Analysis of available data on the effectiveness of ailerons without exposed overhang balance

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Title:
Analysis of available data on the effectiveness of ailerons without exposed overhang balance
Series Title:
NACA WR
Alternate Title:
NACA wartime reports
Physical Description:
16 p., 14 leaves : ill. ; 28 cm.
Language:
English
Creator:
Swanson, Robert S
Crandall, Stewart M
Langley Aeronautical Laboratory
United States -- National Advisory Committee for Aeronautics
Publisher:
Langley Memorial Aeronautical Laboratory
Place of Publication:
Langley Field, VA
Publication Date:

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

Notes

Summary:
Summary: A considerable amount of two- and three-dimensional data on the effectiveness of ailerons without exposed overhang balance has been collected and analyzed. The trends indicated by the analysis have been summarized in the form of a few approximate rules concerning the effectiveness parameter Pyramid symbolalpha/Pyramid symboldelta (at constant lift): Thickening and beveling the trailing edge (as measured by the trailing-edge angle Ø) will generally reduce the effectiveness about 0.3 percent per degree of bevel for ailerons sealed at the hinge axis and about 0.6 percent per degree of bevel for unsealed ailerons. A 0.005c gap at the hinge axis usually reduces the effectiveness approximately 17 percent for flap chord ratios of 0.2. This percentage increases as the flap chord ratio is reduced. The effectiveness is about 14 percent lower at aileron deflections of 20° than at aileron deflections of 10°. At large angles of attack (a = 10°) and for chord ratios of about 0.2, positively deflected ailerons are approximately 20 percent less effective than negatively deflected ailerons. The deflection of partial-span flaps has no consistent effect on the effectiveness. Increases in Mach number and forward movement of the transition point decrease the aileron effectiveness.
Bibliography:
Includes bibliographic references (p. 13-16).
Statement of Responsibility:
by Robert S. Swanson and Stewart M. Crandall.
General Note:
"Originally issued May 1944 as Advance Confidential Report L4E01."
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."

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University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003807486
oclc - 126887221
System ID:
AA00009426:00001

Full Text

1- 1


NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS





WARTIME REPORT
ORIGINALLY ISSUED
May 1944 as
Advance Confidential Report L4EO1

ANALYSIS OF AVAILABLE DATA ON THE EFFECTIVENESS OF
AILERONS WITHOUT EXPOSED OVERHANG BALANCE
By Robert S. Swanson and Stewart M. Crandall

Langley Memorial Aeronautical Laboratory
Langley Field, Va.









^YM AC-A.


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


L 171


ACR No. L4E01



































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/analysisofavaila001ang









iIACA ACR .:o. L4,,01


NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS


ADVANCE CONFIDENTIAL REPORT


ANALYSIS OF AVAILABLE DATA ON THE EFFECTIVZH'SS OF

AILERONS WITHOUT EXPOSED OVER HAUG BALAITCE

By Robert S. Swanson and Stewart M. Crandall


SUM1MAERY


A considerable amount of two- and three-dimensional
data on the effectiveness of ailerons without exposed
overhang balance has been collected and analyzed. The
trends indicated by the analysis have been summarized in
the form of a few approximate rules concerning the effecv-
tiveness parameter Aa/A6 (at constant lift): Thickening
and beveling the trailing edge (as measured by the
trailing-edge angle f) will generally reduce the effec-
tiveness about 0.5 percent per degree of bevel for ailerons
sealed at the hinge axis and about 0.6 percent per degree
of bevel for unsealed ailerons. A 0.005c gap at the hinge
axis usually reduces the effectiveness approximately
17 percent for flap chord ratios of 0.2. This percentage
increases as the flap chord ratio is reduced. The effec-
tiveness is about l4 percent lower at aileron deflections
of 200 than at aileron deflections of 100. At large angles
of attack (a = 100) and for chord ratios of about 0.2,
positively deflected ailerons are approximately 20 percent
less effective than negatively deflected ailerons. The
deflection of partial-span flaps has no consistent effect
on the effectiveness. Increases in Mach number and forward
movement of the transition point decrease the aileron ef-
fectiveness.

