Wind-tunnel investigation of the aerodynamic and structural deflection characteristics of the Goodyear inflatoplane

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Material Information

Title:
Wind-tunnel investigation of the aerodynamic and structural deflection characteristics of the Goodyear inflatoplane
Series Title:
NACA RM
Physical Description:
56 p. : ill. ; 28 cm.
Language:
English
Creator:
Cocke, Bennie W.
Langley Aeronautical Laboratory
United States -- National Advisory Committee for Aeronautics
Publisher:
NACA
Place of Publication:
Washington, D.C
Publication Date:

Subjects

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

Notes

Summary:
Summary: An investigation has been conducted in the Langley full-scale tunnel to determine the aerodynamic and structural deflection characteristics of the Goodyear Inflatoplane over a range of test velocities from minimum stall speed up to speeds giving load factors for wing buckling. Tests were conducted over a range of speeds from approximately 41 to 70 mph with wing-guy-cable loads, wing-distortion photographs, and aerodynamic-force data recorded at each speed for a full range of angle of attack.
Statement of Responsibility:
by Bennie W. Cocke, Jr.
General Note:
"Report date September 10, 1958."

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003833712
oclc - 145593172
sobekcm - AA00006133_00001
System ID:
AA00006133:00001

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NACA RM L58E09

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS



RESEARCH MEMORANDUM



WIND-TUNNEL INVESTIGATION OF THE AERODYNAMIC

AND STRUCTURAL DEFLECTION CHARACTERISTICS

OF THE GOODYEAR INFLATOPLANE1

By Bennie W. Cocke, Jr.


SUMMARY


An investigation has been conducted in the Langley full-scale
tunnel to determine the aerodynamic and structural deflection character-
istics of the Goodyear Inflatoplane over a range of test velocities from
minimum stall speed up to speeds giving load factors for wing buckling.
Tests were conducted over a range of speeds from approximately 41 to
70 mph with wing-guy-cable loads, wing-distortion photographs, and
aerodynamic-force data recorded at each speed for a full range of angle
of attack.

The airplane was longitudinally stable and had adequate pitch and
roll control and normal stall characteristics at the lower speeds giving
maximum load factors between 1 and 1.5. However, as speed was increased,
aeroelastic effects associated with wing twist produced an increase in
lift-curve slope and loss of stability near the stall. For speeds up
to 65 mph, which produced a load factor of approximately 2, the maximum
wing load was limited by stall with moderate wing deflections. However,
at a speed just over 70 mph and at an attitude producing a load factor
just over 2, a column-type buckling occurred on the inboard wing panel
with the inboard wing section folding up and contacting the engine
mounted above the wing. Additional tests were made with modifications
to the wing-guy-cable system which reduced the aeroelastic effects on
the aerodynamic characteristics and allowed load factors up to approxi-
mately 2.5 before a tendency for wing buckling occurred.


iThe information presented herein was previously given limited
distribution.






2 NACA RM L58E09


INTRODUCTION


The need for a means of rescue or escape for fliers downed in enemy
territory has prompted the Military Services to consider a number of
possible rescue concepts. One scheme considered would utilize a small
lightweight airplane made from inflatable structure which when deflated
could be completely contained in a small lightweight package and para-
chuted to a downed man for self-rescue at the most opportune moment.
This idea has been developed under contract by the Office of Naval
Research to the point of successful flight demonstration of a single-
place prototype model which can be deflated and packaged in a size
and weight which can be handled by one man.

As the pneumatic structure used in this airplane is novel and does
not lend itself to existing structural theory, the prototype airplane was
tested in the Langley full-scale tunnel to obtain data on its character-
istics under various aerodynamic loadings up to wing failure to provide
data for correlation with existing theory developed from static-load
tests.

This report presents the results of the wind-tunnel tests, during
which the aerodynamic characteristics and control effectiveness were
obtained in addition to wing-guy-cable loadings and wing-deflection
records for a range of wind velocities from approximately 36 to 70 mph.
The characteristics of the configuration were also determined for a
range of reduced inflation pressures simulating leakage due to battle
damage or compressor malfunction.


