Langley full-scale-tunnel tests of the custer channel wing airplane

MISSING IMAGE

Material Information

Title:
Langley full-scale-tunnel tests of the custer channel wing airplane
Series Title:
NACA RM
Physical Description:
57 p. : ill. ; 28 cm.
Language:
English
Creator:
Pasamanick, Jerome
Langley Aeronautical Laboratory
United States -- National Advisory Committee for Aeronautics
Publisher:
NACA
Place of Publication:
Washington, D.C
Publication Date:

Subjects

Subjects / Keywords:
Propellers, Aerial   ( lcsh )
Aerodynamics -- Research   ( lcsh )
Genre:
federal government publication   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Summary:
Summary: As a part of a research program to study the principles involved in the use of propeller slipstreams and jets to increase lift, an experimental Custer Channel Wing airplane has been tested in the Langley full-scale tunnel to investigate the lift characteristics of a channel-propeller combination and the flow phenomena in and about a channel wing. Some of the general stability and control characteristics of the airplane were also studied at tunnel airspeeds from approximately 25 to 40 mph. Emphasis was placed on determining the airplane static life characteristics (zero airspeed) for the basic configuration and for several modifications. The effect of a ground boundary on the airplane static characteristics was investigated by testing the airplane both on the ground and out of the influence of the ground. Photographs of the tuft surveys made to determine the air-flow distribution around the channel and tail surfaces are included.
Statement of Responsibility:
by Jerome Pasamanick.
General Note:
"Report date April 7, 1953."

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003808518
oclc - 129739157
sobekcm - AA00006172_00001
System ID:
AA00006172:00001

Full Text


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NACA RM L53A09

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

RESEARCH MEMORANDUM

LANGLEY FULL-SCALE-TUNNEL TESTS OF THE

CUSTER CHANNEL WING AIRPLANE

By Jerome Pasamanick


SUMMARY


As a part of a research program to study the principles involved in
the use of propeller slipstreams and jets to increase lift, an experimental
Custer Channel Wing airplane has been tested in the Langley full-scale
tunnel to investigate the lift characteristics of a channel-propeller com-
bination and the flow phenomena in and about a channel wing. Some of the
general stability and control characteristics of the airplane were also
studied at tunnel airspeeds from approximately 25 to 40 mph. Emphasis
was placed on determining the airplane static lift characteristics (zero
airspeed) for the basic configuration and for several modifications.
The effect of a ground boundary on the airplane static characteristics
was investigated by testing the airplane both on the ground and out of
the influence of the ground. Photographs of the tuft surveys made to
determine the air-flow distribution around the channel and tail surfaces
are included.

The significant findings of the static tests are summarized briefly
as follows:

(a) The channel-propeller configuration in the static condition with-
out ground effect produced a resultant force inclined 230 upward from the
propeller thrust axis.

(b) The magnitude of the static resultant force with the propellers
operating at about 2,450 rpm (170 horsepower total for both propellers)
was approximately 88 percent of the static thrust calculated for these
propellers when not in the presence of the channels.

(c) The airplane, having a normal tail configuration, had neither
trim nor control effectiveness at zero forward speeds.


"? : c







NACA RM L53A09


INTRODUCTION


There is renewed interest in airplanes capable of hovering and flying
at very low airspeeds, as well as attaining satisfactory performance at
cruising and high speeds. In this connection the NACA has started a basic
research program to study the principles involved in the use of propeller
slipstreams and blowing jets to increase lift. As a part of this program
the NACA has undertaken tests of a full-scale experimental Custer Channel
Wing airplane to study the static and low-speed lift-producing capabili-
ties of the channel-wing principle as well as to study some of the sta-
bility and control characteristics of the experimental airplane at zero
airspeed and low forward speeds at high angles of attack.

The primary objective of this investigation was to determine the
amount of lift that would be produced by the operation of propellers
located at the rear of open wing channels of semicircular cross section
to induce a flow of air through the channels. The magnitude and direction
of the corresponding resultant force were determined for various power
conditions at zero airspeed. Also a brief investigation was made of the
static stability and control characteristics of the experimental airplane
at zero airspeed and at tunnel airspeeds of about 26 to 41 mph over an
angle-of-attack range of -2o to 460. A few of these tests were conducted
with the airplane yawed 50 and 100. The static tests were conducted with
the airplane mounted in the tunnel (tunnel inoperative) in the absence of
ground effect and also with the airplane on the ground in a three-point
attitude to investigate ground effect. The test conditions of propeller
operation, blade angle, and propeller-channel configuration were carefully
controlled to satisfy the conditions prescribed by the designer for the
experimental flight configuration.


SYMBOLS


The data are presented with respect to the stability axes which con-
stitute an orthogonal system of axes having the origin at the airplane
center of gravity, the Z-axis in the plane of symmetry and perpendicular
to the relative wind in the forward-flight tests or vertical in the static
tests, the X-axis in the plane of symmetry and perpendicular to the
Z-axis, and the Y-axis perpendicular to the plane of symmetry. The posi-
tive directions of the forces and moments measured about these axes are
shown in figure 1. The center-of-gravity location is shown in figure 2.

The data are presented in coefficient form for airspeeds of 26 mph
or greater. It is emphasized that the forces and moments created by pro-
peller operation at these low airspeeds produce large coefficients because
of the low dynamic pressure. This is particularly true of the lift coef-
ficient at high angles of attack because the propellers operating at high







NACA RM L53A09


thrust produce a force component in the lift direction in addition to
the wing lift and therefore produce lift coefficients that are very large
at low airspeeds and that are infinite when the airplane is at rest. For
this reason, actual forces and moments rather than coefficients are used
in presenting the data from the static tests or tests at airspeeds of
11.5 mph or lower. For the static tests where there is no wind, the
vertical component of the resultant force is called the lift and the
horizontal component along the X-axis is called the longitudinal force,
positive when directed forward with respect to the airplane.