No consistent deviation of the experimentally deter-
mined values of static rolling moments from those values
predicted by the lifting-line-theory method could be de-
tected. Because the several factors neglected in the
lifting-line theory apparently are fairly small and
counteract one another, on the average, no additional
correction need be applied.


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INTRCODUT:ON


As a part of' the -nerral lateral-control investi-
g-ation by the '",CA, the large amount of two- and three-
dimensional data on the rolling effectiveness of ailerons
without exposed overhang balance is collected and ana-
lyzed in the present paper. The main purpose of the
analysis is to detQrmin3 any fairly consistent variations
in the effectiveness of these ailerons with the various
design variables and criterions of similitude.

7h. secondary -urpose of the analysis is to evaluate
experimentally the limitations of lifting-line theory v:i
with regard to the estimation of aileron rolling moments
fro~-. section data.





CL wing lift coefficient

CLmax maximum rinr- lift coefficient
"max
c, section lift coefficient

CL wing rolling-moment coefficient

a angle of attack, degrees

6 flap or aileron deflection, degrees

b wing span

y spanwise location

Yo spanwise location of outboard end of aileron

Y1 soanwise location of inboard end of aileron

S wing area

A aspect ratio (b2/s)

X taper ratio, that is, fictitious chord at tip
divided by chord at root


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NACA AT' "o. L. 1- 01


A a/A5




(62/65)


wing chord at any section

flap chord at any section

aileron chord at any section

airfoil trailing-edge angle, degrees

Mach number, ratio of free-stream velocity to
velocity of sound

Reynolds number

slope of curve of section lift coefficient
against angle of attack at constant 6



slope of curve of section lift coefficient
against flap deflection at constant c

section flap effectiveness parameter, that
is, absolute value of ratio of equivalent
change in angle of attack to angle of
flap deflection measured at constant lift

aileron effectiveness parameter, that is,
ratio of equivalent change of angle of
attack to angle of aileron deflection;
su-'cript 5 indicates that values are
computed from three-dimensional test
data by use of lifting-line theory


K theoretical or experimental correction to
lifting-line-theory values of rolling

moment


T wind-tunnel turbulence factor

Subscript 6 = 00 to 100, 6 = 00 to 150,
or 6 = 00 to 200, etc., indicates range over
which Aa/A6 or (Aa/A6)3 is evaluated.


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DATA

Scope


The characteristics of the two-and three-dimensional
models and the air-flow characteristics of the wind tun-
nels in which the models were tested are summarized in
tables I and II, respectively. The data as given in the
original reports (references 1 to 2h) were in many cases
uncorrected for the effects of the jet boundaries or for
model deflections. It was found essential to apply such
corrections before the data could be correlated.


Reduction and Presentation of Data

Section data.- The effectiveness parameter Aa/A6
is the section characteristic of flaps that determines
their ability to provide rolling motion when installed
as ailerons on an airplane, provided the analysis is
based on aileron deflection rather than stick force.
This c'rameter Aa/A6 is equal to the absolute value of
the change in angle of attack necessary to neutralize
the lift produced by a unit flap deflection. The effec-
tiveness parameter was determined from the section data
of references 1 to 15 by plotting a against 6 for a
constant section lift coefficient c7 and measuring the
slope (absolute value of slope us-i) of a straight line
thro'iuh 6 = 00 and 5 = 100 for the effectiveness at
small flap deflections (a/!A6)o=00to 100 and
throucii 6 = 0 and 5 = 200 for the effectiveness at
large flap deflections (Aa/6)6=00to20o6. It is often
convenient in the analysis to consider the limiting case
of 6a/68, which is equivalent to For prac-

tical purposes, the values of (Aa/6)5=0ootol0o are very
nearly equal to the values of ~2/65.

Static three-dimensional data.- In references 25
and 26 are presented charts for estimating the rolling
moment caused by aileron deflection. The charts were
calculated frcm lifting-line theory for various wing and
aileron plan forms for a slope of the section lift curve
of 0.099 per degree. The general method for using the


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N'OA A]CR TTo. LEO1 l


charts to determine the rolling-moment coefficient CL
is to evaluate graphically the following integral across
the aileron span:

C7
ao (A d- -
CL= K d (la)
b/2

or

C- a= 6 K -d1C (lb)
0.099 J A6 a

wh e r e

ao slope of section lift curve, per degree

K experimental or theoretical correction to
lifting-line theory (to be evaluated)

Aa/A5 experimental section lift effectiveness of
ailoron

01/a is determined from the charts of reference 25 or 26,
and y is measured along the wing span. If Aa/65 is
constant across the aileron, the integral is equal
to Aa/A6 times the difference between the end values
of C /a.

