Wind-tunnel investigation of control-surface characteristics

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Material Information

Title:
Wind-tunnel investigation of control-surface characteristics
Alternate Title:
NACA wartime reports
Physical Description:
22, 45 p. : ill. ; 28 cm.
Language:
English
Creator:
Hoggard, H. Page
Bulloch, Marjorie E
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 )
Aerodynamic load   ( lcsh )
Genre:
federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Bibliography:
Includes bibliographic references (p. 19-20).
Statement of Responsibility:
by H. Page Hoggard, Jr., and Marjorie E. Bulloch.
General Note:
"Report no. L-205."
General Note:
"Originally issued April 1944 as Advance Restricted Report L4D03."
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 - 003804344
oclc - 123553411
System ID:
AA00009408:00001

Full Text





















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NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS






WARTIME REPORT

. ORIGINALLY ISSUED
... April 1944 as
Advance Restricted Report L4D03


S: VIND-TUNKEL INVESTIGATION OF CONTROL-SURFACE CHARACTERISTICS

XVI PRESSURE DISTRIBUTION OVE AN NACA 0009 AIRFOIL WITH

l.l- l 0.30-AIRJOIL-CHORD BEVELED-TRAILING-EDGE FLAPS

By H. Page Hoggard, Jr., and Marjorie E. Bulloch


Langley Memorial Aeronautical
Langley Field, Va.


WASHINGTON


ZTIM :REPORTS are reprints of papers originally L
teurAh results to an authorized group requiring then
Lijtd# a security status but are now unclassified.
t ,.All have been reproduced without change in oi


l ,! L 205
IsX,,.J ..,: H,: JiyiX... : .. .


Laboratory


issued to provide rapid distribution of
for the war effort. They were pre-
Some of these reports were not tech-
der to expedite general distribution.



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NATIONAL ADVISORY CO'.ITTTEE FOR AEROIU.:UTICS


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'V D- TT'lTUCL I1D rSTGA/\TIT'n O CONTROL- SURFACE C:;. jRAC i I TIC S

XVT PRtESSURT. ,IST'TrIBUTION OVER AMr NACG!. 000? AIRFOTL "'ITH

0. 0-ATIROTL-CHOrDH BEVELED-TR ILII G-DJE FLAPS

Ey H. Page Hoggard, Jr., and !Parjorie 2. Bulloch


SUfI.TAW'[


Pressure-distribuition tests have been made in the
IHACA 4- bv (-foot vertical tunnel of i plain flan with
interchangeable beveled trailing edges on an
HIACA 0009 airfoil. The flap chord was 50 percent of the
airfoil chod and the bevel chords were 15 andi 20 percent
of the flap chord. The 15-percent bevel w'a tested with
the bevel corner faired with both large and small radii.
The ournose of these tests was to sunplr pressure-
distribution data that may bI used. for structural and
aerodynamic design of' horizontal and vertical tail sur-
faces.

The results are presented as diagrams of resultant
pressure coefficients and of increments of resultant
pressure coefficient for the airfoil with the flap h.'.ving
beveled trailing edges. The diagrams are presented for
the control surface with the gap at the flap nose sealed
and unsealed.

A comparison of the beveled -flan pressure data with
plain-flap data indicated that the addition of a bevel
reduced the nrossures over the entire airfoil, including
the peak at the airfoil nose, and caused a reversal of
pressure over the beveled part of th. flap. ihe
normal-force coefficient for the beveled-trailing-edge
flap was less than th, coefficient for the plain-airfoil-
contour flap. The opn gap produced a tendency toward
overbalance by de crashing the negative pr.ussurjs over
the upper surface of a flap when deflected downward.
The results gen-:rally were in fair agreement with force-
test data pr.-viousl-r published.








2 NACA ARR No. )L+DO0


IN TRODUC TION


The National Aivisory Counmittee for Aeronautics has
instituted an extensive investigation of the aerodynamic
characteristics of various control surfaces. The force-
test data from this investigation have been suirmarized in
reference 1. The two-dimensional pressure-distribution
data obtained as part of the investigation have been
analyzed and the variation with flap chord of the various
aerodvramic characteristics of a flap has been presented
in reference 2.

