Tests of a linked differential flap system designed to minimize the reduction in effective dihedral caused by power

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Material Information

Title:
Tests of a linked differential flap system designed to minimize the reduction in effective dihedral caused by power
Alternate Title:
NACA wartime reports
Physical Description:
16, 28 p. : ; 28 cm.
Language:
English
Creator:
Pitkin, Marvin
Schade, Robert O
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:
Flaps (Airplanes)   ( lcsh )
Aerodynamics -- Research   ( lcsh )
Genre:
federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Summary:
Summary: An investigation has been made in the Langley free-flight tunnel to determine experimentally the effects of a linked differential flap system upon the effective-dihedral characteristics of a 1/10-scale powered airplane model. The differential flap system consisted of two individual flaps so linked as to operate differentially from an initial setting when free and designed to created rolling moments automatically opposing those created by slipstream effects. Tests were made on the Langley free-flight-tunnel balance and on a trim stand that permitted freedom in roll and yaw. The results of the test indicate that the negative dihedral changes caused by power may be materially reduced or completely eliminated by use of a differential flap system. Increasing the flap-differential ratio (ration of upgoing flap deflection to downgoing flap deflection) to a value above unity increased the effectiveness of the differential flap system in opposing the dihedral changes caused by power but diminished the tendency of the flaps to restore themselves to their initial setting of equal deflection. Little effect was observed when the flap-differential ratio was decreased to a value below unity. Differential flap action also increased the static directional stability.
Bibliography:
Includes bibliographic references (p. 16).
Statement of Responsibility:
by Marvin Pitkin and Robert O. Schade.
General Note:
"Report no. L-4."
General Note:
"Originally issued August 1945 as Advance Restricted Report L5F25."
General Note:
"Report date August 1945."
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 - 003631486
oclc - 71672044
sobekcm - AA00006272_00001
System ID:
AA00006272:00001

Full Text



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NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS





ARlTIME REPORT'
ORIGINALLY ISSUED
Aukt 1945 as
Advance Restricted Report L5F25

TESTS OF A LINKED DIFFEIRTIAL FLAP SYSTEM

-4 DESIGNED TO MINIMIZE THE REDUCTION IN
EFFECTIVn DIERAL CAUSED BY POWER *
.. By Marrin Pitkin and Robert 0. Schade

Langley Memorial Aeronautical Laboratory
langley Field, Va.








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NACA ARR No. L5F25 RESTRICTED

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS


ADVaNCE RESTRICTED REPORT


TESTS OF A LINKED DIFFERENTIAL FLAP SYSTEM

DESIGNED TO MINIMIZE THE REDUCTION IN

EFFECTIVE DIHEDRAL CAUSED BY POVER

By Marvin Pitkin and Robert 0. Schade


SUMMARY


An investigation has been made in the Langley free-
flight tunnel to determine experimentally the effects of
a linked differential flap system upon the effective-
1
dihedral characteristics of a --scale powered airplane
10
model. The differential flap system consisted of two
individual flaps so linked as to operate differentially
from an initial setting when free and designed to create
rolling moments automatically opposing those created by
slipstream effects.

Tests were made on the Langley free-flight-tunnel
balance and on a trim stand that permitted freedom in
roll and yaw.

The results of the tests indicate that the negative
dihedral changes caused by power may be materially
reduced or completely eliminated by use of a differential
flap system. Increasing the flap-differential ratio
(ratio of upgoing flap deflection to downgoing flap
deflection) to a value above unity increased the
effectiveness of the differential flap system in
opposing the dihedral changes caused by power but
diminished the tendency of the flaps to restore
themselves to their initial setting of equal deflec-
tion. Little effect was observed when the flap-
differential ratio was decreased to a value below unity.
Differential flap action also increased the static
directional stability.


RESTRICTED







NACA ARR No. L5F25


I'I TRODTU'CTION


The application of power in tractor airplanes
generally causes a large decrease in the effective
dihedral of such airplanes, particularly at low speeds.
For airplanes possessing initially small positive
values of effective dihedral in gliding flight, power
application may lower the effective dihedral to negative
values and induce large and unsatisfactory degrees of
spiral divergence. These adverse effects cannot be
simply eliminatedby the expediency of increasing the
initial amount of geometric dihedral because such a
change may provide an excessive amount of dihedral
with power off and lead to poor or unstable oscillatory
characteristics in power-off flight or in power-on
flight at high speeds.