CL lift coefficient, Lift where lift is the com-
q X Wing area
ponent of the resultant force perpendicular to the
relative wind and the wing area is the projected area
of the channel wings Sc, unless specified otherwise

Longitudinal force
CX longitudinal-force coefficient, Long, where
qSc
longitudinal force is the component of the resultant
force along the X-axis, positive upstream (When the
airplane is at zero yaw (or sideslip), CX = -CD where

CD is the drag coefficient rag. CX is positive when
qSc
the thrust component along the X-axis exceeds the drag
component.)
Lateral force
Cy lateral-force coefficient, where lateral
qSc
force is the force along the Y-axis, positive to the
right

Cm pitching-moment coefficient, M/qSec

C rolling-moment coefficient, L/qScb

Cn yawing-moment coefficient, N/qScb

D drag, Ib

M pitching moment about Y-axis, positive when nose is raised,
ft-lb

L rolling moment about X-axis, positive when right wing is
depressed, ft-lb

N yawing moment about Z-axis, positive when right wing is
retarded, ft-lb







NACA RM L53A09


R resultant force, (Lift)2 + (longitudinal force)2, Ib

q free-stream dynamic pressure, pV2, lb/sq ft
2P

qav average free-stream dynamic pressure during an angle-of-
attack run, pVav2, Ib/sq ft


p mass density of air, slugs/cu ft

V free-stream velocity, ft/sec unless specified otherwise

Vav average free-stream velocity during an angle-of-attack
run, ft/sec unless specified otherwise

S projected area of channel wings, 35 sq ft

St total area including the ailerons and the projected fuse-
lage area intercepted by the channels, 59 sq ft

c channel chord, measured in a plane parallel to plane of
symmetry, 35 in.

b airplane span, 20.25 ft

8e elevator control deflection, measured perpendicular to
hinge line, positive when trailing edge is down, deg

br rudder control deflection, measured perpendicular to hinge
line, positive when trailing edge is left, deg

a angle of attack of channel-wing-chord line (which is paral-
lel to the thrust axis) referred to relative air stream
in forward-flight tests and to the horizontal in static
tests, deg

P angle of sideslip, positive when left wing is retarded, deg

e inclination of resultant force, tan1 Lif deg
Longitudinal force


total horsepower input to motors







NACA RM L53A09


MODEL AND TESTS


The Custer Channel Wing airplane supplied by the Custer Channel
Wing Corporation was an experimental airplane in which Continental C-90
internal-combustion engines of approximately 90 horsepower driving 6-foot-
diameter metal propellers were mounted in two wing channels attached to
a light-plane airframe. The airplane as received with an uncovered fuse-
lage framework weighed approximately 900 pounds empty. The tail arrange-
ment of this configuration was a conventional airplane-tail installation.

For the tests reported herein, the airplane fuselage was covered
with fabric and the internal-combustion engines were replaced by variable-
speed electric motors because the original internal-combustion engines
were not adequately lubricated for operation at very high angles of attack.
The use of electric motors also permitted approximate measurements of the
power input. A drawing of the Custer Channel Wing airplane as tested is
presented as figure 2 and photographs showing the airplane mounted in the
tunnel and on the ground setup are given as figure 3.

The power input to the motors was determined from current-torque
calibrations of the individual electric motors and also checked by the
conventional determination of power by means of wattmeters. It is noted
here that the propeller-operating wind-tunnel and ground tests were con-
ducted under different motor control conditions. In the tunnel both
motors were operated at the same speed because only a single source of
variable-frequency supply and control was available, which resulted in a
maximum propeller speed of 2,450 rpm as dictated by the particular motor
drawing its maximum rated current. When the port motor was operated
alone, however, it was possible to attain a speed of 2,645 rpm before
reaching its limiting current rating. For the ground tests, on the other
hand, an additional power source was made available and it was possible
by special manipulation of the control equipment to run each motor inde-
pendently and even at overload conditions to attain 2,625 rpm, producing
220-horsepower total output for both motors as compared to about 170 horse-
-power at 2,450 rpm.

In the course of the tests in the tunnel a power difference of the
order of 15 to 20 horsepower existed between the two motors. Inasmuch
as both the right-hand (starboard) and left-hand (port) propellers were
carefully set at 11.50 blade angle at the 0.75-radius station, the power
asymmetry can be accounted for by a dissymmetry in either the geometry
of the two propellers or the channels, or perhaps both. Except for some
thrust dissymmetry which introduced lateral and directional trim shifts,
the power dissymmetry does not reflect a primary influence on the sig-
nificance of the lift results of this investigation.







NACA RM L53A09


The elevator and rudder control surfaces were remotely controlled
by electric actuators and the control-surface deflections were recorded
from calibrated control-position transmitters. The horizontal tail had
a fixed 100 nose-down incidence with respect to the channel-chord plane
and the ailerons were locked in a neutral position with respect to the
channel-chord plane.

Propeller-removed tests were conducted over the low airspeed range.
The propeller speed, blade angle, and axial location, with respect to
the trailing edge of the channel wing, were selected and carefully con-
trolled to agree with conditions formerly used by the manufacturer in
level-flight tests. Most of the tests were made with the center of the
propeller tip chord 3/4 inch inside the channel-wing trailing edge
(propeller tip trailing edge at the channel trailing edge) and with a
minimum of blade tip clearance (approximately 1/16 inch). Some tests
were conducted with the center of the propeller tip chord at the channel-
wing trailing edge and for these runs the blade tip clearance was no
more than 3/8 inch. In the latter part of the wind-tunnel test program,
several runs were made with an extensible leading-edge flap and an
extended trailing-edge flap installed on each channel, as shown in fig-
ure 2. It should be noted that the deflection angle, chord, and camber
of the extensible leading-edge flaps are not necessarily optimum configu-
rations for a simulated shroud contour but should at least approximate
the effect of a more generous leading-edge radius.

The airplane was tested over a range of angle of attack and airspeed
from about 2 to 41 mph in the wind tunnel at zero sideslip with pro-
pellers operating, and several tests were made at approximately -50
and -10O angles of sideslip for small negative and high positive angles
of attack. The elevator and rudder control effectiveness was obtained
for the forward-flight conditions at several angles of attack.

Static tests (zero airspeed) were made both in the wind tunnel where
the airplane was tested high enough above the tunnel ground board not to
be influenced by ground effect for most test conditions and also on the
hangar floor in a three-point attitude. The static tests in the absence
of ground effect were conducted with the tunnel entrance nozzle blocked
off with a tarpaulin in order to insure essentially zero tunnel airspeed
(the airplane slipstreams would otherwise create an appreciable circula-
tion in the tunnel). Long hanging tufts indicated undisturbed air ahead
of and below the channels, which effectively represented a zero-velocity
free-air condition. The ground tests of the airplane were conducted in
the Langley full-scale-tunnel hangar with the entire door system opened
to eliminate interference on the propeller slipstream. This location
enabled the airplane to be located in an area free from natural wind.
The airplane was mounted about 2 inches from the ground on a floating
frame suspended from Baldwin SR-4 load cells which measured the lift and
the longitudinal force at the desired test conditions. The attitude of







NACA RM L53A09


this arrangement simulated the three-point attitude (am 190) of the air-
plane resting on the ground.