.Two-dimenslonal force tests have been previously run
on a similar model of an :TACA 0009 airfoil with several
beveled trailing ed-us; the results ,f these tests are
presented in reference 3 (:-lso zuwrmarized in reference 1).
From the results of these force tests of trailing-edge
shapes having various included trailing-edge angles and
other airfoil eosts, a method based on the included angle
at the trailing edge has been found for predicting the
value_ of hinge-moment pi'aneters to be expected from a
bevVl. This correlation: can be found in figure 150 of
reference 1.

The tw'o-dimensional-flcw'i tests presented herein were
made to investigate the pressure acting on a control sur-
face with a beveled trailing edge. Such data should be
valuable for strrct,'ral design of the control Purfaces,
for exyrpan-3tion of the balancin, action of the tevel,
asn. for- 9tudy of boLndary-laye" conditions. The investi-
gation v'is r.,-i*e at all an-les of attack and flap deflec-
tios-is coisid,_.-ed n:_cessar-; for the strucu'ial design of
ail'erons, elevators, an1 rudders.





cp fla- chord rem.'war:l of fl-p hinge axis, percent
a'rfoil chord

c cho--d of basi: airfoil with fl.ap neutral

q dynamic rresnure of free air stream


P pressure coefficient







NACA ARR Yo. LLi.D03


PR resultant pressure !oefficient

APR increment of resultant pressure coefficient

p static pressure at a poirt on airfoil
po static pressure in free air stream
ao angle of attack for infinite aspect ratio

6f flap deflection
M Mach number, ratio of local velocity to speed of
sound

cn airfoil section noriE, l- force coefficient (n/qc)

c airfoil section pitchini.-moment coefficient
about quarter-chord point of airfoil (m/Ac2)

onf flap section normal-force coefficient (nf/qcf)

chm. flap section 1i.nge-r.omear coefficient (i/r-t2)
n normal force of airfoil section per unit span

m pitching moment oC airfoil section about quarter-
chord point per unit span

nf normal force of flap section per unit span

hf hinge moment of flap section per unit .span

C (nCn



cna
nfree a och


n6a6


SnA







NACA ARR !To. TL403


Cn^f








Pc = K-o)
o fh a
_P=

P C






The subscripts outside the parentheses indicate the
factors held constant during the measurement of the
paramo tc r.
Srbscripts:

IT point on upperi surface

L point on lower surface
h resultant

APPARATUS A.iD MODELS

The tests wore m-de in the ikCA 04- by (-foot vertical
tunnel. Th.e test section of this tunrnel has been con-
verted from the crig.iil open, circular, 5-foot-diameter
jet (reference )) to a closed rectangular 4- by 6-foot
test section, as shown in figure 1. The model completely
spanned the tesu section; therefore, two-dimensional flow
was appror:imated.







lIACA ARR "o. L4DO0


The model used for the pressure-distribution tests
of this investigation was designed to be an exact co')y of
the model used for the force tests in reference 5 but
with only the 0.15cf and 0.20cf b3veled-trailing-edge
shapes. The O.15cf bevel was tested with the bevel
corner faired with both large and smill radii. The
2-foot-chord model was made of laminated mahogany to the
modified NACA 0009 profile (table I). The airfoil was
equipped with a 0.50c plain flop, as shown in figure 2(a).
A gap of 0.005c was provided at the flap nose. The flap
was constructed with interchangeable blocks that formed a
beveled trailing edge and a thickened profile, as shown
in figure 3 of reference 5.

A single chordwisc row of pressure orifices was
built into the upper *:,rid lower surfaces of the airfoil
and flap at the midspan location. The orifice loca-
tions are presented in figure 2(b) in percent of airfoil
chord from the leading, edge. The copper tubes from the
pressure orifices were brought out of the :.iodel at one
end through the tor:.ue tube and the tunnel wall to a
multiple-tube, open-faced manometer. Rea dings were
recorded by a camer'i.


TEST


All of the tests, except those with larg-e flap de-
flection and high positive angle of attack (flap deflec-
tion, 500 and 450; angle of attack, 1L.50 and 19.50) were
run at an average dynamic pressure of 15 pounds rer
square foot. The large flap deflections at high positive
angles of attack required more power than was available to
maintain a dynamic pressure of 15, pounds per square foot;
therefore, these tests were run &t an average dynamic
pressure of 12 pounds per square foot. The airspeed in
the test section at dynamic pressures of 15 and 12 pounds
per square foot is about 76 and 69 miles per hour, respec-
tively, at standard sea-level conditions. The corre-
sponding values of effectivee .Ae-rnolds number are
2,760,000 and 2,208,000. (Effective Reynolds num-
ber = Test Reemolds number x turbulence factor; the turbu-
lence factor of the I;ACA 4- by 6-foot vertical tunnel is
1.95.)