From the previous considerations, it is desirable
to seek means of avoiding large dihedral changes due to
power. One possible solution proposed by Dr. H. S. Ribner
of the Langley Memorial Aeronautical Laboratory involves
use of a system of linkage whereby the flaps operate
differentially from an initial setting so as to create
rolling moments automatically opposing those created by
slipstream effects. The results of tests of such a
system in the Langley free-flight tunnel are reported
herein.

A powered model, representative of conventional
single-radial-engine fighter airplanes, was employed
for all tests. Most of the tests were made on a test
stand that permitted freedom in roll and yaw. Measure-
ments of rolling moments were obtained by use of a
calibrated-spring system. Necessary force-test data
were obtained on the Langley free-flight-tunnel six-
component balance. The effects of differential flap
action upon the dihedral characteristics of the model
were studied for various ratios of differential flap
movement. In most cases, the tests were made with
vertical and horizontal tail surfaces removed, although
a brief study was made of the effect of vertical-tail
area upon the effective dihedral.







NACA ARR ITo. L5F25


SYMBOLS


The coefficients and symbols are defined as follows:

CL lift coefficient (L/qS)

CD drag coefficient (D/qS)
GX longitudinal-force coefficient (X/qS)

C1 rolling-moment coefficient (L/qbS)

Cz- rate of change of rolling-momjent coefficient with
angle of sideslip, per degree (6CL/63)

Cjip rate of change of rolling-moment coefficient with
angle of yaw, per degree (6CL/6p)

L force along Z-axis, positive when acting upward,
pounds; moment about X-axis, positive when it
tends to depress right wing, foot-pounds

X force along X-axis, positive when acting forward,
pounds

D force along wind direction, positive when acting
rearward, pounds; diameter of propeller, feet

Y force along Y-axis, positive when acting to the
right, pounds

M moment about Y-axis, positive when it tends to
raise nose, foot-pounds

N moment about Z-axis, positive when it tends to
turn nose to right, foot-pounds

q dynamic pressure, pounds per square foot (pV2)

S wing area, square feet

c mean aerodynamic chord, feet

b wing span, feet

V airspeed, feet per second







NACA ARR No. L5F25


p

a





6f


dSfU/dSfD


6fU/6fD


Tc


V/nD

n

ACLf


St

6r

6rL


mass density of air, slugs per cubic foot

angle of attack, degrees

angle of yaw, degrees

angle of sideslip, degrees

flap deflection, degrees

right-flap deflection, degrees

flap-differential ratio (ratio of upgoing flap
deflection to downgoing flap deflection)

mean flap-differential ratio over the first
200 of incremental up deflection

thrust disk-loading coefficient
(Effective thrust/pV2D2)

propeller advance ratio

rotational speed, revolutions per second

change in lift coefficient caused by flap
deflection

vertical-tail area, square feet

rudder deflection, degrees

left-rudder deflection, degrees


THEORY OF DIFFFZTTIAL FLAP ACTION


An important part of the decrease in the effective
dihedral parameter Cj caused by power is produced
by the lateral displacement of the slipstream over the
trailing wing as the airplane is sideslipped. The
lateral center of pressure of the additional lift
induced by the slipstream moves outboard from its
original center position and creates a rolling moment
about the center of gravity of the airplane. The







NACA ARR No. L5F25 5


variation of this rolling moment with sideslip angle is
such as to reduce the effective dihedral.

On a wing with a flap, the increase in lift of the
trailing wing caused by slipstream displacement is
accompanied by an increase in flap hinge moment. The
hinge moment of the flap mounted on the leading wing
similarly is decreased when the airplane is sideslipped.

The flap system tested is shown in figure 1 and is
so designed as to utilize the change of flap hinge
moments caused by power in order that the dihedral
changes caused by power may be reduced. The system
consists of two flaps connected by a mechanical linkage.
The arrangement of the control rods and singletree is
such that upward deflection of one flap pivots the
singletree about its fixed center pivot and thus causes
downward deflection of the flap on the opposite wing.
The central bar, which provides the fulcrum for the
differential action, is used to deflect or retract both
flaps equally and is extended and locked in flap-down
flight.