Most of the models studied.had ailerons of constant
chord ratio and da/A6 thus was a constant. By in-
serting experimental values of CL and chart values

of CV/a in equation (la), ao (a or its equiva-
a A0.099 A 5
lent a K-- therefore could be evaluated. A few
0.099 A6
erroneous values in references 25 and 26 were corrected
and the curves were refaired to be similar to the known
fairing for elliptical wings. By using section data to
estimate ao and Aa/&6, the value of K could be
experimentally evaluated. If section data for evalu-
ating ao were not available, a was estimated by
usini the measured slope for the finite-span wing in the
lifting-line-theory formulas (reference 27) corrected for


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NACA :.CR No. LTEOl


the edge velocity by the methods of reference 28. For
these cases in which ao could not be satisfactorily

estimated, the data are presented as 0o /
0.099 \A/3

"or the few cases in which the aileron chord ratio
was not constant across the aileron span, the integral
of equation (la) was evaluated by using the section data
of figure 1 to estimate the variation of aa/A6
with ca/c; thus an effective average value of ca/c,
weighted according to the ability of each spanwise ai-
leron section to produce rolling moment, was evaluated.


DISCUSSIC7:

Effect of Viscosity


From figures 1 and 2, the effectiveness Aa/Ab of
sealed plain flaps and ailerons with small trailing-edge
angles is seen to be considerably less than the theo-
retical values for thin airfoils. Most of the decrease
in effectiveness may be attributed to the influence of
viscosity. The effective surface or boundary of the
airfoil is displaced from the actual surface by the
amount of the so-called displacement thickness, which
is the height of the mean ordinate of the velocity dis-
tribution in the boundary layer. Because the shapes
and thicknesses of boundary layers vary with pressure
gradient, transition location, Reynolds number, Mach
number, gap at hinge axis, etc., the effective airfoil
shape varies with these factors.

The flap effectiveness Aa/A6 is less than the
theoretical value because the rate of increase of the
thickness of the boundary layer with flap deflection,
which results from the high adverse pressure gradient
behind the hinge axis, is usually greater than the rate
of increase of the boundary-lay-r thickness with angle
of attack. The slope 6c;/6)a is therefore decreased
more by viscosity than is (c6L/6a)6; Aa/A6 is thus
decreased by viscosity. The larger the flap deflection,
the smaller the effectiveness Aa/A6. -he section data
of figure 3 and the finite-span data of figure l show


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NACAACR r'o. L)4EO1


that, at low angles of attack, the effectiveness at flap
deflections of 20 is approximately 14 percent lower than
the effectiveness at flap deflections of 100. At high
angles of attack, approximately the same reduction occurs
(fig. 5), except for the gap-unsealed condition in which
little consistent reduction is in evidence.

The effect of viscosity upon the aileron effec-
tiveness depends markedly upon the pressure gradient. T
The direction of the deflection of an aileron would be
expected to have little effect at small angles of attack
because the pressure distribution at 6 = 0 is very
nearly the same on both surfaces. The data of figure 6
verify this deduction. At high angles of attack, however,
negative aileron deflections reduce the adverse pressure
gradient whereas positive aileron deflections increase
the adverse pressure gradient. A lower effectiveness
thus may be expected for positive aileron deflections.
The data of figure 7 indicate that, at an angle of at-
tack of 100 and for chord ratios of about 0.2, positively
deflected ailerons are about 20 percent less effective
than negatively deflected ailerons. This effect appears
to increase with aileron chord ratio.