In addition to the tests with both propellers operating or removed,
some tests were made to determine the effect of both propellers wind-
milling at an average airspeed of approximately 26 mph as well as the
effect of asymmetric power at about the same forward speed and at static
conditions in the absence of a ground boundary. Some questions arose as
to whether the airplane channel-flow characteristics might have been
affected by replacing the engines with the smaller electric motors.
Accordingly, a few additional ground tests were made with a nacelle,
roughly representing the engine, built around the electric motors. (See
fig. 4.) The modifications reduced the open frontal area of the channel
to a value slightly less than that with the original internal-combustion
engines.

CORRECTIONS AND RESULTS


The wind-tunnel forward-flight data have been corrected for jet-
boundary and stream alinement effects. The tunnel blocking and support
interference were sufficiently small that corrections were not applied.
The corrected wind-tunnel forward-flight data are determined as follows:


orr = tun 0.248CL 0.5
corr tun L

CLorr = CL 0.0087CX


Cxorr = Cx + 0.0087CL + 0.0043CL2 0.0012
corr

The figures showing the basic data are presented in the following
order: The effects of propeller rotational speed, angle of attack, and
control-surface deflection on the force and stall characteristics of the
airplane in static or near-static conditions as determined in the absence
of ground effect are presented in figures 5 to 8. Figures 9, 10, and 11
show the effect of the horizontal tail on the static force and flow
characteristics of the airplane with and without the modified nacelle in
the presence of a ground boundary. The basic aerodynamic characteristics
at low forward velocities are given in figures 12 and 13 for the propellers
inoperative and operative, respectively. Results of tuft observation
studies of the flow behavior at low forward velocities are shown in fig-
ure 14. The effects of control-surface deflection on the model charac-
teristics for low-speed conditions are presented in figures 15, 16,
and 17 for the rated, windmilling, and asymmetric power conditions,
respectively. The airplane characteristics in sideslip are given in fig-
ure 18 and the effects of the presence of the ailerons on the airplane
longitudinal characteristics are given in figure 19.







NACA RM L53AO9


DISCUSSION OF RESULTS

Longitudinal Characteristics


Static tests in the absence of ground effect.- The static lift,
longitudinal force, and pitching moment of the airplane as obtained in
the wind tunnel for the propeller located with the center of its tip-chord
trailing edge 3/4 inch inside the trailing edge of the channel wing and
operating at 2,450 rpm were 340 pounds, 800 pounds, and -350 foot-pounds,
nose down, respectively, at &-= 00 (fig. 5(a)). The corresponding
resultant force and inclination of the resultant-force vector with respect
to a were 880 pounds and 230 (fig. 7). The nose-down pitching moment
would require a tail force in the negative lift direction for trim in
hovering which would further reduce the magnitude and inclination of the
resultant-force vector. Thus, provided a suitable tail could be obtained
for the airplane to provide the negative tail force, the airplane, in
order to hover, would have to be inclined at some angle greater than 670
and the weight would have to be less in magnitude than the resultant force.

There is a small reduction in the magnitude of the resultant force
with angle of attack which would not be expected if power were constant
and if the slipstream and flow about the channels experienced free-air
conditions. Although there is a small power decrease, it is reasoned
that the decrease in the lift force, which at a = 460 is about
40 pounds, is caused by a certain amount of ground effect on the flow
about the airplane in the tunnel when the airplane is at very high
angles of attack. (See fig. 3(a).)

The results presented in figure 8 show practically a constant
pitching moment for all elevator deflections (zero elevator effectiveness)
which indicates that the propeller slipstream passed completely below the
tail and, therefore, the measured airplane static lift and longitudinal-
force characteristics in the absence of ground effect were dependent only
upon the wing and propeller characteristics. It should be noted in fig-
ure 8(c) that the pitching-moment curve indicates that there was some
flow in the neighborhood of the tail at 460. Reference to figure 3 again
shows that for this high angle the tunnel ground board was probably not
far enough below the tail to provide completely free-stream conditions.
A visual tuft survey in the region behind the propeller plane at zero
angle of attack showed that the propeller slipstream for the static
flight condition was deflected well downward and underneath the tail.
These results show that in the hovering or very low speed flight condi-
tions, a problem exists with regard to obtaining longitudinal and direc-
tional control.







NACA RM L53A09


As mentioned previously, there was a power dissymmetry that occurred
in these tests which, in the static condition at m = 00 and at 2,450 rpm,
caused rolling moments and yawing moments corresponding roughly to
50 pounds more lift and about 35 pounds more thrust on the starboard side
than on the port side of the airplane (fig. 8). No basic significance is
attached to this particular condition of test since, in flight, the pilot
would be expected to control the thrust adequately by variable propeller
pitch or by throttle adjustments.

Exploratory studies made early in the test program indicate that
the selected locations of the propeller in tne channel did not affect
the static lift characteristics of the airplane (fig. 5(a)). The channel
and propeller-blade deformations that occurred at the maximum propeller
rotational speed (2,450 rpm) actually resulted in zero tip clearance over
a small portion of the channel for each of the two propeller locations
described previously. At 2,000 rpm, however, the data show approximately
20 pounds higher lift for the more forward propeller location having the
nominal tip clearance of 1/16 inch. Visual flow studies, figure 6, indi-
cate that shortly ahead of the propeller plane the flow along the bottom
of the channel was steady but a region of stalled flow occurred along
the sides of the channel. It was evident from the behavior of the tufts
forward on the channel that leading-edge separation extended along the
bottom of the channel and along the outboard side of the channel (down-
ward rotation of the propeller). The fuselage fillet provided a more
gradual inflow into the channel and eliminated the leading-edge stall
along the inboard side of the channel. The channel leading-edge stall
for this static condition is attributed to the high inflow angle in spite
of the fairly favorable leading-edge radius and camber of the airfoil
(NACA 4412). Inasmuch as an effective control of inflow into a fully
shrouded propeller has been found to require a much more generous leading-
edge radius than that used on this channel, an attempt was made to deter-
mine how much relief of the leading-edge stall would be possible by the
addition of an extensible leading-edge flap. The results presented in
figure 5(b) show that this particular application of an extensible
leading-edge flap did not favorably affect the airplane characteristics,
but the addition of an extended trailing-edge flap increased the lift by
about 25 pounds, which would be expected from a more downward deflection
of the slipstream. Any such attempt to increase the static lifting
capabilities of the channel wing, either by a change in airfoil section
to one having larger leading-edge radius and camber or by an external
flap-type device, might not be compatible with good performance at the
higher-speed forward-flight condition.