When the apnlied hinge moment on the differential
flaps changes because of slipstream displacement, the
trailing-wing flap tends to rise and the leading-wing
flap tends to fall. As the trailing-wing flap rises,
its aerodynamic hinge moments decrease whereas those
of the leading-wing flap increase. At some differential
setting, equilibrium is again obtained. The aileron
effect of the differential-flap deflections produces
rolling moments that tend to compensate the rolling
moments created by slipstream displacement. In
addition to direct slipstream-displacement effects,
augmentation by the slipstream of the wing-fuselage
interference may make an important contribution to
the loss in effective dihedral due to power for low-
wing airplanes. The effect, however, of such a con-
tribution upon the flap hinge moments would probably
be similar to that due to the direct effects of slip-
stream displacement and, consequently, the basic
theory would not be greatly altered.

The preceding considerations indicate that dif-
ferential flaps are fundamentally a linked-aileron
system drooped to some initial downward setting and
having an upfloating tendency. Aileron-linkage theory
(reference 1) shows that the operating moments of such







0 IIACA ARR No. L5F25

a system can be reduced if the ailerons, starting from
equal deflections, are so linked that the upgoing aileron
deflects at a progressively greater rate than the
downgoing aileron (that is, when the values of the flap-
differential ratio d6fU/d6fD are greater than unity).
The effectiveness of a differential flap system in
producing rolling moments opposing those created by
power can therefore be increased by increasing the
differential ratio of the system above unity, because
such a change reduces the restoring moments of the
system and thus results in greater incremental flap
deflections for a given hinge-moment change induced by
power effects. At some differential ratio, the flap
system is neutrally balanced and has no tendency to
return to the original condition of equilibrium.
Differential ratios greater than this value create an
overbalanced flap system.


APPARATUS

Wind Tunnel


The tests were conducted in the Langley free-
flight tunnel; a complete description of the tunnel
is given in reference 2. The free-flight-tunnel six-
component balance used in the force tests is described
in reference 5. Figure 2 shows the test model mounted
on the balance strut in a yawed attitude. All force
and moment measurements obtained from this balance
are with respect to stability axes. The stability
axes (see fig. 3) are a system of axes having their
origin at the center of gravity of the airplane and
in which the Z-axis is in the plane of symmetry of the
airplane and is perpendicular to the relative wind, the
X-axis is in the plane of symmetry and perpendicular
to the Z-axis, and the Y-axis is perpendicular to the
plane of symmetry.


Trim Stand

Most of the tests were made on a trim stand that
was so constructed as to allow the model freedom in roll
and yaw about the stability axes. The construction of
the stand is illustrated in the sketch shown as figure 4..








NACA ARR No. L5F25 7

A photograph of the model mounted on the trim stand is
shown in figure 5- As shown in figure 4, a calibrated
spring was attached to the roll-free bearing for the
tests to provide for stability in roll and to permit
unbalanced rolling moments to be obtained as a function
of the angle of bank. The angle of bank was read
visually by means of the calibrated indicator card
shown in figure 5. Flap deflections were also read
directly from an indicator card by means of a pointer
rigidly attached to the inboard end of the left-flap
segment. (See fig. 5.)


Model

The model used in the investigation is generally
representative of low-wing radial-engine fighter air-
1
planes and corresponds to a ---scale model of a 40-foot-
10
span airolane. A three-view drawing of the model is
shown as figure 6 and photographs of the model are shown
in figure 7. The dimensional characteristics of the full-
scale airplane as represented by the ---scale model tested
10
in the Langley free-flight tunnel are as follows:

Propeller:
Diameter, feet . .. 11.7
Number of blades . . 2

Wing:
Area, square feet . 266.5
Span, feet . . .. 0
Aspect ratio . . 5.71
Airfoil section . Rhode St. Genese 55
Incidence at root, degrees .. 0
Dihedral, degrees . . 0
Sweepback at quarter percent chord line, degrees 5.2
Taper ratio . . 2:1
Mean aerodynamic chord, inches ... 85.90
Root chord, inches . 107.80

Center of gravity:
Back of leading edge of root chord, inches 32.91
Below fuselage center line, inches 0
Percent of mean aerodynamic chord . 25










Flaps:
Type . .. .
Span, feet .
Percent wing span .

Tail:
Vertical tail 1
Total area, square feet
Percent wing area .

Vertical tail 2
Total area, square feet
Percent wing area .