The tests were made at angles of attack ranging from
-20 to 200 at intervals of 5 and at angles giving maxi-
mum positive and negative lift. It may be noted that all








HTACA ARR No. L4D03


angles of attack are offset from the exact values of
00, 5o, 10, 150, and 200 by -0.7 owing to an error
in setting the zero angle of attack. This error was
found to be consistent throughout the tests and the data
were corrected accordingly. The model was tested with
the 0.50c Dlain flap deflected O0, 1, 20, 50, 100,
150, 200, 25, 500, and 450. 'he tests were run with
the flap gap both open (O.O05c gap) and sealed with
plasticine. During the tests with 300 and 450 flap
deflection, pressure orifice 15 for the lower surface
(fig. 2(b)) was sealed because its position at both
large flap deflections was inside the gap.

Check tests were made for each flap deflection as
an indication of the accuracy of the test results.
'T.hen the 0.005c cap was used, the check tests were made
after both angle of attack and flap deflection had been
reset. The sealed-2ap check tests had only the angle
of attack reset, because the plasticine seal would have
to be refaired if tne flap deflection were changed.

The speed of the tunnel was maintained at the test
value of Q for approximately 2 minutes before readings
were recorded in order to allow the alcohol. in the
manometer tubes to reach the correct height.


RESULTS

Presentation of Data

The results of the pressure-distribution tests are
given in the form of diagrams of resultant pressures with
flap neutral and resultant-pressure increments caused by
varying the flap deflection. The resultant pressures and
increments of resultant pressure are presented for the
various bevel and gap combinations and for various angles
of attack in figures 5 to 1C. The resultant normal
pressure at any point along the chord of the airfoil was
determined by taking the algebraic difference of the
pressures normal to the upper and lower surfaces of the
airfoil at that point. All diagrams of resultant pres-
sures or resultant-pressure increments of the airfoil
and flap comoInations are plotted as pressure coeffi-
cient P1 cr as ,P." The resultant pressure coeffi-
cient is defined as


PR = PL P








?TACA ARR To. IJ4DO3 7


whe re

PTT PO
P- -
q


SPL- Po
q

P pressure coefficient

p static pressure at a point on airfoil

Po static pressure in free air stream

q dynamic pressure of free air stream

and the subscripts

U upper surface

L lower surface

R resultant

The resultant-pressure diagram for any condition may
be obtained by adding the distribution at a given angle
of attack and the distribution at a given flap deflection.
A comparison of resultant-pressure distributions over the
bevel juncture with large and small radii is presented in
figure 11 at several angles of attack and flap deflec-
tions.

Pressure d..stributions fnr th*e urner and lower sur-
faces of the fl.ap hrW' i a G.l5cI I hrvcl with sealed gap
are presertet in figl-re 12 for .'T:hXoV.s angles of attack
and flan defec.tion.. The r.-su.tant pressures over the
NACA OGOQ airfoil with 0,50c p:'a n flep and sealed gap
(reference 5) are compa-ed wiich :he i-esultant pressures
over the modified airfoil with 0.15cf-bevel flap in
figure 15. Figure I4 presents upper- and lower-surface
pressures over the plain flap and LWe 0.15cf-bevel flap
for the came conditions for which resultant pressures
are given in figure 15.

The rates of change of pressure coefficient with
angle of attack and with flap deflection are presented for








NACA ARR No. L4DO0


the various bevel and gap combinations in figures 15
to 18 for convenience in calculating distributions at
small values of ao and 6f. The flap section normal-
force coefficient as a function of flap deflection is
presented for all combinations of' bevel and gap in
figures 19 and 20 at several angles of attack. Com-
plete chordxise pressure distributions for various
combinations of ao and Of that might occur on the
horizontal tail of a dive bomber in highly accelerated
maneuvers at various speeds are presented in figure 21
for the 0.15cf-bevel flap with sealed gap.