Effect of low airspeed.- The effects of a small forward wind velocity
on the airplane characteristics at zero angle of sideslip with only the
port motor operating are shown in figure 5(c). It can be seen from this
figure that a wind velocity of 3 mph increases the lift by 20 pounds at
2,450 rpm and a velocity of 4 mph increases the lift by 30 pounds at







10 NACA RM L53A09


2,625 rpm. The increased lift resulted, in part, from the stabilization
of the flow over the wing leading edge and into the channel. An example
of the effects of a forward wind velocity on the lift and thrust capa-
bilities of the airplane at several angles of attack with both propellers
operating at 2,450 rpm is shown in the following table:

a = 0 a = 200 a = 46
V,
mph Lift, Longitudinal Lift, Longitudinal Lift, Longitudinal
lb force, lb lb force, Ib Ib force, Ib

0 340- 800 580 635 770 350
4 360 795 650 600 840 280
11.5 385 745 735 540 980 190
26 470 590 940 395 1375 -210

At the normal three-point attitude of the airplane (a s 200) wind veloci-
ties of 4 mph and 11.5 mph increased the airplane total lift by 70
and 155 pounds, respectively, over that obtained in the static condition.
It is considered that the improved leading-edge flow and more uniform
flow into the propeller disk are at least partly responsible for this
increased lift.

It is necessary to state that these tabulated lift results are for
out-of-trim conditions and therefore can only be considered in approxi-
mate relationship. The lift requirement of the horizontal tail for trim
would tend to decrease somewhat the total lift values shown at zero for-
ward speed and to increase somewhat the total lift at a forward speed of
26 mph.

Static tests in presence of ground effect.- The lift characteristics of
the airplane in the presence of the ground were appreciably influenced by
the horizontal tail. (See fig. 9.) The data show that, with the tail off
on the ground-test arrangement, the same lift and horizontal forces were
obtained as with the tail on in the tunnel tests, which verifies that the
wing flow and propeller flow are relatively unaffected by the presence
of the ground and agrees with the previously mentioned observation that,
in the tunnel tests at zero airspeed, the slipstream was below the tail.
At the ground angle of 190 and a propeller speed of 2,450 rpm the total
lift of the channel wing and vertical component of the propeller thrust
(with Se = 00) is about 690 pounds. Increasing the propeller rotational
speed results in forces proportional to the square of the speed and horse-
power proportional to the cube of the speed, which thereby verifies that
no appreciable extraneous winds were present. Deflecting the elevator
control down 200 resulted in an additional increase in lift of about
100 pounds. The removal of the horizontal tail reduced the lift of the
airplane 104 and 135 pounds for the propeller operating at 2,450
and 2,625 rpm, respectively. The measured power input for the propeller






NACA Ri.: L53A09


rotational speeds of 2,450 and 2,625 rpm were approximately 170
and 220 horsepower, respectively.
In order to visualize the air flow in the vicinity of the channel
and to help explain the effectiveness of the tail in ground effect, the
photographs presented in figure 10 show the flow pattern at various
locations of the ground-test arrangement, and the following sketch
illustrates the general flow pattern into the plane of the propeller.
These flow characteristics were also present for the wind-tunnel arrange-
ment where, however, it was not possible to survey the flow as completely
as was done in the ground tests.








-3^


SGround plane



The air flow over the channel leading edge and extending approximately
2 inches above the channel surface is quite rough.

In the region ahead of and above the propeller hub axis and espe-
cially near the top of the propeller plane (fig. 10(a)), the inflow angle
into the disk is very high (estimated to be about 750). In fact, at the
top of the propeller disk (fig. 10(b)) the air is seen to flow into the
propeller from the downstream face. Inasmuch as the propeller is
operating in nonuniform flow, its loading is asymmetrical and its effi-
ciency, correspondingly, is lowered. A few preliminary surveys of the
propeller slipstream immediately behind the plane of rotation indicated
that the upper half, on the average, exhibited about one-half the dynamic-
pressure rise of the lower half. It appeared that the outer 50-percent
radius of the propeller in the lower 1200 sector was the most effective
portion of the propeller. The flow asymmetry observed is probably the
major cause for the large difference between the measured longitudinal
force of about 400 pounds per propeller and the calculated thrust of
about 500 pounds for the same blade angle and tip speed. The acute down-
ward deflection of the propeller slipstream directly behind and on the
center line of the propeller plane and the large positive angle of upwash







NACA RM L53A09


in the region of the horizontal stabilizer can also be seen from the
photographs in figure 10. It was evident that the horizontal tail was at
a positive angle of attack in a flow field of fairly high dynamic pres-
sure. The lift of the horizontal tail (fig. 9) was measured to be 15 per-
cent and 25 percent of the total airplane lift with the elevator neutral
and set at 200 (trailing edge down), respectively, for a propeller rota-
tional speed of 2,450 rpm.


As was previously stated, the horizontal tail on the experimental
airplane tested did not contribute at all to the airplane static lift
when tested in the absence of ground effect in the wind tunnel. A con-
cise comparison of the results with and without ground effect can be
determined from table I. The ground effects on the present configuration
are of such major importance that considerable study would have to be
given to the problem and especially to the controls in attempting to
design such an airplane for vertical ascent at take-off. In the absence
of ground effect the airplane tested has a large nose-down pitching
moment (about -350 foot-pounds) and, in order to trim the airplane
longitudinally, the tail would be required to produce a down force of
approximately 30 pounds, so that a net reduction of lift results.
Accordingly, the large lift increments observed on the ground are of no
particular significance since a negative tail lift is required for trim
in flight.

Effect of nacelle installation.- The electric-motor-nacelle instal-
lation used in the present investigation resulted in a greater open-
channel frontal area (measured in a plane parallel to the plane of rota-
tion) than that for the internal-combustion-engine installation of the
experimental flight airplane. In order to determine the effects of
channel blockage on the propeller and channel static characteristics, a
modified nacelle installation was tested on the ground-test arrangement.
The frontal area blocked by the modified engine-nacelle mock-up was
approximately two and one-half times the frontal area blocked by the
electric-motor-nacelle installation used in the wind tunnel (see fig. 4)
and was slightly greater than the blockage by the internal-combustion
engines installed originally. The results at 2,450 rpm (figs. 9 and 11)
show that the modified nacelles reduced the airplane lift by 40 pounds
and also reduced the airplane longitudinal force by about 40 pounds.