Vertical tail 5
Total area, square feet
Percent wing area .


17ACA ARR No. L5F25



. Split, partial span
. .... 20
. . 50


. ... .3
. . 5


. . 26.68
. . 10


. 40.0
. . 15


The model was equipped with a 14.0-inch diameter,
two-blade propeller set at an angle of pitch of 100 at
0.75 radius and was powered by a direct-current
1
controllable-speed electric motor rated 1- horsepower
at 12,000 rpm. The propeller was attached to the motor
by direct drive, and an electrical tachometer was
installed on the motor to permit direct measurements
of propeller speed. Right-hand propeller rotation was
used for all tests. The layout of the isolated nower
unit mounted on the roll bracket is shown in figure 8.

The model was equipped with partial-span split
flaps of 25 percent chord and of total span 50 percent
of the wing span. The flaps when locked and not in
differential operation were at an initial setting of 40o
The right- and left-wing flaps were linked together
through a differential linkage located in the fuselage.
Details of the flap linkage are shown in the photograph
presented in figure 9. This linkage system was designed
to permit variation of the flap-differential ratio
(ratio of upgoing flap deflection to downgoing flap
deflection). This result was accomplished by moving
the two end pivots of the singletree rearward with
respect to the fixed central pivot. The linl:age system
was so arranged as to permit rmxirnum differential flap
deflections of 37.5 down and 250 up from the initial
flap setting of l,00 down.


* *








NACA ARR No. L5F25


The Rhode St. Genese 35 airfoil section was used
on the model wing because of the high maximum lift coef-
ficient of this section at the low Reynolds numbers at
which the tests were run. The geometric dihedral of
the uodel measured from the lower surface was set at 00
for all tests.

No horizontal tail surfaces were used on the model.
Three similar vertical tail surfaces of different areas
were installed on the model for some tests. Sketches of
these tail surfaces are shown in figure 10.


TESTS

Test Conditions


All tests were run at a dynamic pressure of
1.90 pounds per square foot, which corresponds to an
airspeed of about 27 miles ner hour at standard sea-level
conditions and to a test Reynolds number of 172,000
based on the mean aerodynamic chord of 0.67 foot. All
forces and moments measured in the tests are with respect
to the stability axes (fig. 5), which intersect at a
point located at 25 percent mean aerodynamic chord and
on the center line (thrust line) of the fuselage. In
order to obtain sizeable power effects upon the effective
dihedral all power-on tests were made at Tc = 0.96,
a value that represented the maximum thrust obtainable
from the motor-propeller unit. This value simulated
full-scale brake horsepowers ranging from approximately
3000 to 9000 over the high-lift range.


Force Tests

Force tests were made with nower on and with pro-
peller off, and with flaps undeflected and deflected L0
for various angles of attack and yaw. Some tests were
made with power on at angles of attack of 1.00 and 15.50
and at yaw angles of 0 and 100 to determine the
effectiveness of the flap system. For these tests tha
flap on the left wing was locked at 00 and the deflection
of the flap on the right wing was varied from 00 to 700.

A complete thrust calibration of the propeller-motor
unit was made to determine the model power characteristics.







IIACA ARR No. L5F25


A plot of the results of this calibration is presented
as figure 11.


Trim-Stand Tests

Trim-stand tests were made to determine the effect
of freeing the differential flaps on the effective
dihedral of the test model. The influence of flap-
differential ratio unon the effective-dihedral char-
acteristics was also studied. Other tests were made
to determine the effect of vertical-tail area upon
the effective dihedral and the influence of differential
flap action upon the directional stability.

Test orocedure.- The effective-dihedral charac-
teristics of the model with tail surfaces removed were
determined as follows:

The model was set at various angles of yaw on the
test stand and the corresponding trim angles of bank for
the propeller-off and power-on conditions were noted by
visual observation. The values of the angles of bank
thus obtained were converted to rolling-moment coef-
ficients by means of the roll-spring calibration. The
effective-dihedral parameter CL, (or -CL) was then
directly determined from a plot of these rolling-'ioment
coefficients against the corresponding angles of yaw.
The same procedure for determining the effective-dihedral
characteristics of the model with vertical tail surfaces
installed was followed except that the model was free
in yaw and was trimmed at the different angles of yaw
by rudder deflection.

