Q.31
9. 71 -9.471
9.515 -9.515
9.021 -9.024
8.597 -8.597
r.059 .059
7.570 -7.570
b.612 -6 612
5.791 .791
2.922 -4.922

.123 -8.128
2.247 -2.247
1.416 -1.416
.63, -. 688
.139 -. 15
0 0
2_1 9 Q


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I ._


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NACA ACR No. L5G31a Fig. 1










Il k z*








'-4




1-1
0






.Z
.O

z ~
C. 'U







4L r




s
<-> ; o 's








NACA ACR No. L5G31a


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x Upper surface
0 Lower surface


/"& /= 0434.


fe 0. e727.


0 20 40 60 60


M f= 0. 6/.










cr




) = 0. 7703.















,'= O. 740.


/ 0 0 o d40 60 0 /00


Percent/ chord
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Figure 2.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a -20; 8 00.


MrTIOAL ADnSOm
CrNtmTTl MlM AIMAUTICS


Fig. 2a-f








NACA ACR No. L5G31a


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x UZoer sUr/face
0 Lower Surface


. --- Q





0 =0./L4.


0 20 40 60 80


/00 0 20 40 60 60 /00


PPrce n chord


IAT IONAL ADVISlORY
CONNITTLE FT0 ALROIATICS


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Figure 3.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a = -2; S = -4.


/=


Fig, 3a-f







NACA ACR No. L5G31a


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x Upper surface
0 Lower surface


(ci M=06 7/.


_ -- --





I /= 0 6/3.















W O. 70/.










(PX-



ft) =O. 745.


0 20 40 60 0 /00 0 2.0
Percent chord


40 60 50 /00

NATOGMAL AVTSbQY
aMMttEE FM mMIAWICL


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Figure 4.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a -2. S -80.


cr
K__t2-~i2


Fig. 4a-f


I








Fig. 5a-f


0 20 40 60 80 /00 0 20 40 60 60 /00


Percent chord


NATIONAL ADVISOIV
CONir'liL 101 AIONilUfTCS


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Figure 5.- Pressure distribution fur a modified NACA 65,3-019
airfoil with 0.20-chord flap. a = -20; S = -120


NACA ACR No. L5G31a


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x Upper sur/ace
0 L owner surface


[/c) "-O 674.


/


;)/ ;=O) C/0.

















(d) /1--0.700.







NACA ACR No. L5G31a


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x Upper surface
0 Lower surface


(o) /1=0442.


-/ b- 1i-/


Cr~


77)


(e) /1=0726.


P-


(0) /7=0646.










cr




() ---=0706.


O 20 40 60 0 /00 0 20 40 60 d6 /00

Percent chord

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Figure 6.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a 00; S = 00.


A1


Fig. 6a-f


~b~F~


S 7- 7 .








NACA ACR No. L5G31a


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x Lge,' sur/face
0 Lower Surface


(0) I= 0. 435.


cr





(c) M/= 0677.

















(Q) M= 0733.


(d) /= -0.709.
I7 /= 0.709.


0 20 40 60 80 /0 0 ZO 40 60 80


PGrce;n/ caord
CONFIDENTIAL


NATIONAL ADVISORY
CONNITTEE n AUgAWTCrKs


Figure 7.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0 .20-chord flap. a = 00; 8 = -40.


Fig. 7a-f








NACA ACR No. L5G31a


CONFIDENTIAL

x Upper surface
o Lower surface


eW/= O.4.'i7


(E) fc =0. 736.


0 20 40 6o d60 Ao7


0 20 40 60 Wi /O


Percent chor d

CONFIDENTIAL


Figure 8.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a 00; 8 -80


4=0.7


mATPOMAL AVISOAT
COWM7TTIU F0 MUTIKS


-C






M-O


Fig. 8a-f








NACA ACR No. L5G31a


CONFIDENTIAL
, Upper surface
0 Lower surface


K t1=0.648.


:-
cr


0 Zo 40 60 60


/Wo 0 20 40 60 80 ,O0


Percent chords
CONFIDENTIAL


NATIONAL ADVISORY
COmITTEE I JAEI OkUTICS


Figure 9.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a = 00; 8 = -12.


k I


S/V -0.6 76.


Fig. 9a-f








NACA ACR No. L5G31a


CONFIDENTIAL
x Upper surface
0 L wer surface


(a) /10.4.36.
