The gap at the flap hinge axis allows the low-energy
boundary-layer air to leak from the pressure side to the
suction side of the airfoil. The boundary layer on the
pressure side is thus thinned and on the suction side is
further thickened with a resulting reduction of the lift
increment. The effect of the gap on the lift-curve slope
due to angle of attack (6cl/6a) is fairly small,
because the pressure difference across the hinge axis is
small. The slope (6c7/66)a, and consequently Aa/A6,
is considerably decreased, however, because the maximum
pressure difference due to flap deflection is usually
located at the hinge axis. Figure 8 shows that a 0.005c
gap at the hinge axis decreases.the effectiveness about
17 percent for flap chord ratios of 0.2. This reduction
appears to be larger for flaps of smaller chord.

A forward movement of the transition point usually
increases the thickness of the boundary layer and thus
decreases the flap effectiveness Aa/A6. This effect is
shown qualitatively in figure 9, in which data are pre-
sented from tests with and without the nose of the air"-.
foil roughened in order to fix transition. The position


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8 C0"ID'7TTAL NACAAR No. LLE401


of the transition point on the unroughened airfoil was
not determined. Some unpublished computations and ex-
perimental data indicate that a reduction of about
2 percent in Aa/A6 for a forward transition movement
of 0.1c may be eypoected with sealed plain flaps. The
effects of viscosity are usually greater with increased
thickness and beveling of the airfoil trailing edge.
The effect of transition movements and gaps thus are
greater for airfoils with lar:- -railing-edge angles 0.
Gaps at the aileron hinge axis also increase the loss
in Aa/&6 that results fror transition movements.

The effect of beveling the trailing edge of the
flap is presented in figure 9, in which the effectiveness
(Aa/A6)5=00to100 is shown as a function of the trailing-
edge angle '. Reductions of about 0.4 percent per
degree of bevel for sealed flaps and of about 1 percent
per degree of bevel for unsealed flaps are indicated.
Th- three-dimensional data of figure 10 show a decrease
in aileron effectiveness of about 0.3 percent per degree
of bevel for sealed ailerons and approximately
0.6 percent per degree of bevel for unsealed ailerons.

It should be noted that, under some particular con-
ditions, viscosity may increase Aa/,o to values even
greater than those for the theoretical thin airfoil. The
explanation for this rather astonishing fact is quite
simple. The effectiveness parameter Aat/A is equal to

the ratio of the lift-curve slopes If vis-

cosity decreases (6,0j/6a)6 more than it decreases
f06 2, the effec iveness. parameter. Aa/a6
is increased. For a few conditions, markedly low lift-
curve slopes (6cL/6a') occur over a small range of
an-le of attack a. Also, the slope bcL/66')a is less
affected or is affected over a different range of a.
Over a limited range of a, very large values of Aa/A/6
may therefore occur. A few cases in which this phenomenon
has been observed are: (1) negatively-deflected ailerons
at large angles of attack near the stall, (2) so-called
linked-balance ailerons with which a gae through the wing
occurs well ahead of the hinge .axis and allows very low
values of (6c7/6a) but has little effect on (6c/66)a


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NACAACR ifo. IJ._E Ol


near a = 00, and (53) ailerons on low-drag airfoils
with large trailing-edge angles, which usually have a
very large value of effectiveness Aa/A6 near the angles
of attack where the transition point suddenly shifts for-
ward (near boundary of low-drag region) and causes a
break in the curve of cl against a.


Effect of Compressibility

Data on the effect of Mach number on
(&a/A6)_=-lOOto l00 are shown in figure 11. The data
are rather limited and subject to some doubt because it
is extremely difficult to determine accurately the wind-
tunnel corrections at large values of Mach number. Cor-
rections for model twist and deflections were applied to
the data. Increasing the Mach number usually de4rsn6-0'
creases Aa/A6. From figure 11, it may be seen that an
increase in ITach number from 0.2 to 0.-5 reduces the
effectiveness about 7 percent.

The simple theory of Glauert and Prandtl indicates
no effect of Mach number on Aa/A6 because (6cL/6a)
and (6c7/668) are assumed to be increased equally by
compressibility. Experimental data indicate, however,
that (6cz/6a), is usually increased a little more
and (60/68)a a little less than the Glauert-Prandtl
relation would account for. The explanation appears to
be related to the thickening of the boundary layer and
the transition changes that have been observed at high
Mach numbers. It is believed, therefore, that below the
critical speed the main effect of compressibility cp
upon Aa/A6 is to increase the effects of viscosity.