Calibration curves were obtained for each flap
linkage by measuring the upgoing flap deflection
produced by a given downgoing deflection on the opposite
wing. Representative calibration curves obtained in this
manner are shown in figure 12. It should be noted that
these curves are nonlinear. This nonlinearity is char-
acteristic of the linkage system employed and results
in a change of flap-differential ratio d6fu/d6fD with
incremental flap deflection. For definiteness, the term
"flap-differential ratio" was defined by the mean slope
of the differential curves over the first 200 of incre-
mental up deflection. This procedure is considered







NACA ARR HIo. L5F25


sufficient to identify the general effects of altering
flap-differential ratio in the te3ts.

Scope of trim-stand tests.- Trim-stand tests were
made to determine the dihedral characteristics of the
model for the following conditions:

(1) Flaps locked at 400; propeller off; a = 1.00;
CL = 095

(2) Flaps locked at 400; T = 0.96; a = 1.00 and
15.50; CL = 1.4 and2.7

(5) Flaos free; T = 0.96; a = 1.0 and 15.50;
CL = 1.4 and 2.7

Propeller-off tests were not run at 15.50 angle of attack
because the model was completely stalled at that angle.
The effect of varying the flap-differential ratio
between 0.8 and 1.4 was studied for condition (5). The
effect of vertical-tail area upon the nower-on
(Tc = 0.96) dihedral characteristics was investigated
at an angle of attack of 15.50 with flaps locked and with
flaps free at a differential ratio of 1.0.


RESULTS AND DISCUSSION


The results of the tests are given in figures 15
to 25. Lift and drag data obtained from the force tests
are given in figures 15 to 15 for various power and flap
configurations. These data show that maximum lift coef-
ficients comparable with those obtained on full-scale
airplanes were obtained for all test conditions.


Effective-Dihedral Characteristics

Effect of power.- The effect of power application
upon the effective-dihedral characteristics of the model
with flaps locked at 400 is shown in figure 16. These
data show that application of power increased the
negative slope of the curve of rolling-moment against
yaw angle from -0.00063 to -0.00205. This change
corresponds to reduction of about 7 in effective
dihedral and illustrates the usual effect of power
upon the dihedral parameter.







NACA ARR No. L5F25


Because of the particular geometric configuration
used in the tests, the test model possessed 20 negative
effective dihedral in the prooeller-off condition. The
model differed therefore from conventional full-scale
airplanes, which generally possess moderately large
positive effective dihedral in the propeller-off condi-
tion. This difference, however, is only of academic
interest inasmuch as the reduction in effective dihedral
caused by power is an incremental effect that is con-
sidered independent of the initial value of dihedral in
the propeller-off condition.

The data of figure 16 also indicate little effect
of lift coefficient upon the dihedral characteristics
of the model under conditions of constant thrust coef-
ficient. This phenomenon is unusual inasmuch as an
increase in lift coefficient generally results in an
increase in power effects. The increase in effective
dihedral generally associated with increasing angle
of attack (lift coefficient) was probably sufficient at
high angles of attack to offset the increased power
effects caused by the same increase in angle of attack.

The action of the slipstream in producing dihedral
changes, previously discussed in the section "Theory of
Differential Flap Action," appears to be verified by the
results of force tests made to determine the flap
effectiveness with power on (figs. 17 and 18). These
data show that when the model was yawed to the right
(* = 100) the lift increments contributed by the flap
on the right (trailing) wing were considerably increased
because of the action of the displaced slipstream. The
reverse was true when the model was yawed to the left
( = -100).

Effects of differential flap action.- Trim-stand-
test data showing the effect of freeing the differential
flaps on the effective dihedral are given in figure 19.
Figure 19(a) shows little effect with propeller off;
whereas figure 19(b) shows that with power on for a
differential ratio of 1.0 (equal up and down flap.
deflections), the negative dihedral change caused by
power at a lift coefficient of 1. was reduced by over
80 percent. A similar effect of differential flap
action was also attained at a lift coefficient of 2.7.
Additional data showing the effect of varying
flap-differential ratio are shown in figure 20. The
slopes of these curves, which are indicative of the