(c) n=0.678.


ab) t-fr0;2643 -


-/


0


/Z




Q.





Q0
'JO


0 ZO 40 60 80 0 00 20 40 60 80 /00
NA TIONAL. ADB 0Y6V
PorcQ enf chord coMITTU I itt
CONFIDENTIAL


Figure 10.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a a 2o; 8 0O


S--- 705.


Fig. lOa-f


/







NACA ACR No. L5G31a


Fig. lla-f


CONFIDENTIAL

0x Uoe,- su-face
0 lowe- surface


/ (a ) = O.4 J


















(c) M 0.675













-/



(e) '--0.75/


-r


0 ZO 40 60 80 /00 0


PeSrcenf cor-o


(do) = 0.7/0


20 40 60 80 /00


NATIONAL ADVISORY
OMITItL FNM AIEONATUCS.


CONFIDENTIAL


Figure 11.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a = 20; 8 = -4.


Aj


U







Fig. 12a-f


NACA ACR No. L5G31a


CONFIDENTIAL

x Upper surface
o Lower surface


-I-

0)



(c) /f7= 0675.


f---^-^














f(6) 0.7 07.


0 20 40 60 60 /00 0 20 40 60 80 i00


,-'erce n chord


nATTO r A &OVSOAY
COmnu F. MM AMIUUTrC&


CONFIDENTIAL


Figure 12.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a 20; 8 -80







S NACA ACR No. L5G31a


CONFIDENTIAL
x Uppe.P surface
D Lower surface


N


( 1) =0. 433.


































e)C -0.73 .
f --- --- --- *S C



/() =0.77.











/ =O.7


0 ZO 40 60 80 /00 0 20 40 60 60


Percent/ chord


NATIONAL ADVISORY
COMMITTEE fo A MAUTICS.


CONFIDENTIAL


FiBure 13.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a = 20; 8 = -120.


d) 0=-70-7.
















S/ 1=0. 742.


Fig. 13a-f








NACA ACR No. L5G31a


CONFIDENTIAL
x Upper surface
0 Lower surface


(a) M= 0433.
















(c) /1=0680.


(6) M=a061/4.
















0d) N=o .704-.


0 20 40 60 90 /00 0 20 40 60 S0 /OI


Percent chord


nATtOAma WDISOVY
coMMITMT Fa WIrTS


CONFIDENTIAL


Figure 14.- Pressure distribution for a modified NACA 65, -019
airfoil with 0.20-chord flap. a 40; 8 00.


-I


0


Q.



0c


Fig. 14a-f








'NACA ACR No. L5G31a


CONFIDENTIAL

x Upper su~i4ce
- Lopwer surface


-/


*; 0



":*:
/










A)
Is-
Q

c

u

a)
0
C^ n


(a) /=0.436.


-2










/ (e M=0.L725. (f) /A=0.739.

0 20 40 60 80 /00 0 20 40 00 80 /00

Percent chord NI. .s
CONMITTEE FOI MlAUIICS
CONFIDENTIAL

Figure 15.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a = 40; 8 = -4.


rpncn




) W r= 0.644.












rr


Fig. 15a-f







NACA ACR No. L5G31a


CONFIDENTIAL
x Upper surface
0 Lower surface


(a) / 10.437.


) = 0.675.


(C) M=0675.


--
cr


t 1=0.648.


(dJ /= 0 706.
















ff) -=-074/.


0 2O 40 60 80 /00 0 20 40 60 80 /00


Percent chord


NATImAL mVISORY
COMMITTM pm *UMAIWT K&


CONFIDENTIAL


Figure 16. Pressure distribution for a modified NACA E5, 3-019
airfoil with 0.20-chord flap. a 40; S -8g.


Fig. 16a-f








NACA ACR No. L5G31a


CONFIDENTIAL


x Upper surface
a Lower surface


0


I











a
*1



c3
0



Q)

(t


(a0) /M0.435.








/_____ "-i',,, "2- ___ .

.^ .- ^ \ -^


0 XO 40 60 W0 /00 0 ZO 40 60 SO /00
NATIONAL ADrISOAY
Perc en / chord co...nEl i ME-MUT
CONFIDENTIAL

Figure 17.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a = 40; 8 = -12.


Fig. 17a-f


(cI .-U.o/Y.







NACA ACR No. L5G31a


CONFIDENTIAL
x Upper surface
0 Lower surface


( t) /=O. 75.