Forward Flight

Characteristics with propellers removed.- The maximum lift coeffi-
cient (fig. 12) of the Custer Channel Wing airplane in the basic condi-
tion, propellers removed, was 2.2 and 2.1 for tunnel velocities of
approximately 26 and 40 mph, respectively. Actually, in any one run, the
tunnel dynamic pressures varied from about 2.2 pounds per square foot at
= 00 to 1.5 pounds per square foot at the highest angle of attack for
the low-airspeed data. For the higher-airspeed data, the tunnel dynamic







NACA RM L53A09


pressures varied from 4.4 to 3.9 pounds per square foot from the lowest
to the highest angle of attack. The small decrease in Cax (0.1) at
the higher forward airspeed is considered to be the result of some
increased roughness at the channel leading edge caused by the previous
installation and removal of the flap (the data for 40 mph were obtained
after the tests with the extensible leading-edge flaps were made). The
results (fig. 12) indicate that, with regard to C1, there was probably
some asymmetry near and beyond the stall, resulting in the irregularity
of the C, curve, and that most of this asymmetry appears to have dis-
appeared at the higher speed. The airplane had stable nose-down pitching
moments at the stall for both airspeeds, but there is an unstable pitching
tendency in the low CL range.

It should be emphasized at this point that in the reduction of data
the coefficients are based on the channel-wing area alone instead of the
total area which is the normal procedure. Such a presentation, however,
does not afford a direct comparison with other airplane configurations
since the forward-speed results are influenced by the ailerons and the
fuselage. In order to illustrate the point, the lift results in fig-
ures 12 and 13 are plotted also in coefficient form based on the total
effective wing area, which include the ailerons and the projected fuse-
lage area intercepted by the channels. It will be seen that the maximum
values of CL are comparable with those normally attained by conven-
tional airplanes with thick wings and without high-lift devices.

Characteristics with propellers operating.- Varying the longitudinal
location of the propeller plane in the channel had no significant effect
on the lift characteristics of the airplane in forward flight, as was
also found to be the case in static conditions. The rearward-located
propeller in the forward-flight condition, however, decreased the
longitudinal-force coefficients by approximately 0.3 and produced an
unstable pitching-moment shift throughout the lift range (fig. 13(a)).
Unless otherwise noted, the data of the following discussion were obtained
with the rearward-located propeller in the channel.

With the propellers operating at 2,450 rpm, increasing the forward
velocity resulted in more linear aerodynamic characteristics throughout
the angle-of-attack range (fig. 13). Propeller operation contributed a
destabilizing longitudinal-trim shift but apparently the stall remains
stable. The irregularities in lateral force and yawing moments, which
occurred at the lower speed (a ; 24, fig. 13(a)), were completely elimi-
nated at the higher speed (fig. 13(b)). The large variations of the air-
plane lateral characteristics at the lower speed result from asymmetrical
flow separation in the region just ahead of the propeller. The tuft sur-
veys of figure 14 show the flow to be rough at the channel trailing edge
for several degrees below the stall, followed by a complete and instantane-
ous flow breakdown on the lower part of the channel at a = 25.20. It is
of interest to note from figure 14 that the tufts indicated the direction







NACA RM L53A09


of flow on the channel surface to be inclined toward the lower part of
the channel. This drainage of the flow from the side surfaces delayed
the separation on these surfaces until higher angles of attack were
reached. The forward-flight surveys of figure 14 (up to am s 250) show
the elimination of the leading-edge flow separation which occurred in
the static flight condition (fig. 6).

A comparison of the lift-coefficient curves of figures 12 and 13
shows about a sevenfold increase in the values of the coefficient at a
given angle of attack when power is applied which results primarily from
the lift produced by the component of the propeller thrust in the lift
direction. The rather large values of the lift coefficient are not of
fundamental significance since the dynamic pressures for the low-velocity
conditions are very low and ultimately the lift coefficients reach infi-
nite values as the static condition is approached. It should be clearly
defined that high values of lift coefficient are not unique for this
particular configuration but would occur for any aircraft which approaches
the hovering-flight condition.

It is emphasized that the large variation in the static longitudinal
stability over the angle-of-attack range for the two forward-speed con-
ditions shown in figure 13 signifies a serious problem of control in
flight at the very low airspeeds. The untrimmed pitching moments (shown
for neutral elevator) are large and, as shown later, cannot be trimmed
with full control deflection. Some more effective tail configuration
suitably located with respect to the existing flow field would have to
be provided for these low airspeeds.

Miscellaneous stability and control measurements.- Although it was
shown by the flow studies of the slipstream in the static-thrust investi-
gation that the tail on this particular arrangement would not be in the
slipstream and, therefore, could not trim or control the airplane in very
low speed flight, it was still of some interest to obtain a little addi-
tional information on the effect of asymmetric power, power failure,
higher forward velocity, and yaw on the elevator and rudder control
effectiveness at a representative high angle of attack at a velocity of
about 25 mph. The angle of attack was arbitrarily chosen to be in the
vicinity where the component of the thrust along the X-axis equaled the
drag component (CX t o).

In general, these results show a very low control effectiveness for
the windmilling propeller condition which is increased somewhat by pro-
peller operation. The increased dynamic pressure at the tail is also
accompanied, however, by an increased downwash and the configuration
could not be trimmed in pitch for this power condition at the high angle
of attack. Also, for the asymmetric thrust condition that prevailed in
the forward-flight tests (about 25 pounds at a = 310), full rudder
deflection was required for directional trim. For single-engine oper-
ation, which would markedly increase the thrust asymmetry, the aerodynamic







NACA RM L53A09


controls would, therefore, be completely inadequate and flight would not
be possible with single-engine operation.

The objective of this investigation did not include a study of the
over-all performance at low or high speeds, inasmuch as the test configu-
ration was not a prototype and embodied several compromises which pro-
duced high drag. Since single-engine flight provokes such serious prob-
lems that power-off landings would be required in event of failure of
one engine, it is of interest to note that the maximum power-off lift-
drag ratio for the configuration tested is very low (about 1.7), although
this test configuration could undoubtedly be redesigned to give lower
fuselage and engine installation drag. Nevertheless, the power-off
landing would be a difficult maneuver with the small wing area of the
experimental airplane tested if the landing speed were not sufficiently
high to permit a flare, since for a given wing span a considerable amount
of the channel surface contributes more drag than lift, and also since
an engine nacelle supported by struts in a channel wing would be expected
to have greater over-all drag than a conventional nacelle-wing configu-
ration.