NACA ARR Ho. L5F25


effective-dihedral characteristics, are shown plotted
against mean flap-differential ratio in figure 21. The
results presented in figure 21 indicate that, although
decreasing the differential ratio below unity slightly
reduced the efficacy of the differential flaps in
opposing dihedral changes due to nower, increasing the
flap-differential ratio above unity was beneficial.
The data show that the adverse effects of power upon
the effective dihedral were completely eliminated when
a differential ratio of about 1.0 was used at a lift
coefficient of 1.4 or when a differential ratio of
about 0.96 was used at a lift coefficient of 2.7. Use
of ratios greater than these values reversed the effect
of power and resulted in positive increases in the
effective-dihedral parameter with power application.
Slightly larger effects of differential ratio were
usually encountered at the high-lift condition
(CL = 2.7). These differential-ratio tests showed that
the differential flap system employed in the tests
became overbalanced at differential ratios of about 1.4.
When overbalance occurred, the flans locked violently
against their stops as soon as power was applied, thereby
inducing large and abrupt rolling motions. This action
occurred for all angles of yaw.

The beneficial effect of increasing flap-differential
ratio is in agreement with the theory of reference 1.
As shown by the data in figure 22, increasing flap-
differential ratio generally resulted in greater incre-
mental flap deflections at a given angle of yaw as a
result of reduced unbalance of the flap system.

Effect of vertical-tail area.- Results of tests
made to determine the influence of vertical-tail area
upon the power-on effective-dihedral characteristics
of the model are presented in figure 25 and are
summarized in ,figure 24. These data show that adding
vertical-tail area up to 15 percent of the wing area
had little effect. The general tendency of such addi-
tions, however, was to increase the effective dihedral.


Directional Stability Characteristics

Rudder-deflection data from the yaw-free tests
(fig. 25) indicate that differential flap action con-
siderably increased the static directional stability.
This increase in stability is attributed to the drag







NACA ARR No. L5F25


changes accompanying the differential action. As the
airplane is yawed, the flap on the trailing wing moves
up and reduces the drag of that wing, whereas the flap
on the leading wing moves down and increases the drag.
These drag changes produce stabilizing yawing moments.
Although no yaw-free tests were made at differential
ratios other than 1.0, it appears reasonable to assume
that increasing the differential ratio will increase
the stabilizing action of the flaps in yaw because the
larger flap deflections encountered will cause greater
drag increments.


Remarks about Design

Dynamic response.- In brief tests made to determine
the dynamic response of the flap system to sudden and
sharp yawing motions, the results (obtained by visual
observation) indicated no appreciable lag of flap
deflection with yawing motion. It should be emphasized,
however, that the friction in the flap system in the
current tests was held to small values, perhaps smaller
than those encountered in full-scale designs. Inasmuch
as excessive friction in the flap system could cause
the controls to "freeze" in a differential attitude
(particularly for small degrees of balance) and thus
to induce violent rolling maneuvers, the designer should
attempt to limit the friction in the flap system to as
small a value as is practical.

Linkage design.- The principle of the differential
flap linkage is not necessarily restricted to the
mechanical, pin-jointed type of system. Although
differential ratios other than unity are readily
obtained with this type of system, differential flaps
could also be linked by means of hydraulic, cam, gearing,
or electrical systems.

Aerodynamic balance.- Because of the severity of the
rolling motions caused by an overbalanced flap system,
care should be taken in the design of differential flap
systems to allow a safe margin of unbalance.. Further
analytical and experimental work with particular refer-
ence to the effects of nonlinearity of flap load and
moment characteristics is required to establish a
quantitative design procedure for differential flap
systems.








NACA ARR ITo. L5F25


C NCLUSTONS


The following conclusions have been drawn from
tests of a differential flap system installed in a
--scale powered airplane model in the Langley free-
10
flight tunnel:

1. A flap system in which the right and left flaps
were linked together and were free to deflect differ-
entially materially reduced or eliminated negative
dihedral changes caused by power application.

2. Increasing the flap-differential ratio of the
flap system (ratio of upgoing flap deflection to down-
going flap deflection) above unity increased the effec-
tiveness of the flap system in opposing dihedral changes
caused by power.

5. Increasing the flap-differential ratio of the
flap system reduced and eventually reversed che aero-
dynamic tendency of the flaps to restore themselves to
their initial setting of equal deflection.

4. Little effect upon the effective-dihedral char-
acteristics was observed when the flap-differential
ratio was decreased to a value below unity.