(e) f/=0.724.


h-) /-= 56.
cr






(b) 17=0. Q6.


CONFIDENTIAL


0 20 40 60 80 AK0


0 ZO 40 O0 WO 100


Percent chord


Figure 18. Pressure distribution for a modified NACA 65, 3-019
airfoil with 0.20-chord flap. a = 60; 8 00.


--fri
CELr


fo) = 0.4 35.


/d) /0


t /"= 0.703.


Fig. 18a-f


~p~-~







NACA ACR No. L5G31a


CONFIDENTIAL
x ,Upper ur ace


0 Lower surface


0 W0 40 O6 80 AIO 0 ZO 40 60 80 /00


Percent chord
CONFIDENTIAL


NATIONAL ADVISORY
cONMrmTT E AEAt M KS


Figure 19.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a = 6; S = -4.


0











-I
0.


o
c0


"-. -.1 Ipc













09 F1=0.703.










T'^'^-F


Fig. 19a-f







Fig. 20a-f


NACA ACR No. L5G31a


CONFtIDENTIAL
x Upo,-r surface
0 Lower surface


0 20 40 0 60 /W00 0 20 40 60 W0 A0


Percent chord


NATIONAL ADVISO Y
CONWITTtt fo MIUAUTICt


CONFIDENTIAL


Figure 20.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a = 60; 8 -8


0















Q.
o 0
Q,


(0) = 02440.








NACA ACR No. L5G31a


Fig. 21a-f


CONFIDENTIAL


x UpDePr s,1*c7,,
0 -ower Sur'o-e


) -=0.704.
,-



(o)l 11-0 .704.


0 20 40 60 80


/00 0 20 40 60 80 /00


Percent chord


NATION L ADVISORY
COMIIIT111 FM uAOUICS


CONFIDENTIAL


Figure 21.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a = 60; 6 = -120.


N.,

Z -/


o
SU
Al


fc) / '1 0.674.

















fe) =0.72 7.


-CP
1 cr








Fig. 22a-f


NACA ACR No. L5G31a


x Upper surface
O .L owner surface


CONFIDENTIAL


(e) 1= 0.703.


-p
Icf


0 X 40 W 60 80 /00 0 20 40 60 80 /00


/F rcen/ chord
CONFIDENTIAL


NATION. *nISOXv
."17Tm 0l AMMUTKS


Figure 22.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a 100; S 00.


-Q -z
.Q)



0
U


\




























(d) /=0-674.

















(i) ( = 0720.
\f)7=O O









NACA ACR No. L5G31a


CONFIDENT

UzDver acSr'L

Low er sur oce


'IAL





( Cr






b,,

(b) /4= 0 566-


0 20 40 60 60 /00 0 20 40 60 d0 ,'00

Perc er,'I chord ATIONA. .... VI.D
CONlITI[I 101 aUI0OMiur
CONFIDENTIAL

Figure 23.- Pressure distribution for a modified NACA 65,3-019
airfuil with 0.20-chord flap. a = 10U; 8 = -40


-2-


A

-0-


-3


-2


-1


0





















-2










-2


-i



0


f


fa)M a 0 4.54.


Fig. 23a-f







NACA ACR No. L5G31a


Fig. 24a-f


CONFIDENTIAL


0 20 40 60 60 /00 0 20
Percent chord

CONFIDENTIAL


40 60 80 'Cc


NATIONAL ADVISORY
coUlIatt FOB UAUItICS


Figure 24.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a = 100; 8 -8








NACA ACR No. L5G31a


CONFIDENTIAL


x Ucper 5srfa.:e
0, Lower surface


.r'
I = 4 9.
A (dl fr'. 0.439.


^___I


- f) /F=0.72.


Ct


0 O0 40 60 c0 /00 0 20 40 60 80 /00


Perce n/ chord


NATIONAL ADVISORY
COMMITTEE Fi AEAUMLUTICS


CONFIDENTIAL


Figure 25. Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a = 100; 8 = -120.


-/


0









o



, -


kj


Fig. 25a-f


j~






NACA ACR No. L5G31a


CONFIDENTIAL


x L/O -'" surface
0 Lower surface


P7=0620


S- .


(d0 /M=O. 702


0 20 40 60 80


/00 0 ?0 40 60 6i0 O?


Percent chord


CONMITTH f AlM.OMKS


CONFIDENTIAL

Figure 26.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a 20; 8 4.


~)r4


r~ k-


(l M=0.733.