The section aerodynamic coefficients of the airfoil
and flap are presented as functions of angle of attack
for all bevel and gap combinations in figures 22 to 24.
The coefficients were obtained in each case by mechanical
integration of the original pressure diagrams.

The parameter values fcr beveled flaps are pre-
sented in table IT along .-/ith values for the plain-
airfoil-contour flap for convenient comparison. The
plain-flap parameter values wore obtained from refer-
ences 1 and 6.


Pre c i son

The angles of attack are believed accurate within
0.10. "lap deflections are believed accurate within
+0.20. Plotted values of pressure coefficient P are
correct within 2 percent except for peaks at the
leading edge and flap hinge a:i-s or for stalled ccn-
di tions.

Coefficient values calculated from check test points
have been plotted in figures 19 an.d 22 and are designated
by flagged symbols. Many of the points come within the
accuracy of the plot; others vary a negligible amount.
The accuracy of the corrected zero angle of attack is
indicated by the deviation from zero of lift and moment
coefficients at zero angle of attack. From figures 19
and 22, It aopears that the maximum error in setting the
angle oe attack at zero lift is 0.20. This discrepancy
may be caused by flow misalinement in the tunnel or by an
asymmetrical nodel.

Two-dimensional flow having been approximated, the
results may be considered as section characteristics.







NACA ARR No. LjD03


Experimental tunnel corrections were applied only to the
airfoil section normal-force coefficient Cn. Although
no corrections were made for the other coefficients, the
tunnel values are believed to be higher than the free-air
values and hence are on the conservative side for struc-
tural purposes. The magnitude of the airfoil resultant
pressure coefficients as represented in the resultant-
pressure diagrams (figs. 5 to 10) is known to be too large
by about 7 percent because these curves were clotted
directly from manometer records without the application
of the experimental tunnel correction, which allows for
the increase in lift produced by tunnel-wall interference.


DISCUSSIG"i

Resultant-Pressure Distribution

The resultant-pressure diagrams should prove useful
in determining loading conditions for the structural
design of ailerons and horizontal and vertical control
surfaces. Tests have indicated that the increments of
pressure and the increments of section aerodynamic
coefficients caused by flap deflection are approximately
independent of the airfoil section for airfoils of
approximately the same maximum thickness and thickness
distribution (references 7 and 8). It is therefore
believed that, for structural design, the incremental
data presented herein may be applied to other basic
sections of approximately the same thickness and thick-
ness distribution. The increments of the section aero-
dynamic coefficients may be taken from figures 22 to 24
by using the flap-neutral curve as a reference line.

From a study of the incremental-resultant-nressure
curves for the stalled conditions (ao = 19.50 and -20.7)
for both bevel chords and :-ap conditions (figs. 4, 6, 8,
and 10), it appears that the bevel continues to reduce
the flap hinge moment in the stalled condition from the
hinge moment for a plain flap under the same conditions.
The tests of beveled elevators on the fuselage of a
typical pursuit airplane also indicated that the bevel
was effective in the stalled attitude and reduced the
floating angle of the elevators by about 100 (reference 9)
from the angle at which airfoil-contour elevators would
float. The resultant-pressure curves (figs. 5 to 10),
especially for the 0.005c gap, show a tendency toward a
decrease of resultant pressure over the main airfoil just
ahead of the flap.







FACA ARR No. L4DO3


The results indicate that the size of the radius at
the bevel juncture is relatively unimportant in its
effect on the loads over a beveled-trailing-edge flap
(fiC. 11).


Pressure Distribution over Upper and Lower

Surfaces of Beveled Flap

The distributions presented at various angles of
attack and flap deflections in figure 12 indicate that
only on the surface of the flap which is deflected
against the relative wind does the bevel affect the
pressure distribution to any great extent. The only
exceptions occur at low angles of attack and small flap
deflections, for which the upper- and lower-surface
distributions show nearly equal effect of bevel. The
pressure distribution on the side avvay from the rela-
tive wind, when at large angles of attack or flap deflec-
tion, resembles that of a flap and tab in a stalled
condition.

It will be noticed in figure 15 that the resultant-
pressure peak at the flap hinge axis is higher for the
beveled flap with the 0.005c gap than for the beveled
flap with the sealed gan. Inasmuch as the resultant
pressure is the algebraic difference of the upper- and
lower-surface pressures at any point, the positive peak
on the lower surface makes the resultant-pressure peak
higher. (See fig. 1l.)