SYMBOLS


CL lift coefficient,
L qS

CD drag coefficient, Drag
qS


Cm pitching-moment coefficient, Pitching moment
qSE

C, yawing-moment coefficient, Yawing moment
qSb

CZ rolling-moment coefficient, Rolling moment
qSb






NACA RM L58E09 5


q free-stream dynamic pressure, Ib/sq ft

S wing area, sq ft

c mean aerodynamic chord, ft

b wing span, ft

a angle of attack (angle between relative wind and fuselage
water Line 50), deg

V stream velocity, mph

Pf fuselage inflation pressure, Ib/sq in.

P wing inflation pressure, lb/sq in.


Be elevator deflection angle, positive when trailing edge
deflected down, deg

6a aileron deflection angle, positive when trailing edge
deflected down, deg

&6 control stick deflection angle, deg

Subscripts:

L left

R right

av average


AIRPLANE AND APPARATUS


The Goodyear Inflatoplane used in this program is composed of
pneumatic structure throughout with exceptions of the engine, engine
mount, landing gear, and miscellaneous short control members. All
inflatable components are interconnected in a manner allowing a small
compressor on the 40-hp air-cooled engine to maintain a constant regu-
lated pressure in the system even with moderate leakage. The wing and
tail surfaces are woven in a manner such that the upper and lower air-
foil surfaces are connected internally by nylon drop threads varying
in length to produce the approximate shape control desired in any






NACA RM L58E09


surface when inflated. A circular fuselage is utilized with a fuel bag
internally mounted and the cockpit section is constructed using sections
of air mat material 2 inches thick. Design gross weight of the airplane
is 550 pounds with 120 pounds of fuel and 240 pounds payload.

Each wing panel is restrained by two guy cables on the upper and
lower surfaces with the two upper cables anchored to the engine pylon
and the two lower cables anchored to the landing gear. Both upper and
lower cables attach to patches bonded to the wing surface approximately
0.57b out from the center line. A general layout with pertinent
geometric data is shown in figure 1 and a photograph of the airplane is
presented in figure 2. The propeller was removed for this load program
primarily for safety considerations.

The airplane was mounted for tests on the conventional six-component
mechanical balance as shown in figure 3. A special yoke (fig. 5(b)) was
utilized to mount the airplane so that strut restraining loads were
transmitted to the fuselage through strap attachments located beneath
the wing quarter-chord point thus leaving the wings free to deflect while
being restrained only by the normal wing-fuselage and guy-cable attach-
ments as in flight. The tail strut was attached to a saddle strapped
to the rear of the fuselage (fig. 5(c)) and was connected by cables to
the front support yoke thus preventing longitudinal tail strut loads
from being transmitted into the fuselage.

An actuator system was installed in the cockpit to allow remote
operation of the elevators and ailerons which were equipped with control-
position transmitters located on the respective surfaces to record the
position settings of each control. Control-position transmitters were
placed on both the right and left side of the elevator surface to give
indication of the amount of twist occurring under load since the eleva-
tor was actuated by a single horn. Strain-gage units were installed
in all wing guy cables for cable load evaluation, and cameras were set
up to record the deflection of the left wing panel under the various
loading conditions. The left wing was chosen for photographic study
as the contours and wing geometry of the panel were more uniform than
those of the right panel.


TESTS


The objective of this program was to determine the aerodynamic
and wing deflection characteristics of the Inflatoplane under various
loading conditions. Tests were conducted at various airspeeds ranging
from approximately 36 to 71 mph with the airplane angle of attack
increased by small increments at each airspeed until either wing stall






NACA RM L58E09


occurred or the wing buckled. This sequence of tests made with the
airplane at normal inflation pressure was repeated for two airplane con-
figurations, namely, the basic original airplane and the airplane with
an additional guy cable added to the lower surface of each wing panel
to provide additional rigidity. For the configuration with additional
guy cables, tests were also made with wing and fuselage pressures
reduced for wind speeds near minimum flight speed to determine a safe
minimum inflation pressure for maintaining flight. In conjunction with
the tests made at normal pressure, aileron and elevator control effective-
ness were measured on the original airplane configuration for speeds
chosen to represent the minimum and cruise flight regions.