Corrections to Lifting-Line Theory

The limitations of lifting-line theory for the esti-
mation of aileron hinge-moment characteristics from
section data were discussed in'reference 28. The aspect-
ratio corrections to the hinge 'moment determined from
lifting-line theory were shown to be inadequate whereas,
for the cases in which lifting-surface-theory calcula-
tions (reference 28) are available-, the aspect-ratio cor-
rections to the hinge moment determined from lifting-
surface theory are shown to be satisfactory. The large


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,.GA ..CR No. L-lO1


difference between the results of the two theories may
be illustrated by the fact that the a-je t-ratio cor-
rections to the slope of the curve of Linge moment
against angle of attack determined from lifting-surface
theory are about twice as great as the corrections
determined from lifting-line theory.

The aspect-ratio corrections to the slope of the
lift curve .aiIr.st angle of attack (references 29 and 30)
for moderate aspect ratio as determined by the lifting-
line and lifting-surface theories differ by only about
7 or 8 percent. The aspect-ratio corrections to the
da:.ping moments of elliptical wij7 rotating about the
lateral plane of symmetry as determined by the two theo-
ries also differ by only about 7 or 8 percent (unpub-
lished correction determined by the methods of refer-
ence 30). The difference bsteen the two aspect-ratio
corrections to the slope of the lift curve age:rst flap
deflection is about 5 to 4 percent, which is only about
one-half as much as that for the slope of t'he lift
curve against angle of attack. This difference exists
primarily because the distance to the three-quarter-
chord point (point for best measure of effective angle
of attack of wing) from the center of load that results
from aileron deflection is roughly one-half the distance
to the three-quarter-chord boint from the center of load
that results from changes in angle of attack. The ef-
fective lc.--th of the trailing vortices thus is less for
the load that results from flap deflection. It might
therefore be expected that the aspect-ratio correction
to the static aileron rolling moments determ-i1ed from
lifting-surface theory would be of the same o3er, 5
to 4 percent greater than the value determined from
lifting-line theory. In any case, the aileron rolling
moments determined from lifting-line theory should be
much closer to the experimental values than the aileron
hinge ::omnents would be.

It may be seen that the section data (fig. 1) and
the finite-s;psn data with the lifting-line-theory aspect-
ratio corrections applied (fig. 2) are in fairly good
agreement. Although there is considerable scatter, the
curve faired through the section data represents vtry
well the finite-s.:.an data, especially for aileron chord
ratios of 0.2 or less. (See fig. 2.) -.n experimental
evaluation of the over-all- spect-ratio corrections shows,
on the average, no serious discrepancies (exceeding
10 percent) with the lifting-line-theory values; that is,
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NACAACR No. LICEO1


on the average, K = -= 1.0. The 3 or 4 percent
SAa/A6
increase in the aspect-ratio correction that might be
expected from a qualitative study of lifting-surface
theory (actual numerical values have not yet been calcu-
lated) may either be masked by the scatter of the data
in figures 1 or 2 or may be counteracted by three-
dimensional boundary-layer effects or by the effect of
the vertical location of the trailing-vortex sheet (ref-
erences 28 and 31).

Lifting-line theory indicates no change in aileron
effectiveness with deflection of partial-span flaps.
Some effect might be expected because of cross flow;
however, figure 12 shows that the 'deflection of partial-
span flaps generally has no consistent effect on aileron
effectiveness.

The available data on the effect of sweep and taper
(figs. 15 and 14) show that, insofar as aileron rolling
moments are concerned, no large corrections are to be
applied to lifting-line theory for the effects of taper
and sweep. For wings of low taper (large values of X),
it appears that the aileron effectiveness is slightly
greater if the wing is swept forward.