5. Differential flap action caused an increase
in the static directional stability of an airplane.

6. Further analytical and experimental study is
required to develop a quantitative design procedure for
differential flap systems.

Langley 1'emorial Aeronautical Laboratory
National Advisory Committee for Aeronautics
Langley Field, Va.






NACA ARR No. L5F25


REFERENCES


1. Jones, Robert T., and Nerken, Albert I.: The Reduc-
tion of Aileron Operating Force by Differential
Linkage. NACA TN No. 586, 1936.

2. Shortal, Joseph A., and Osterhout, Clayton J.:
Preliminary Stability and Control Tests in the
NACA Free-Flight Wind Tunnel and Correlation with
Full-Scale Flight Tests. NACA TN No. 810, 19)1.

3. Shortal, Joseph A., and Draper, John W.: Free-
Flight-Tunnel Investigation of the Effect of the
Fuselage Length and the Aspect Ratio and Size of
the Vertical Tail on the Lateral Stability and
Control. NACA ARR :To. 3D17, 1945.







NACA ARR No. L5F25


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direc/on


NATIONAL ADVISORY
rZ COMMITTEE FOR AERONAUTICS


Fiqure 3. Syslern of /qbl////y axes. Arrows Irnt/ca/e
posi/wve direc-llons of momnen/3 nd forces.


Fig. 5






NACA ARP No. L5F25 Fig. 4




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NATIONAL ADVISORY
COMMITTEE FM AERONAUTICS


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Fig. 6







NACA ARR No. L5F25


Figure 7,- Photographs of powered model employed
in Langley free-flight-tunnel tests.


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NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS
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Fig. 11






NACA ARR No. L5F25


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NATIONco ADVISORY
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NATIONAL ADVISORY
COMMITTEE FOE AERONAUTICS
0 (3 /6 24 32 40
11pgoing Ai/ap ru7'7en/lafon;

F/qgL# /2. ReynlolIV&e ca/,lbro/,on curves
used lo de/erm/ne /;lap c/lferen/la/ ra/l/o
on powered Aesl model. 0 equo/s 40 in//ol
de fe c //on .


Fig. 12





NACA ARR No. L5F25


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NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS
_TIONAL A _O~


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F/'gure /3.-Lift alnd draq choroccer/st/c.s
Pr _ol/ered t-es mctrde/ npo/oy9o0 n d
Iets of /llT/erlentlo/ a /TI sysy ms.
Prope//er'-of coo lf/on; Y- -O 9 190 pounds
per square AVI/.


Fig. 13







NACA ARR No. L5F25


2.2


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Angle of oa2'/c* oc eg


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Fig. 14





NACA ARR No. L5F25


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NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


4 8 12 /6
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figure /5.- L/1 QoWY' g39 chA2rQcter/st/Cs Oc ao
/fx,4tqeQ "aest mofe/ eorpbyec/ /) reas
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10


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Fig. 15


1.4






NACA ARR No. L5F25


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NACA ARR No. L5F25


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NACA ARR No. L5F25


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Fig. 18





NACA ARR No. L5F25


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NACA ARP No. L5F25


o- X/ops locked a/o 40
- -/- FIo s. free
(differenhal ft/ao -/0)
- -- Prope//e, off; -f/ops
/oc/

NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS
-8 -4 0 4 8
Angle of yaW) V/), deg

(T) = 0 L96 (uIn/e55 / c- ermse rdwca/ed).
FrQure /9. nc2r/aoed.


Fig. 19b






Fig. 20a


NACA ARR No. L5F25


1




NACA ARP No. L5F25


Fig. 20b


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NACA ARR Nlo. L5F25


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Fig. 22a





NACA ARR No. L5F25


56








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Figure 22 .- Concluded.


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S/ / NATIONAL ADVISORY
Z COMMITTEE FOII AERONAUTICS
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77 / / /// l /lll Ill II l///7 //l/17 li// I/I///////I / / /I- T


Fig. 22b






NACA ARR No. L5F25


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NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS

Angle l .y.ov t odg

Fjure ?23. Effect o" ker-ll/cfl 7U11/ orea on m1e
r d//ro-moment chur cter/s3Icz oa a. 'o-vereCo
-est /nod/ In yo 7c =O.96; Cl --.7:
q/=.9 ftoofndJ per aquare fot


Fig. 23






NACA ARR No. L5F25


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NACA ARR No. L5F25


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