Aerodynamic force and moment data were recorded for each of the
runs, and for a representative range of loading conditions, wing-guy-
cable loads were recorded along with photographic records of the deflec-
tion of the left wing panel. For conditions where wing buckling was
reached, motion pictures were used to record the wing motions after
collapse.

All data presented in this paper have been corrected for wind-
tunnel buoyancy, jet boundary, and stream misalinment. Support-strut
tares were not measured since major emphasis was placed on obtaining
loads information. All drag results, therefore, include the tare drag
of the support system. Pitching-moment data shown are computed for a
center of gravity located longitudinally at fuselage station 72.7 and
vertically at water line 45.5.


RESULTS AND DISCUSSION


Longitudinal Aerodynamic Characteristics and Wing Buckling Tests

Original airplane configuration.- The results of force measurements
(fig. 4) made at various airspeeds for the original Inflatoplane con-
figuration at normal inflation pressure (7 lb/sq in.) showed a marked
effect of dynamic pressure on the variations of lift coefficient and
pitching moment with airplane attitude. These variations are attributed
to aeroelastic effects. At the lower speed (qa = 4 Ib/sq ft) which
would closely approximate a minimum steady flight speed (approximately
1 g at CL,max), the lift curve was linear and the airplane was stable
through stall. With increasing speed, an increase in lift-curve slope
is apparent and the airplane becomes unstable in the high lift coeffi-
cient range representing accelerated flight. For the speeds corre-
sponding to average dynamic pressures of 4, 7, and 10, wing stall was
reached at each speed and the value of CL,max obtained was reduced
Climax






NACA RM L58E09


with increased speed. The CL,max value of 1.0 at qav = 10 lb/sq ft
was reached at an a of approximately -20 and gave a load factor slightly
less than 2 with moderate wing deflections but no signs of wing failure.
With the tunnel speed increased to approximately 71 mph (av = 12), a
run was made with angle of attack increased by 10 increments, and when
an angle of attack of approximately -5 was reached wing buckling
occurred suddenly after approximately 50 seconds time had elapsed at
this condition. The wing recovered quickly when load was reduced after
buckling without any apparent damage, however observation of the wing
behavior indicated that if a propeller had been installed and operating
the wing would have been destroyed. As it was important to obtain the
loadings for this condition, the run was repeated with angle of attack
increased by increments of 0.250 up to the -5o attitude of which three
load readings were taken prior to collapse. This information showed
a steady increase in load with time with the fuselage attitude held
constant thus indicating that stretch in the nylon fabric at high
loadings was allowing the wing to increase effective attitude with
respect to the fuselage. The last recorded load prior to wing buckle
for this condition was 1,154 pounds for a load factor slightly in excess
of 2.

No apparent damage to the airplane resulted from these first two
buckling experiences; therefore, an additional buckling test was pro-
gramed with more complete motion-picture and still-photographic cover-
age for study of the rapid motions of the wing during the 2 or 5 cycles
of buckling and recovery which the wing went through despite prompt
shutdown of the wind tunnel. For this additional photographic run the
airplane attitude was slowly increased from -100 to -50 and continuous
movie coverage was taken as the wing loaded and buckled. In this
sequence the rear wing-guy-cable patch on the lower surface of the left
wing tore on the second buckling cycle and the wing contacted the engine
and was punctured by the spark plugs and propeller shaft. Photographs
showing the wing at the onset of buckling and just after puncture are
shown in figure 5. Motion pictures and still photographs of the wing
buckling showed a column-type failure inboard approximately midway
between the fuselage and wing-guy-cable-attachment points with the wing
folding inboard and moving up and in so as to bring the inboard wing
sections well into the propeller disk area. Close study of photographs
of the buckling runs made prior to the failure of the rear guy-cable
patch and wing puncture showed further that the rear guy cable had
fouled on the model support system during these runs and actually snub-
bed the wing in its upward travel, thus probably preventing wing punc-
ture due to contact with engine during the first buckling tests. Initial
buckling in all cases occurred on the left wing panel which developed
a slightly higher loading than the right panel. This load asymmetry,
also indicated by the higher magnitude of the left-wing-guy-cable loads






NACA RM L58E09


and by positive airplane rolling moments, was attributed at least in
part to negative camber evident in the airfoil sections in the region
of the right wing tip.