~7~-"


~) N=L~6 74.


' "


Fig. 26a-f


I


I -









NACA ACR No. L5G31a


Fig. 27a-f


CONFIDENTIAL

x U,'oer s5Jr/oca-

0 Low~ver surface


.-,


0
ct
* U


?C-




I

eC) l1--0676.




















fe) f1-O0732.


0 20 40 60 80 /00 0 20 40 60 80 /00


Pl-er ce n Chor


NATIONAL ADnvSORY
CONMITTEI roe AIUPOUTICS


CONFIDENTIAL


Figure 27.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a = 20; 8 = 80.


f





;) "7=-0744 .








NACA ACR No. L5G31a


CONFIDENTIAL

x p ce" surface


0 Lower surface


0 20 40 50 80 /00 0 20 40 60 0 'CO


Percent chord


NATIAL AVISOYv
Car aTT MnurKS


CONFIDENTIAL


Figure 5P.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.PO-chord flap. a 20; b 120.


V/1 0 6 /15.

















d) M=0.7C2.


-/


0


/










a-


- /77 z0o 753.


Fig. 28a-f







NACA ACR No. L5G31a


CONFIDENTIAL
x Upor sur face
0 Lower surface


a /I= 0430.


0 20 40 60


O0 A90 0 20 40 60 80 lW


Percent c/orfd
CONFIDENTIAL


NTmIImAL ADISOTI
COMITY FN MtlaMUTUIS


Figure 29.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a = 40; 8 = 4.


-/ I


S

.1 -

.I -


Fig. 29a-f


~pb~fi-







NACA ACR No. L5G31a


CONFIDENTIAL
x paperr surface
0 Lower surface


o(a) N 0.434.


0 20 40 60O 50


SM =0. 705.


/00 0 20 dO 60 60 /00


P'erc/nf chord


NATtlONrm ADVISORY
COMNITTit FM AMUMITS


CONFIDENTIAL


Figure 30.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a 40; 8 80.


-I



s 0


Fig. 30a-f









NACA ACR No. L5G31a


CONFIDENTIAL
x Upper surfoc
0 Lower suratoce


li)7"1 :0.'t5,3








(c) /- 0643.








' iCJ A^- 0.4j.


(b) : 0.j 4.

















(d) 1= 0. 70/.















SCOMITM M L UTICS

/ 0.- 74.


0 2O 40 60 80 /00 0 0O 40 60 60


Percent chord
CONFIDENTIAL

Figure 31.- Pressure distribution for a modified NACA 65,3-019
airfoil with 0.20-chord flap. a = 40; b = 120
I; -


Fig. 31a-f










NACA ACR No. L5G31a


Fig. 32 a-e


0

4O





O
r-+
0







0
t




0


O,'-*

S0o
COI


0

O0


00










OO
0









Q.
41
a
















-b
05















-1







(II
CM

















a)
Hr














*^l
(x.











NACA ACR No. L5G31a


Fig. 33, a-e


,-4
-I
0


rl






(d
-4




to


.










( W

0
Z I







*00
oo
E


ti
U
0


















,C

oC






+4




*H
0 u





S-4


to





to
a)


Li
I







bo
*-I
V4


o













, 1



z



f-4
r.4
z

0
0
o



uo
I-



-S
'-N

-4c
4-
55




















0










NACA ACR No. L5G31a


Fig. 34a-f


O
I






<*
C)O
0-0
0













ZE
z I










Cd
F03
0
E


ti
0

CO
Q.





0 x
0 -



















C-)
00r





















.-1
C3
1) -.-




C 0



rd(
CO



r-1












ad
Lc,


r-.











NACA ACR No. L5G31a


Fig. 35a-f


0

Ll
**-4

Cd




I



to



UN











0-
2 I











-Co
a u



CO







Cd








0.
ci







-I









-C
Lm








c'-.
3





CL
cu"









0



I
a.
LD _)









"Z











av'
*-
1-1
rZ.











NACA ACR No. L5G31a


Fig. 36a-e


*-4




0
1--
L.
..-r
id



oL)








iD

C,
<0












-o


uS
Sd
La
0

a
,C
3U
o r-


'cl








W .
u


CL
0







bo .-
Q.
a






c0Z
.CO 0


u


03
C 0
0-C



Li)i


4

CI
a












;
-a
a o



M











-t
a-U

1a
C,














-a
E.






r

o


E-
cs)











NACA ACR No. L5G31a


Fig. 37a-f


*I--
"-4
0
-4


Cd


1-4
0

O



i0


zo




-4
o'2
E


d
-1l





o ,-4
10












4.