The pressure distribution produced over the upper
and lower surfaces of a flap by a beveled trailing edge
is compared with the pressures over a plain flap in
figure l4. The effect on the pressure distribution of
the bevel on the surface deflected against the relative
wind is more pronounced when the .;ap is open. The main
effect of the open gap on the flap pressure distribution
anpears to be the decrease in magnitude of the negative
pressures over the upper surface of the flap, which
results in a tendency toward lower or even overbalanced
hinge moments.


Curves of Pa and P6

For convenience in calculating the pressure distri-
butions over both surfaces for small values of ao







NACA ARR No. L4DO3


and 6f, the curves of Pa and P6 were calculated and
are presented in figures 15 to 13. From the experi-
mental data, it was found impossible to predict with any
degree of accuracy the variation of pressure with angle
of attack over the nose of the airfoil because the
stagnation point moves considerably and the pressures
change rapidly and are not linear with angle of attack.
The variation of pressure with angle of attack over the
rest of the airfoil appeared from these tests to remain
a linear variation only from 00 to o0; therefore, the
Pa-curves should not be used for calculating pressures
beyond a value of ao of 5.

The variation of pressure with flap deflection for
any point on the airfoil contour appeared from these
tests to be linear to 50. The P6-curves therefore
should not be used for flan deflections greater than 50.
The final pressure distribution:required is found by
multiplying the values of Pa and P6 by the values
of ao and 6f for which the distribution is desired
and adding algebraically to the basic distribution
(P at to = 6f = GO) given in the louer part of fig-
ures 15 to 18.


Flap Section Normal-Force Coefficient

For all combinations of bevel and gao tested, the
values of cnf were smaller than for the plain flap with
sealed gap at the same angles of attack. The values
of Cnfl and Cnf for beveled and plain flap may be
conveniently compared in table ITI. The variation
of cnf as a function of angle of attack is clearly
shown in figures 19 and 20. The effect of a is small
at 6f = 280 with the gap open and at 6f = 200 with
the gap sealed.


Pressure Distribution on Horizontal Tail For

Highly Accelerated Maneuvers

The flight condition during which high structural
loads and the formation of a compression shock on the
horizontal tail are most likely to occur is a highly







NACA ARR No. LFDO0


accelerated maneuver in which the horizontal tail is
operating at a high angle of attack at high speed. The
pressure data presented herein are not applicable to
tail design for high-speed flight unless they are cor-
rected for the variation of pressure with Mach number,
which is given approximately by the relation 1i/ M2.
Theoretical variations of pressure' with i'\ach number are
compared with experimental pressure-distribuuion data at
various Mach numbers in reference 10. The pressure
distributionspresented in figure 21 at angles of attack
of -0.70, 5.70, and 10.70 and with flap deflections
of 0, -5, -10, and -15 are test data that cover the
highly accelerated maneuvers estimated from unpublished
dive-bomber test data.


Aerodynamic Section Characteristics

Normal-force coefficient.- The force-test lift data
of reference 5 are given in terms of section lift
coefficient whereas the pressure-distribution data are
given in terms of normal-force coefficient. Inasmuch as
the lift coefficient and normal-force coefficient have
nearly the sarme value, this value is referred to as "lift"
in the following discussion.

The slope of the lift curve -- from table II

for the airfoil with O.15cf beveled trailinga'odge and
sealed gap is 0.088 as compared with 0.091 from the
force-test data in reference 3. These results are in
fair agreement if account is taken of the fact that
different models and methods of calculation were used for
the force and pressure tests.

The lift-curve slopes from the force and pressure
tests for the 0.20cf bevel wi-th sealed gap have the same
value, 0.092. For the open cap the lift-curve slopes
from the force and pressure tests are, respectively,
0.088 and 0.087 (table II and reference 1). The lift-
curve slopes obtained from the pressure-ctistribution tests
appear to check very well with the force-test results.
Opening the gap appeared to change the angle of attack at
which the stall occurred by about 10. This angle of
attack, approximately 120 with flap neutral, was not
affected by bevel chord.