Airplane with additional wing guy cables.- On the basis of the
basic tests it was desirable to modify the wing attachment system to
improve the aerodynamic characteristics of the airplane at higher speeds
and to improve its load-carrying ability and arrest its motions after
wing failure. From studies of the data and photographs it was felt that
the aeroelastic effects shown in the data were largely associated with
the deflection of the inboard wing sections which resulted in increasing
wing incidence with load. This increase was believed to produce an
unfavorable downwash at the tall resulting in the loss in stability at
higher loadings. As the wing failure was similar to that of an eccen-
trically loaded column it also was obvious that some additional restraint
inboard should give higher load capability, while, at the same time,
offering some possibility of improving stability.

During the time used for patching the wing punctures, provisions
were made for addition of two new guy cables on the lower surface of
each wing panel. Attachment points for these cables were located on
the wing at chordwise locations approximately 0.25E and 0.60E at a span-
wise point approximately 4 feet from the fuselage beneath the point
where buckling was first observed. The front cable was rigged taut to
take load and the rear cable was left slack to serve only to reduce wing
motion should buckling occur. This approach was taken as it was felt
that adding the cable near the center of pressure should provide the
necessary restraint and offer the greatest chance of reducing unfavorable
wing twist on the inboard wing sections.

The results of the tests made at normal inflation pressure
(7 lb/sq in.) with the additional guy cables installed (fig. 6) showed
an appreciable improvement in the aerodynamic characteristics along with
higher load capability. For this configuration the degree of instability
resulting at the higher loadings is greatly reduced and the increase in
lift-curve slope with speed was noticeably less. Also for the higher
speed condition, wing buckling finally occurred only after a condition
of intermittent stall of the left wing developed which produced a series
of wing oscillations which were followed by buckling along a chordwise
line approximately 2 feet out from the fuselage center line. For this
configuration a lift load of approximately 1,500 pounds was recorded
for an airplane attitude approximately 10 below stall (a = -3.1). Stall
and buckling occurred as the airplane attitude was further increased
to -2.1. Loads for this condition can only be estimated but it is
reasonable to believe that a value of CL,max of at least 1.0 was
reached with the maximum load approaching 1,400 pounds for a load factor
of approximately 2.5.







NACA RM L58E09


Although buckling occurred at this condition, it is conceivable
that the longer time lapse and the intermittent stall preceding buckling
would be an effective warning of the approaching buckling boundary. It
is also felt that by further modification to the wing-guy-cable and wing-
root attachments, additional improvements in the stability and load limit
could probably be obtained; however, such development was beyond the
intended scope of this program.

As an additional point of interest in connection with load char-
acteristics of the pneumatic structure, a short series of tests was made
with wing and fuselage pressures reduced from normal pressure to ascer-
tain the possibility of maintaining flight near minimum speed in case
of an emergency caused by loss of air pressure. The results of the
reduced pressure tests (fig. 7) did not show any drastic changes in
aerodynamic characteristics which should rule out flight down to the
lowest test pressure of 5 lb/sq in. At this pressure the wing did not
buckle for the test speed with the maximum load factor reaching approxi-
mately 1, and brief aileron and elevator control checks indicated that
control could be maintained.


Wing-Guy-Cable Loads and Deflection Studies

Loads measured in the wing guy cables for the two cable configura-
tions tested are summarized in figures 8 and 9 where the individual
cable loads are plotted as a function of total configuration lift. Wing-
deflection photographs for some of the more pertinent conditions are
presented in figures 10 to 14. These photographs have airplane angle
of attack and total configuration lift noted on each to permit correla-
tion with the proper cable loads and aerodynamic data plots. Horizontal
stripes on the vertical deflection target bar shown at the wing tips
in the photographs are spaced 2 inches apart. The long stripe on the
horizontal bar provided general horizontal reference.