SUMMARY OF RESULTS


The results of the investigation in the Langley full-scale tunnel
to determine the characteristics of an experimental Custer Channel Wing
airplane at static conditions and over the low airspeed range are summa-
rized as follows:

1. The resultant force and the pitching moment at static conditions
(zero airspeed) in the absence of ground effect for the propeller oper-
ating at 2,450 rpm (170 horsepower total for both motors) were 880 pounds
and about -350 foot-pounds, nose down, respectively. The resultant force
was about 88 percent of the static thrust which the propellers were cal-
culated to develop in the absence of the channels and was inclined 230
upward from the propeller thrust axis. Thus, provided a tail configu-
ration could be obtained to trim the airplane, thereby reducing the
magnitude and inclination of the resultant force, the airplane, to hover
in the absence of ground effect, would have to be inclined at an angle
greater than 670 and the weight would have to be less in magnitude than
the resultant force.

2. There was no appreciable ground effect on the propeller-channel
characteristics for the static condition.

3. The controls were completely inadequate under static lift con-
ditions and the airplane could not be flown in hovering flight.







NACA RM L53A09


4. With the airplane resting on the ground there was an upward flow
at the tail which increased the total lift about 200 pounds for the static
condition. This lift force, however, is not significant since it would
be completely eliminated for longitudinal trim as soon as the airplane
tended to become airborne.

5. Small longitudinal displacement of the propeller plane did not
appreciably affect the airplane static lift characteristics. Leading-
edge stall occurred in the static flight condition and the single attempt
to eliminate the stall by an extensible leading-edge flap did not meas-
urably improve the airplane static lift characteristics. Small wind
velocities did stabilize the flow over the leading edge and increased
somewhat the lifting capabilities of the configuration.

6. Flow studies in the region of the propeller plane indicated high
inflow angles into the upper unshrouded portion of the propeller with
reversed flow occurring at the top of the propeller disk. The lower seg-
ment of the propeller disk was the more heavily loaded.

7. Low forward velocities greatly improved the flow into the channel
wing and increasing the tunnel airspeed from about 25 mph to 40 mph
resulted in more linear aerodynamic characteristics throughout the angle-
of-attack range.

8. The airplane control effectiveness was low but essentially linear
with control deflection in forward flight, so in the event of power fail-
ure the reduced control effectiveness caused by the loss of the propeller
slipstream is an important design factor to consider. The asymmetric
power condition also reduced the control effectiveness, and of consider-
able significance is the fact that the configuration would be uncontrolla-
ble with more than a small amount of power asymmetry resulting in a thrust
dissymmetry of the order of 25 pounds at an airspeed of 25 mph.


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
















10
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M__ ___I r IJI I
4H c ^' __


H

la
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d,

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aS
,
*r


0 0

CM
d,-



Ca)




to 0


-- H


1 4 4


\'0
0
-4


o I I
H CDH


co
UO


\0




cn
LO


- 4 1 & .4 1 4. .


0 10 j CC CM -
*>.. f3 H O \ '.0 CM
H 0 -\ o0 00 0 d





) r- '.0 ( O O
0 r0

- 1 1 4 10 co m m




to D o t-


r O: 0 w 0 0
0 -H 4) (, 0 C(3 0
o O 0





r ,I ) 0 L
0 0 0
441- IS.0

____________3


NACA RM L53A09


T r 1


0
r-4
-r4


r-I
P,


CO



0
-t


0\
4-.
II



4',
0(







o-


a)


0
0
H
ri
E-1


*sM)


m n o |
\(o I \0D I \D


I


Sto
4za9






NACA RM L53A09


WIND DIRECTION



W9i0 DIRECTION


AZIMUMT REFERENCE


Figure 1.- The stability system of axes. Arrows indicate positive direc-
tions of moments, forces, and angles.







NACA RM L53A09 19















o


t a
'It











--4-
So r-














S. ,\
S/ -ff H








B7^ w







20 NACA RM L53A09












I 111


















4



d
I"I 4





,,--.4. 7 As








NACA RM L53A09


h co








(6 0

I I
U
4 CO
)0

U
0. *d

o *r-

Cp







NACA RM L53A09


(a) Motor installation for original tests.


(b) Modified nacelle mock-up.

Figure 4.- Photographs of the original and modified electric-motor-nacelle
installations on the Custer Channel Wing airplane.







NACA RM L53A09 23


206














M





O Prope//er tip chord ( loca.fed -/14asi/e
of frai/ing edge of channel/

1006 -- Propea/er 1fo cho,.d /ocafed
af fra-i/ng edge of channel/


800--



600 e-



4 Q




I----- iii
,zO __ ___xnL f

0_dJ


400


800


1200 /600
RPP


2000 2400


(a) Location


of propeller in channel. p = 0.002326 slug/cu ft;
q = 0 lb/sq ft.


Figure 5.- Effect of propeller rotational speed on the static or near-
static characteristics of the Custer Channel Wing airplane in the
absence of ground effect. 8e = 00; r = 0; a = 00.


2800







NACA RM L53A09


0 Leading-edge fl/aps (excens1h/e)

/I -0 Trai/ing-edge f/aps (extendedf) anizd
iea.din/y edge f/ap.





iLong/fd/ .a
6002



40







O7
0 400 800 1200 1600 2000 2400 2800
RPM

(b) Leading-edge and trailing-edge flaps installed.
p = 0.002324 slug/cu ft; q = 0 lb/sq ft.


Figure 5.- Continued.


-4nr,


HP







NACA RM L53A09


400 800 /200 /600 2000 2400 2800

(c) Port motor operating p 0.002328 slug/u t.
(c) Port motor operating. p = 0.002328 slug/cu ft.


Figure 5.- Concluded.






NACA RM L53A09


e /f/me


cz


U 5-ta/11


-Popeler p/ane


Figure 6.- Stalling characteristics of the Custer Channel Wing airplane
in the absence of ground effect. Propellers operating at 2,450 rpm.
q = 0 lb/sq ft.






NACA RM L53A09


zoo.


/0

/06c


6C



66


S




-4


20


I-- I L I L-


60



0__ 240


^ ^ - -

^b._ ^r



--- -- -:-- -- ---l' ^ -- --


I AI I I

0 /0 20 3O 40
oC, d


Figure 7.- Effect of inclination of thrust axis on the static character-
istics of the Custer Channel Wing airplane in the absence of ground
effect. Propellers operating at 2,450 rpm; configuration untrimmed
in pitch. q = 0 lb/sq ft; be= -100.