0 .-4

o (
Jl o
CL
be
*I
o
/IM
0.














,a-
0o ..





-4




r-4
a,








0-







cr.







NACA ACR No. L5G31a


1 T I~J


IA


o


L


-Zu
4 \ b









%,COO* %t tog
1 \ S


- I
a.
LL.
z
0
U







_Q_


Y/7


S


t )


y O .Q W Q i

'^ *^ywru i^ow k)W


Fig. 38a,b



0
o
CD
Cu
0)

I-











0
cd
0
-I







0 0






C (d
ua)


4,
a,











C)

cui

:s








Fig. ,39a










-3

C2-.



-J


Cm0






'0
-i0


NACA ACR No. L5G31a


./ .2 .3 .4 .6 .7 .
Mach number, /M


(al a = -20.

Figure 39.- Effect of compressibility on section normal-force
and moment coefficients.


-4C
-8
-4


CONFIDENTIAL
I


~ N_ RY


rrr 7.. ~.- --l=r.SZi


~








NACA ACR No. L5G31a




./ -


Fig. 39b


--4




-2







-.4







-.6


Mach number, /M


(b) a = 00.


Figure 39.- Continued.


CONFIDENTIAL


6
/awr)


cr







Fig. 39c







.3 -



.2







0-

C
-/









./2

-.4 -



.4-





cm<^


NACA ACR No. L5G31a


(c) a = 20.


Figure 39. Continued.








NACA ACR No. L5G31a











.4



.3












.2 ------













C _


0 ./ .2 .- .4 .3" .6
Mach number, M



(d) a = 40.


Fig. 39d













*


.7 .8


Figure 39.- Continued.








Fig. 39e




.a-



.7-







.S-



.4

'


I I
NATIONAL ADVISORY
COMMITTEE FOI AEONAUTICS



-4 +-+ +-
mpi


- 0
/ L
V


NACA ACR No. L5G31a


.2 .3


CONFIDENTIAL

.4 .5 .b
Moach number, M


(e) a = 60.


Figure 39.- Continued.







NACA ACR No. L5G31a


CN
u-4,









-0 1









(f) a = 80.

Figure 39.- Continued.
./~-/ -- --------- --t-^____^.










Figure 39,- Continued.


Fig. 39f








Fig. 39g












.8



.7


CM


NACA ACR No. L5G31a


NATIONAL ADVISORY
COMMITTEE OI AERONAUTICS





Y- -



CONFIDENTIAL

? .4_ .5- .-- .7 .-


AMach


number, Ml


(g) a 10.


Figure 39.- Continued.


.7


r








NACA ACR No. L5G31a












.9






C,
.7



.6


-/
.5



.4







.Z2



--- -/

ifl I


Fig. 39h


NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS.


0 2 3 4 5- 6
Mach number ,M

Ih) a = 120.

Figure 39.- Continued.


.7 .6








NACA ACR No. L5G31a


6
fd


CONFIDENTIAL -


Mcr----


I \
1_ .


.5






,--
4













NATIONAL ADVISORY
SI I CONNITTEL F0m AE t cAUTKS


7-


-


CONFIDENTIAL


I I __ I __ __ L __ __ __


. .4 .5 .
Mach number, M


(i) a 20; positive flap deflections.


Figure 39.- Continued.


0---0-
4 V-
8 A-
/2 .


0



-/


-2


. -------"


1 I


7


1 1 41


Fig. 39i


.8a


./ .6


i)







NACA ACR No. L5G31a







/0



.9







.7

c,


-./



* -2
O
-0


Fig. 39j


NATIONAL ADVISORY
COMMITTEE FOr AEROiUTICS


CONFIDENTIAL
I I I I I


./ .1? .j .4 .- .6 .7 .6
Mach number,MA


(j) a = 40; positive flap deflections.


Figure 39.- Concluded.


-0--c


I







NACA ACR No. L5G31a


CONFIDENTIAL




-./1 -4 + + ------- K/


-2,






-.4









-.7





.I
NATIONAL ADVISORY
CO-- xMITTEE FA ,o TICS-







0- 4,----------

CONFIDENTIAL
O M T~ FOI ... .....S


0


/1 .2 .3 .4 .5 .6 .7 .8
Mach nmMbe/M


(a) a -2.