CONCLUDING 'L-u.FYS


The trends indicated by the analysis of available
data on the effectiveness of ailerons without exposed
overhang balance have been summarized in the form of a
few approximate rules concerning the effectiveness
parameter Aa/A6 (at constant lift): Thickening and
beveling the trailing edge (as measured by the trailing-
edge angle $) will generally reduce the effectiveness
about 0.5 percent per degree of bevel for ailerons sealed
at the hinge axis and about 0.6 percent per degree of
bevel for unsealed ailerons. A 0.005c gap at the hinge
axis usually reduces the effectiveness about 17 percent
for flap chord ratios of 0.2. This percentage increases
as the flap chord ratio is reduced. The effectiveness
is about 1l percent lower at aileron deflections of 200
than at aileron deflections of 100. At large angles of
attack '(a = 100) and for chord ratios of about 0.2,
positively deflected ailerons are about 20 percent less


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effective than negatively deflected ailerons. The de-
flection of partial-si,-n flaps has no consistent effect
on the effectiveness. Increases in "'Ach number and for-
ward movement of the transition point decrease the
aileron effectiveness.

To consistent c-rr-ction to the lifting-line-theory
method of estimating aileron rolling moments could be
detected. Because the several factors neglected in
lifting-line theory apparently are fairly small and
counteract one another, on th3 average, no additional
correction need be applied.


LenIley :Temrorial Aeronautical Laboratory,
National Advisory Comminttee for Aeronautics,
Langley Field, Va.,


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REFERENCES

1. Street, William G., and Ames, Milton B., Jr.:
Pressure-Distribution Investigation of an N.A.C.A.
0009 Airfoil with a 50-Percent-Chord Plain Flap
and Three Tabs. NACA, TI No. 734, 1959.
2. Ames, Milton B., Jr., and Sears, Richard I.:
Pressure-Distribution Investigation of an N.A.C.A.
0009 Airfoil with an 80-Percent-Chord Plain Flap
and Three Tabs. NACA TN No. 761, 1940.

z. Ames, Milton B., Jr., and Sears, Richard I.:
Pressure-Distribution Investigation of an N.A.C.A.
0009 Airfoil with a 50-Percent-Chord Plain Flap
and Three Tabs. NACA TN No. 759, 1940.

4. Sears, Richard I.: Wind-Tunnel Investigation of
Control-Surface Characteristics. I Effect of
Gap on the Aerodynamic Characteristics of an
NACA 0009 Airfoil with a 0-Percent-Chord Plain
Flap. NACA ARR, June 1941.
5. Jones, Robert T., and Ames, Milton B., Jr.: Wind-
Tunnel Investigation of Control-Surface Charac-
teristics. V The Use of a Beveled Trailing
Edge to Reduce the Hinge Mloment of a Control
Surface. NACA ARR, March 1942.

6. Sears, Richard I., and Liddell, Robert B.: Wind-
Tunnel Investigation of Control-Surface Charac-
teristics. VI A 30-Percent-Chord Plain Flap on
the NACA 0015 Airfoil. h--,CA ARR, June 1942.

7. Gillis, Clarence L., and Lockwood, Vernard E.: Wind-
Tunnel Investigation of Control-Surface Charac-
teristics. XIII Various Flap Overhangs Used
with a 30-Percent-Chord Flap on an MTACA 66-009
Airfoil. NACA ACR No. 5G20, 19435.

8. Sears, Richard I., and Hoggard, H. Page, Jr.: Wind-
Tunnel Investigation of Control-Surface Charac-
teristics. XI Various Large Overhang and
Internal-Type Aerody-naamic Balances for a Straight-
Contour Flap on the 3NACA 0015 Airfoil. NACA ARR,
Jan. 1943.


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9. Purser, Paul E., and Riebe, John M.: Wind-T..nnel
Investigation of Control-Surface Characteristics.
XV Various Contour Miodifications of a 0.5G-
Airfoil-Chord Plain Flap on an NACA 66(215)-01o
Airfoil. NACA ACR To. 5L20, 19435.

10. ;.nzinger, Carl J., and Delano, James B.: Pressure
Distribution over an N.A.C.A. 23012 Airfoil with
a Slotted and a Plain Flap. 'iCh Rep. io. 655,
1958.
11. Wenzinger, Carl J., and Harris, Thomas A.: Wind-
Tunnel Investigation of an N.A.C.A. 25012 Airfoil
with Various Arrangements of Slotted Flaps.
AGICA Rep. No. 664, 1959..