IOC


Lon, it nql
o:00ofrce f ^

I E!


AnA


<^f







NACA RM L53AO9


/00



O

O


































ELif
20
__ to --*""- ,*--,-* --- \ --- /- ----\ -mm --^6....... ........ \^ rf- _






















200


OL
-40
up


-- -- -- -- -- -- AC ^ _
I I I

-30 -20 -/O 0 /0 20 30
de, deg down


(a) a, = 00; r = 0; p = 0.002308 slug/cu ft.

Figure 8.- Effect of control-surface deflection on the static character-
istics of the Custer Channel Wing airplane in the absence of ground
effect. Propellers operating at 2,450 rpm. q = 0 Ib/sq ft.


___


I I


J,







NACA RM L53A09


200






O





















I -X
100 -





S-2a -----





-2_00 __- --(--- --- ----o












400o




200- 1i1


oL


-JO
r/ght


I I


I I
I I I


0 /0 20 30
sr, deg


(b) a = 00; be = 00; p = 0.002308 slug/cu ft.


Figure 8.- Continued.


-I I I I I I I I I L


"''''''""'








NACA RM L53A09


H





O
-200






0
-0 -------- -- -- --i- --- -







O


L













-2
0f









600-







4 --


200
-Z 6 )-?- Z ''O _






0 __ __ __ __ __ __ __ _o___ e


-P O O __ ___ y ^ ___ ____ ____ ___ ____ ___ ____ ____ ___ ____ ___


- I-I


I i I


-30 -2 -0 -/ 0 /0 20 30
6e), de9 down


= 460; 6 = 00; p = 0.002312 slug/cu ft.

Figure 8.- Continued.


0
-40
up


(c) a


vcrM/l







NACA RM L53A09 31


0
( ---- ----4 )--(|---( )--<)---< ---( W-- ----( ---- ----()p/













0-2











0
0 -- ---- -- --.- ____



-400 -- m


-o--- .--F<---- ---^.if



600




() _---- <" -- --( ----)--- -- 0Lon gi,


fo


-20 -/0


0
6Sr deg


/O 20 30 40
le f-


(d) a = 460; 6e = 0; p = 0.002312 slug/cu ft.

Figure 8.- Concluded.


'd/l7a/


-,T
-30
ri hf-








NACA RM L53A09


80-- ----- --0
60






40














STail off






-/ ---


600-


6000






400-
n ~ ~~~ To ./ toff A


up
up


u
6e ,deg


don
down


40eg


dTU
down


(a) 2,450 rpm.


(b) 2,625 rpm.


Figure 9.- Effect of the horizontal tail on the static characteristics
of the Custer Channel Wing airplane in the presence of a ground
boundary. Propellers operating; basic electric-motor nacelle.
q = 0 lb/sq ft.


s
j
PE
~io
uY-


f





NACA RM L53A09


r


(a) Above and in front of the propeller disk.


(b) Behind the propeller disk near the tip radius.
Figure 10.- Flow studies of the Custer Channel Wing airplane in the static
ground tests. Propellers operating at 2,450 rpm. a % 190; 8e = 00.


NACA


01,








NACA RM L53A09


.mu~
*hT? 'i~-Nl~;


(c) Behind the propeller on the center line of nacelle.


(d) Ahead of horizontal tail.


Figure 10.- Concluded.


.~NACA/
"'







NACA RM L53A09 35











o< oo








S0 r-I

J0 0
I .I o



400l




4 0 ; q.-
-- -^ -- -- ^ -- -4,--- r.. o
dom
\ 1 0i o

0 N



\ H0 -1o





0 ^ ^P)
--m r-4
0 00 0l
iICJ c? 'n o


'/a-v -/ Io./p/A/uo7


9/ '4t/7
















































































0 O
c3r ~j 4
I I


NACA RM L53A09














I
_------ 0 --
r-1











1
o






to
o O

0N 0
+ 1.



a & 0

0 0 It
p) 4- OH
r-1 r4-3
M o




N ID
P -I








---4
01
0
-- I na a

? ,L (





1g ~~ da






















-d



















11
\\













-- t--
F ~t







(O .














E Er|E1 =E


0o~
A9 U 2


N
4


o !

rS'


NACA RM L53A09


t-4




0
0






H
t-(
0

II





*r


Tl
b


c o '
1. 1.







38 NACA RM L53A09









C





l II
'00
-- -- -- -- -- ------ -- -- -- -- --

--,-------- -- n g /o







o


S,-I I I




+ o
0 4O 0






-U
t
c *a ,4






-- --- ---- \- __ J __ ^






^ -- -v ~b
^ -^ -4 ^ -. I?
(a __ i __ __ i ^ __ S f
---- -- -- -- \ ^ o


-ira 'arodas o X







NACA RM L53A09 39







'- ,






0 El


/ ) O
--. w













-- 4-
-- -- I-- -- -- I -- ---




-- -0--6 -- -- p n -- -- -- -d fi -








NACA RM L53A09

-o




- J


- 0

J






J o
4.



0



O-

0 o




3 i A


mo' )




) .+


\I dI Px


N^ ^








NACA RM L53A09 41










0


z:
\o













0 <
-- -- ---- ^ ---- .--0






N)








00
K)
i ____ 2 b -_ _<)-- -





\V10








b ~
------------------------------------------4-_ -_ N
6 e. V ^

o o p

__ __ __ __ __ 4 __ __ __ 4r-


.dal/oda9s0oH







NACA RM L53A09


OC =30./O


Rouyh

Intermfftten fall


Sfta//


Figure 14.- Stalling characteristics of the Custer Channel Wing airplane.
Propellers operating at 2,450 rpm. 6e = 00; 6r = 00; q = 1.74 lb/sq ft:
p = 0.002323 slug/cu ft.


use /age
/'n e-.


Prv'pe//er ,v/a/7e


OC 2.5.2







NACA RM L53A09


-40 -30 -20 -/O
up 6e, e9


down


(a) 8r = 00; q = 1.71 lb/sq ft; p = 0.002280 slug/cu ft.

Figure 15.- Effect of control deflection on the aerodynamic characteristics
of the Custer Channel Wing airplane. Propellers operating at 2,450 rpm.
a = 31.1







NACA RM L53A09


-30 -20 -/0 0 /0 20 30 40
rght Sr, de9 left


(b) 5e = 00; q = 1.71 Ib/sq ft; p = 0.002270 slug/cu ft.