NACA APR No. 14DO3 13


The values of lift effectiveness given in

table II were taken at zero lift and show the expected
decrease in effectiveness as a result of the beveled
trailing edge. The small radius on the bevel juncture
increased Cna about 0.003 for open and sealed gap when
compared with the lift-curve slope for the large-radius
bevel. Reducing the radius at the oevel juncture
decreased the effectiveness from -0.56 to -0.52.

The parameter nafree (tabls II) is a measure of
free
control-free stability cnly at ao = Of = 0. The
values in table II indicate the cxn-cted tendency of the
beveled flap to float upward at a s.naller angle than the
plain flap.

A method for estimating. the. pressure distribution
(and normal force) over a bevel from available tab
pressure-distribution data is given in the appendix.
The results of this method are 11 ustiated and a com-
parison is made in figure 25, at several angles of attack
and flap deflections, between actual and estimated pres-
sure distributions for a 0.20cf bevel with sealed gap
and a:n included angle at the trailin- edge of 250

Flan hinge-rmoment coefficient.- The values of chf

(table II) were taken over the linear part of the hinge-
moment curve, which was over a sr.all range (t5) for
the O.005c-gap tests and a larger range (1100) for the
sealed-gap tests (figs. 22 and 24). The values of chf
(table II) were taken from 6f = 0 to 6f = 50 because
the curve appeared linear over this range. For a com-
plete picture of the effect of various bevel and gap com-
binations, all the hinge-mcment curves (figs. 22 to 24)
must be taken into consideration anad too much reliance
should not be placed on the slope values measured over a
small part of each curve, except for stick-free stability
calculations.

The values of chf and ch as found for the
O.15cf and 0.20cf bevels with sealed gap are in fair
agreement with the values of reference 5. Values of both
hinge-moment parameters for the 0.20c bevel with 0.005c gap







WACA ARR No. L4D05


were road from the curves in figure 19 of reference 1 and
were found to be in fairly close agreement. The values
of Chfa and chf6 for both bevel chords were found
to fall near the correlation curve of figure 150 in
reference 1 vrith less scatter than tne average scatter
of the correlation points.

From the values of hinge-moment parameters in
table II It appears that decreasing the radius at the
bevel juncture tends to decrease the negative values
of chf6 for bouh gap conditions. Decreasing the
-6
radius had no effect on the value of chfa when the gap
was open but decreased the positive value when the gap
was Sealed.

Pitching-momecnt coefficient.- The slopes of the
curves of pitching-mo-,rent coefficient as a function of
lift coefficient at a constant tngle of attack and at a
constant flap deflection arie given in table II. The
aerodynsmic cent-er of additional lift caused by varying
the angle of attack generally was located at approxi-
mately the 0.22c stac'on for the sealed-gap tests and ths
0.21c station for the O.00'c-gap tests. The bevel chord
had little effect on the location of this aerodynamic
center.

The aerodynar:ic center at which the lift produced by
flap deflection may be considered to act is located at
approximately the 0.-lc station for either gap condition.
All aerodynamic-center locations for the gap-sealed condi-
tion are in fair agreement with the values presented in
reference 3.


C 0NC LUS ION S


Pressure-distribution tests have bean made in the
NACA 4- by 6-foot vertical tunnel of a plain flap with
interchangeabl.- beveled trailing edges on an
NA' 0009 airfoil. The flaz chord was 50 percent of the
airfoil chord and the bevel chords were 15 and 20 percent
of the flap chord. The results of these tests indicated
the following conclusions:







NACA ARR No. 14D03


1. At a given angle of attack and flap deflection,
the addition of a bevel reduced the resultant pressures
over the entire airfoil, except for the pressure at the
flap hinge axis, including the peak at the airfoil nose,
and caused a reversal of pressure over the beveled part
of the flap.

2. The normal-force coefficient for the beveled-
trailing-edge flap was less than the coefficient for the
plain-airfoil-contour flap with the airfoil at the same
angle of attack and the flap deflected through the same
angle.

3. The open gap at the flap nose gave the flap a
tendency toward overbalance because of a decrease in the
negative pressures over the upper surface of a downward
deflected beveled flap and because of a slight increase
in the negative peak on the lower-surface bevel juncture.

4. The size of the radius used to fair the bevel
juncture appeared to have no appreciable effect on the
pressure distribution developed.

5. The results obtained from the pressure-
distribution tests generally were in fair agreement with
force-test results of a comparable arrangement.