12. Crane, Robert M., and Holtzclaw, Ralph W.: Wind-
T'r.nel Investigation of Ailerons on a Low-Drag
Airfoil. I The Effect of Aileron Profile.
FACA ACR Fo. 4A14, 1944.

13. Denaci, TH. G., and 13r1'9., J. D.: jTind-Tunnel Tests
of Ailerons at Various Sneeds. II Ailerons of
0.20 Airfoil Chord and True Contour with 0.60
Aileron-2C>rd Sealed Internal Balance on the
EACA 66,2-216 Airfoil. '.-.CA ACR No. 5F18, 1945.
14. Davidson, Miltcn, and Turner, Harold R., Jr.: Tests
of an :'.CA 66,2-216, a = 0.6 Airfoil Section with
a Slotted and Plain Flap. NACA ACR No. 5J05,


15. Rojllo, F. M.: Collection of Balanced-Aileron Test
Data. I..\. ACR F.. _All, 1944.

16. Rogallo, F. 7., and Purser, Fsul 1.: .ind-Tivrinel
Investigation of 20-Percent-Chord Plain and, Frise
Aiiexrons on an '2A 25012 Airfoil. 2NACA ARR,
Dec. 1941.

17. Rogallo, F. T., and Schuldenfrei, Ia rvin: '."ir.d-
T'.: el Investigation of a Plain and a Slot-Lip
Aileron on a "'ir.C with a Full-Span Flap Consisting
of an Inboard Frwler and an Outboard Slotted Flap.
FACA ARR, June 1941.


CC '"TDF"TTAL








NACA ACR No. LIEOl1


18. Weick, Fred E., and '"enzinger, Carl J.: Wind-Tunnel
Research Comparing Lateral Control Devices,
Particularly at High Angles of Attack. I -
Ordinary Ailerons on Rectangular Wings. NACA
Rep. No. L-1.9, 1952.
19. 'Jeick, Fred E., and Shortal, Joseph A.: Wind-Tunnel
Research Comparing Lateral Control Devices,
Particularly at High Angles of Attack. V -
Spoilers and Ailerons on Rectangular Wings. NACA
Rep. No. 459, 1952.

20. X1Jeick, Fred E., and Shortal, Joseph A.: Wind-Tunnel
Research Comparing Lateral Control Devices, Par-
ticularly at High Angles of Attack. VIII.
Straight and Skewed Ailerons on v.ings with Rounded
Tips. NACA TN No. 4415, 1933.

21. Weick, Fre,3 E., and Wenzinger, Carl J.: Wind-Tunnel
Research Comparing Lateral Control Devices, Par-
ticularly at riHigh Angles of Attack. IX. Tapered
Wings with Ordinary Ailerons. NACA TN No. 449,
1935.
22. Wenzinger, Carl J.: ".'ind-Tunnel Investigation of
Tanered ,Tings with Ordinary Ailero0s and Partial-
Span Split Flaps. NACA Rep. No. 611, 1957.
23. Wenzinger, Carl J., and Ames, Milton B., Jr. Wind-
Tunnel Investigation of Rectangular and Tapered
N.A.C.A. 25012 C!ings with Plain Ailerons and Full-
Span Split Flaps. NACA. TN No. 661, 1958.
24. Irving, H. B., and Batson, A. S.: A Comparison of
Aileron Control on Tapered Wings with Straight
Leading Edge and Straight Trailing Edge. R. & M.
No. 1837, British A.R.C., 1938.
25. Weick, Fred E., and Jones, Robert T.: Resume and
Analysis of N.A.C.A. Lateral Control Research.
NACA Rep. No. 605, 1937.
26. Pearson, Henry A., and Jones, Robert T.: Theoretical
Stability and Control Characteristics of Wings
with Various Amounts of Taper and Twist. NACA
Rep. No. 635, 1938.


CONFIDENTIAL


C07'TLE!- TTIAL









I:ACA ACR No. L4E01


27. Anderson, Raym.ond P.: Determination of the Charac-
teristics of Tapered '"ings. 'ThCA Rep. To. 572,
1956.
28. Svanson, Robert S., and Gillis, Clarence L.: Limi-
tations of Lifting-Line -'eory for Estimation of
Aileron KFnge-?'cLent Characteristics. NACA
CB No. 5L02, 19'5.