Figure 40.- Effect of compressibility on section hinge-moment
and flap normal-force coefficients.


Fig. 40a







NACA ACR No. L5G31a


0 .i .2 .J .4 .5 .6
MAoch number, /A1


(b) a = 00.


Figure 40.- Continued.


Fig. 40b








Fig. 40c


.2 -



./ -



0-



--./



c-.2 -


NACA ACR No. L5G31a


NATIONAL ADVISORY
COMMITTEE FWO AERONAUTCS.


-12
-4 -
0-


CONFIDENTIAL
/ x F 7


/4och number, /41


(c) a 20.


Figure 40.- Continued.


. / .-,,


.^i .^/








NACA ACR No. L5G31a


741..aai


.J .4 .5 .6
Mach number, M


(d) a = 4 .

Figure 40.- Continued.


Fig. 40d


5 CONFIDENTIAL





-0

-4




-2a







NACA ACR No. L5G31a


CONFIDENTIAL


NATIONAL ADVISORY
COMMITTEE FO AEUt AUTICS.



o 0I


CONFIDENTIAL
/,- .. .3 .4 .3__


M./ .c .4 ., .A
Moch number, M


(e) a 6.


Figure 40.- Continued.


Fig. 40e


.7 .a








NACA ACR No. L5G31a


Fig. 40f


NATIONAL ADVISORY
COMMITTEE FOM AERONAUTICS


/ -' CONFIDENTIAL
S .4 .3' .
Mach number, M


.7 .8


( f) a = 8.


Figure 40.- Continued.








Fig. 40g





.J



.2



.1







-.1





-.












ch 6


NACA ACR No. L5G31a


NATIONAL ADVISORY
COMMITTEE FM AEAIIAUTICS.


0 ./ .2 .3 .5 .6

Moch number,M


(g) a 100.

Figure 40.- Continued.


.7 .Q







NACA ACR No. L5G31a


.5



.4



.3



.2


Fig. 40h


CONFIDENTIAL
I I


1 *1


NATIONAL ADVISORY
COMNIT1EEI T0 AENAIUTICS



S|
k ----


0 ./ .2 .3 .4 .S .6
Mach number, M


(h) a = 12.


Figure 40.- Continued.


^ I


.7 .6







Fig. 40i




.7


.6






.4


.3



C-7


.1










Co


NACA ACR No. L5G31a


0 ./ 2 .4 .5 .6 .7 8
Moach number, /M


(i) a 2o; positive flap deflections.

Figure 40.- Continued.








NACA ACR No. L5G31a


"- -- -- CONFIDENTIAL



.7









./






4.4






.1 -------------- __ ---



0 ----------------- -


.4 .5 .6
Mach number, M1


(j) a = 40; positive flap deflections.


Figure 40.- Concluded.


Fig. 40j










Fig. 41a


NACA ACR No. L5G31a


-c



LO
**(
















o
-14















E





0


C




jo1
Eo

I..










[-co








U
0
O






U
0o
zd

.C


'0 N '0 O








NACA ACR No. L5G31a


Fig. 41b


































-~ -
t o


.,."t






b
0 rl


^ tS


N 0 C\


~o $









Fig. 41c


lzqll


NACA ACR No. L5G31a

























ONL
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NACA ACR No. L5G31a


Fig. 41d


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Ls









NACA ACR No. L5G31a


oo
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0 0-




r~
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Fig. 41e








NACA ACR No. LSG31a Fig. 41f








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z -: z
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Figs. 42,43


NACA ACR No. L5G31a


Q.
0













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NACA ACR No. L5G31a


f(a) a -- -.


(b) a=O-


III
I___-- CONFIDENTIAL

0 .2 .4 .5 .6 .7 .9
Mach number M


Figure 44.- Effect of compressibility on balance-chamber
pressure-coefficient differerrtial, PL-PU"


Fig. 44a-c








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NACA ACR No. L5G31a


(d) a = 4


(e) cr = 6


CONFIDENTIAL
f) cc= 8.
/lach number, 11


Figure 44.- Continued.








NACA ACR No. L5G31a



















4-.
I





S0












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tQ 0__./__ .2.


Fig. 44g,h


(9) c =io/.


Figure 44.- Continued.







NACA ACR No. L5G31a


.6 .7. .


Figure 44.- Concluded.


Fig. 44i,j







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