Langley Memorial Aeronautical Laboratory,
National Advisory Committee for Aeronautics,
Langley Field, Va.







NACA ARR No. L4DO3


AFPENDIX


METHOD FOR CALCULATING PRESSURE DISTRIBUTION OVER A BEVEL

F-Or,! TAB PPRSSURE-DISTRIBUTION DATA


When an elevator, aileron, or rudder is designed,
the general practice is to use the total load over the
surface. '.lotion pictures of bulged fabric on ailerons
in high-speed dives indicate that the pressures along
the chord should be used to determine how securely the
covering must be fastened to the structural members. In
the case of a beveled surface for which a pressure peak
occurs at the bevel. juncture, a sti:dy of the chordwise
distribution might prevent a covering failure. A method
for predicting the chordwise pressure distribution over
a beveled surface without having to test it is advan-
tageous, particularly as such a method supplements a
method already established for predicting the hinge-
moment characteristics.

4 netho'd for predicting the ucord-wse load distribu-
tion on the flap is described herein. Ilo attemot is
made to oradict flap section hinge-moment coefficients;
the hline-.nonent correlation based on the included angle
at the trailing edge (for sealed-gap condition) may be
found in figure 150 of reference 1.

The bevel contour was dt-velped (fig. 3 of refer-
ence 5) by d';leoinz a 0.20)'> tab 100 and deflecting
the flap slightly each vay ;o keep the tab trailing edge
ccntered on ti-e airfcll chord lini. Inasmuch as the
bevel prcfile was developed by using deflected-tab
contours, it was decided to use tab pressure diagrams to
etLituaUe the pressure distribution of a beveled flap.
Only the upper-surjace distribution for a tab deflected
dov.nward aid tne lower-surface distribution for a tab
deflected upward are considereJ. It is necessary to
correct these pressures by means of P6 to allow for the
small flap deflections necessary to keep the tab trailing
edge centered on the airfoil chord line. The resulting
diagrams (fig. 2-,) were integrated and found to give
values of cnf that were in good agreement with the
bevel test data for flap deflections of 100 and 200 at
values of co of -0.70 and 4.50 (figs. 25(c), 25(d),
25(g), and 25(h)). rhe value of cnf based on tab







NACA A 17 No. L$D0O


data was in general somewhat larger than the bevel test
value.

At the smaller flap deflections, the values of ,nf

from tab data were generally much larr than from bevel
test data but, from a comparison of t:.e values with those
for a plain flap in figure 20, the estimated values were
found to be closer to the bevel test value than to the
plain-flap values.

In order to use thLe present correlation method, it
is necessary to have )ressuire-distribution diagrams for
a flap and tab of the desired chords. The tab chcrd
should approximately equal the distance from bevel
juncture to trailing rdr-e.

The included angle of the bavel :lust oe reproduced
by the correct tab and flan deflections. These deflec-
tions must be found in order that the tab-deflection
diagram may be chose and corrected. The following
equation gives the amount that the flap must be deflected
to keep the tab trailing ede centered cn the a'rfoil
chord line:
= .-s evel ~ airfoil
1 Ct Si.11 e---1-2-a ----f-C---
66f = sin-
687 = 3 in -------=-------
cf ct

where

Obevel included an-le at trailing ea,'e of bevel (for
which prediction is bTein; mad3)

Vairfoil included angle at trailing edge of airfoil
from tests of which Mlao -nd. tab pressure
diagrams are to oe used

ct chord of tat, percent airfoil chord

cf chord of flap, percent airfoil chord

V.'ith 66f, ,bevel, anrd airfoil known the uangle
through which the tab is deflected 6t to reproduce
the included anrle of the bevel may be found by the fol-
lowing equation:


6= AL + wtevel Oairfoil
t- .^







18 NACA ARR No. L4D03


It may be noticed in figure 25 that the tab data used
*vere for 6- = 10 whereas equation (1) gives
6t 3.40. By using the diagrams for 6b = 100, the
included angle was found to he 27.60 instead of the
correct value of 250; but, inasmuch as the correlation
for the hinge-moment naramreters based on included angle
shows a change of 0.001 in the value of the hinge-moment
parameters for a change of 2" in the included angle, there
could be only a 31ight change in the size or shape of the
pressure diagram.