29. Cohen, Doris: A I.:ethod for DeteImInrl.-i, the Camber
an- Twist of a Surface to Support a Given Distri-
bution of Lift. iT..;i TIN No. 855, 1942.

30. Jones, Robert T.: Theoretical Correction for the
Lif-t of -Jliptic -in:g. Jour. Aero. Sci., vol. 9,
no. 1, Nov. 1941, pp. 8-10.

51. Bollay, .illiam: A Non-linear Wting Theory and its
Application to Rectangular Wings of Small Aspect
Ratio. Z.f.a.M.M., Ei. 19, Heft 1, Feb. 1959,
pp. 21-35.

32. Jacobs, Eastman N., Abbott, Ira H., and Davidson,
'"I1ton: Supplement (loose-leaf) to I3.CA Advance
Confidential Report, Preliminary Low-Drag-Airfoil
and Flap Data from Tests at Larcg Reynolds
NFuibers and Low Turbulence. -jACA, March 1912.


CC'7T DENTAL


C OT DF I-ETTAL








NACA ACR No. L4E01


TABLE I.- SUPPLEMENTARY INFORMATION REGARDING
TESTS OF TWO-DIMENSIONAL MODELS


Model e
Basic Air-flow characteristics &
Desig- Sym- airfoil Type of flap "
nation bol 7 M 3

1 0 NACA 0009 Plain 1.95 0.08 --------- 1 to 5
2a + NACA 0015 Plain 1.93 0.10 1.4 x 106 6

2b NACA 0015 Internally balanced, 1.93 0.10 1.4 x 106 8

3 X NACA 23012 Plain 1.60 0.11 2.2 x 106 10,11
NACA Plain, 6
4 E 66(2x15)-009 straight contour 1.93 0.10 1.4 x 10

5 NAC 66-009 Plain 1.93 0.11 1.4 x 106 7

6 lNACA Internally balanced Approach- 0.17 2.5x 106 15
low drag ing 1.00
NACA
7 V 66(2x15)-216, Internally balanced Approach- 0.18 5 x 106
a = 0.6 ing 1.00
NACA
8 > 66(2x15)-116, Internally balanced Approah- 0.14 6.0 x 106 15
a = 0.6 ing 1.00
NACA
9 < compromise Plain Approach------ 13.0 x 106 ----
low drag ing 1.00
NAC A Appro6ach-
10 7 low drag Internally balanced Approach- 0.14 6.0 x 10 15

NACA Approach-6
11 635420)-521 Internally balanced ing 8.0 10 ---
(approx.)
aNACA 0.20 2.8 x 106
12 66(215)-216, Internally balanced pproach- 0to to 1
a = 0.6 ing 1.o0 0.48 6.8 x ln6

13 b 66(215)-216, Plain ng 0.1 8 x 10 12
a = 0.6 ing 1.0.

14 /1 66(215AC01o Plain 1.93 0.09 1.2 x 106 9


15 6 66,2-216, Plain ingApproa --- 6.0 x 14
a = 0.6


aThis designation has been changed from the form in which it appears in reference
to the form described on p. 21a of reference 32.


CONFIDENTIAL


CONFIDENTIAL











NACA ACR No. L4EOi 18





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CONFIDENTIAL


Fig. 1


(a) 6 range from 00 to 100.


Flap chord ratio, cf/c

(b) 6 range from 00 to 200.
Figure 1.- Variation of section flap effectiveness with
flap chord ratio for small Mach numbers and a small
range of trailing-edge angle. Gaps sealed; c = 0.
(Symbols designating two-dimensional models are
identified in table I.)


CONFIDENTIAL


'7-


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I Experimental









NAUA ACR No. L4E01


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Experimental-
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0
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----------------------------- -----------------------------------------zzh -- -I--

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0 .1 .2 .3

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(b) Gaps unsealed.
Figure 2.- Variation of aileron effectiveness with
aileron chord ratio for small Mach numbers and a

small range of trailing-edge angle. am 0.
(Symbols designating three-dimensional models
are identified in table II.)


CONFIDENTIAL


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CONFIDENTIAL


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UNIVERSITY OF FLORIDA


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