Figure 15.- Concluded.








NACA RM L53A09


U -.


/






o __ ___ ____ ^e==) _.c _













4
2
I- T In










C4- -








S __ _-NACA
o__________ -LaJ _L


(a) 5r = 0;


q = 1.70 Ib/sq ft; p = 0.002324 slug/cu ft.


Figure 16.- Effect of control deflection on the aerodynamic character-
istics of the Custer Channel Wing airplane. Propellers windmilling.
.a = 360.


-40
up


-i o
-/0 0
o' eg


20 30
o'oA1n


A







NACA RM L53A09


./-----__--______-





O ---17 JCz

0-tF



0
9
03--------- 0--.--. .0-"(--.-- )----OC,,









LL


2 _(-- -- <)-(-. ..)-(-( ---(--- ---() -_




2l

--IACA
0 __ __ __ __ __ __ __ __ __ __ I I


-50
rnghf


-20 -10


(b) 8e = 00; q = 1.74 Ib/sq ft; p = 0.002323 slug/cu ft.

Figure 16.- Concluded.


0 /0 20 30 40
6r, deg left


ty


US~







NACA RM L53A09


200




























-11 --- -20-- /---O /0 20 ) O
U0 -de down
4-










-Z


















-40 -.0 -20 -/0 0 /0 20 30
up 6c, de9 down


(a) 8r =' 0; q = 1.78 lb/sq ft; p = 0.002330 slug/cu ft.

Figure 17.- Effect of control deflection on the aerodynamic character-
istics of the Custer Channel Wing airplane with asymmetric power.
Port propeller operating at 2,450 rpm. a = 33.80.







NACA RM L53A09


-20 -/0 0 /0 20 30 40
,, deg left


(b) 5e = 0; q = 1.78 lb/sq ft; p = 0.002330 slug/cu ft.

Figure 17.- Concluded.


-30
nght







NACA RM L53A09


j


0-I-- -I-I--I-T-IjI--
LI _I --
T I


0 /0 20 30
-9 down


(a) p = -5.060 a = -2.00; 8r = 0; q = 2.20 lb/sq ft;
p = 0.002330 slug/cu ft.

Figure 18.- Effect of control deflection on the aerodynamic character-
istics of the Custer Channel Wing airplane in sideslip. Propellers
operating at 2,450 rpm.


C


















6
9





C",


6-
e- -- _l.^> -- -- -- --









9(
__ -- -- -^ l~ --
I I I


e-,dd








NACA RM L53A09


U


















u"


-/0 0 /0 20
r ,deg


Sa= -2.0; 8e = 00; q=
p = 0.002346 slug/cu ft.

Figure 18.- Continued.


30 40
/eff


2.19 lb/sq ft;


.6 0
.-2>-- ---j





.2













0


- --- ---- -- ----- -- -











8






-I I-
61r h A-L __ _,
o -o -o -o -(>- <)- -CA y
0 __ __ __ __ __ __ __ _

6 ___ ____


(b)



(b) :


-30 -20

= 5h06

= -5.06,


9
9
9
pr







NACA RM L53A09


5e, deg


down


(c) p = -5.06; a = 31.10; or = 0; q = 1.71 Ib/sq ft;
p = 0.002317 slug/cu ft.
Figure 18.- Continued.








ACA RM L53A09


cCy


0 0










CCn









0
-- -- -- -- -- -- -"t- -.-o. -- -- -









24



P2 '
Nnc
2 4 _ _


-20 -/0


/0 20 30 40
leff


(d) p = -5.060; a = 31.1; ,e = ; q = 1.71 lb/sq ft;
p = 0.002315 slug/eu ft.


Figure 18.- Continued.


2.


-JO
right


0
Sr, deg








NACA RM L53A09


HP

/00 F




/.o



V- --- --14C-


Q __ __ ___ ___ __ __ ^ __ __ -- ( Cy _


-2


4


0


6






8


O 6


-40 -30 -20 -/0 0
UP Se deg

(e) p = -9.83o; a = -2.00; r = 00;
p = 0.002320 slug/cu


cm













---,NACA


/0 20 30
down


q = 2.20 Ib/sq ft;
ft.


Figure 18.- Continued.








NACA RM L53A09


6.
.6 _________________ _200
















4-
-- =---O- ----- .---^i"--- -- --------C




-. 0----- ---0----- 0-- --------C





_ ~^ ^ cL

- __ _


-20 -/O 0 /0 20 30 40
6,, Oe9g /eff


2
o


(f) p = -9.830; m = -2.00; be = 0; q = 2.17 lb/sq ft;
p = 0.002324 slug/cu ft.

Figure 18.- Continued.


-30
PrhOf







NACA RM L53A09


-30 -20 -/0 0 /0 20 30
6,-, deq do'n


(g) p = -9.830; a = 31.10; br = o0; q = 1.73 lb/sq ft;
p = 0.002337 slug/cu ft.

Figure 18.- Continued.







NACA RM L53A09


.8 1nn

8 _I -- -<>(- ( -- o0








0- 0









-2








2












CL
20
s---


-30
nght


-20 -/0 0
&-, deg


/0 20


30 40
leff


(h) ~ = -9.830; a = 31.10; 8e = 00; 1.74 lb/sq ft;

p = 0.002337 slug/cu ft.


Figure 18.- Concluded.








NACA RM L53A09 57



8 /2







O 4





6 0













/0 2.4-

Ailerons

on
8-- 22 -Z0-o




----i 0 -1
0 3-





Z. /6 I
-/0 -8 -6 -4 -2 0 -/0 -6 -6 -4- -2 0

o,deg @, do9

oC=-2.0; iZ2.20 /b/sq fc; C -- 31// al 75316/.s ff;
VWs2.7mph a, 2L ^*4ph

Figure 19.- Effect of ailerons on the longitudinal aerodynamic charac-
teristics of the Custer Channel Wing airplane. Propellers operating
at 2,450 rpm. 5e = 0; r = 00; aileron area, 17.5 sq ft.


NACA Langley Field, V.A.







*t





I:


































































':
':
V.














I
ir
i-


i:



ii







|:





UNIVERSITY OF FLORIDA


312620!


UNIVERSITY
DOCUMENT
120 MARST
P.O. BOX 11
GAINESVILL


81066168 J .:......".:. : i :. i:
... ** :.. Y'4: ":
-------- -: --,-
:. .


OF FLORIDA '
rS DEPARTMENT
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