NACA ARR No. LIDO0


REFERENCES


1. Sears, Richard I.: Wind-Tunnel Data on the Aerody-
namic Characteristics of Airplane Control Surfaces.
NACA ACR Ho. 5LOS, 19135.

2. Ames, Milton B., Jr., and Sears, Richard I.: Deter-
mination of Control-Surface Characteristics from
NACA Plain-Flap and Tab Data. UACA Rep. No. 721,
1941.

5. Jones, Robert T., and Ames, Milton B., Jr.: Wind-
Tunnel Investigation of Control-Surface Character-
istics. V The Use of a Peveled Trailing Edge
to Reduce the Hinge Moment of a Control Surface.
NASA ARR, March 19.2.

4. Wenzinger, Carl J., and Harris, Thomas A.: The
Vertical Wind Tunnel of the National Advisory
Committee for Aeronautics. NACM Rep. No. 587,
1031.

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

6. Sears, Richard I.: Wind-Tunnel Investigation of
Control-Surface Characteristics. I Effect of
Sap on the Aerodynamic Characteristics of an
NACA 0009 Airfoil with a 30-Percent-Chord Plain
Flap. hACA AiRR, June 19ll1.

7. Allen, H. Julian: Calculation of the Chordwise Load
Distribution over Airfoil Sections with Plain,
Split, or Serially Hinged Trailing-Edge Flaps.
NACA Rep. Ho. 63), 1953.

8. Allen, H. Julian: A Simplified Method for the Calcu-
lation of Airfoil Pressure Distribution.
NACA TN No. 708, 1959.

9. Gillis, Clarence L.: Characteristics of Beveled-
Trailing-Edge Elevators on a Typical Pursuit
Fuselage at Attitudes Simulating Normal Flight and
Spin Conditions. NACA ARR, Dec. 1942.







20 NACA ARR No. I4DO3

10. Stack, John, Lindsey, W. F., and Ltttell, Robert E.:
The Compressibility Burble and the Effect of
Compressibility on Pressures and Forces Acting on
an Airfoil. NACA Rep. No. 646, 1938.






IACA ARR ND. L"D05


cTriFIAT1E" C'F n'EFTIi.L' "AC7 0009

Stations and ordinates in percent of


A Ti Fnl T1L

airfoil chord


Station


0

1.25

2.5

]ICA 0009
7.5
airfoil
1I 10
section
15,
20


50





60
70
Straight
So
90

100


L.L. radius: 0.39


TALE I


Orcdinate



1 .12

1.96

2.67

5.15

3.51

14 .01

.50


1;.50

1+. 55

+5.97



2.83

-I .02

1 .08












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NACA ARR ?ro. LLDO3




NACA ARR No. L4D03


Model mountedd in NA CA 4-by 6 -fool iedrcal tunnel


Model


---4"--24
.4 A48"


NATIONAL ADVISORY
CIIMM11lEE FOR AERONAUTICS


F gure I .- Insa/aloion of beveled- raisingg edge pressure-
dcsf/rbu//on model in NVACA 4-by 6-fool verhcol /unnel.


Fig. 1





NACA ARR No. L4D03


S/raliht contour from Aer-e
Io here/ed rai/ing edge
Airfoil wi//l .15ci bevel


R283 R= /33
.20c, bevel /15c bevel wilh
small radius


(a) Two-fool-chor NAMCA 0009 a/r/ al mode/ vil/h a
0.30 c pla/4 flap having 01/5c, and 020cc beveled
Ira/ing edges. Dimensions are in percent of a/rfoil
chord.
NAIIORAL ADVISORY
CuMMIIEt FORl AERONAUTICS
/5
r6 7 1/ 12 /314 67 ,. 2/ 23
-----~\8-----20 22


loca-
fhon
5500
6750
70.00
72S0
7500
WODO
90,00
8500
9000
92M
9400
9600
98DO.


(6) Chordwise locarlons of pressure orfices on airfoil
and on the flaps Aaving 0/6cland 020cF hebe/s.

Fgure2. -Di'ensfons and chordnwse pressure-or/ice /ocal'ona
for NA/CA 0009 beveled-/f roa/n-etdge pressure-d/s/ribuhon model/
Dimensions and ori/ce local/ons are /. percent of akrfol chord


Fig. 2




NACA ARR No. L4D03


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