The cable-load data for the original configuration (fig. 8) indi-
cate that the wing bending due to lift is primarily restrained by the
front guy cables and the bending due to chordwise forces by the rear
guy cables. At zero lift the rear-guy-cable load is therefore larger
than the front-cable load at all speeds; but as lift is increased, the
front-cable load increases rapidly and for the high loading condition
(fig. 8(d)) the front-cable loads reach values over three times as large
as the rear-cable loads. With this cable configuration the front-cable
load was approximately twice the rear-cable load for the 1 g condition
(550-pound lift) condition at all test speeds.

Cable-load data for the modified cable system (fig. 9) show that
the additional cable attached forward and inboard on the wing







NACA RM L58E09


appreciably reduced the load in the original front guy cable at the
higher load conditions but had only a small effect on the rear-cable
load. The maximum load reached on the inboard cable was approximately
200 pounds for the maximum loading condition (fig. 9(c)) which imposed
loads of over 600 pounds on the outboard front wing cable. The inboard
cable could probably have been made to carry more load by preloading;
however, the additional cable as installed raised the allowable wing
load to a value slightly above stall onset at maximum design speed
(71 mph). If a higher allowable load is required, it is felt that some
further improvement could best be obtained by adding another light guy
cable at the new initial failure point which for the modified cable
system was approximately 2 feet out from fuselage center line.

The wing-deflection photographs for the two cable systems at all
speeds tested (figs. 10 to 14) show only small differences in deflec-
tion for the two cable systems at load conditions approximating steady
level flight (L = 550 pounds). At the higher loading conditions obtained
at the higher speeds, however, the wing deflections are noticeably dif-
ferent. For the original cable installation the deflection inboard is
seen to build up with load (fig. 12) until failure is reached at a load
just over 1,100 pounds. For the modified cable system at the same speed
(fig. 14) the deflection inboard is less and at the higher loadings
reached with this cable system the tip sections show a more pronounced
deflection.


Control Characteristics

The static longitudinal characteristics of the airplane with the
elevators deflected are presented in figure 15 for test speeds averaging
41 and 64 miles per hour. The data for these speeds chosen to represent
minimum-speed and cruising-speed flight conditions, respectfully, show
a marked change in static stability with speed but little change in the
effectiveness of the elevator to produce trim. Comparison of the eleva-
tor angles obtained for a given deflection of the control stick indicated
somewhat lower response of control motion to a given stick motion at the
higher speeds. Adequate control would appear to be available, however,
and the loss of response is apparently the result of stretch within the
semirigid control system at higher loads. The amount of twist occurring
in the elevator control surface was small as may be seen in figure 15.

Longitudinal and lateral aerodynamic data obtained with the ailerons
deflected for the same test velocities previously discussed are shown
in figures 16 and 17. A reduction in rolling-moment coefficient for a
given stick deflection is evident at the higher speed condition. This
reduction apparently comes from aeroelastic effects and from a reduction
in control motion with stick deflection at the higher loading condition.






NACA MN L58E09


The reduced control motion is most apparent for the up-aileron in all
cases because the up-aileron is acutated by a simple bungee cord attached
to the upper surface so that preset tension in this cord pulls the con-
trol up when tension is relaxed in the lower actuating cable by deflec-
tion of the stick. Rolling-moment coefficients higher than those shown
for the condition at qav = 10 could have been obtained by utilizing
full control travel; however, with the model rigidly mounted through
the fuselage for these tests, a nondesign condition existed with the
wing rolling moment applied with fuselage restrained. Maximum rolling-
moment tests, therefore, were not made at the higher speed.


SUMMARY OF RESULTS


The results of tests on the Goodyear Inflatoplane may be summarized
as follows:

1. The airplane, with original guy cables, was stable through stall
for low speed conditions but for the higher speed conditions exhibited
instability in the lift-coefficient range representing load factors
greater than 1.

2. With the original wing-guy-cable configuration, wing stall occur-
red without any wing buckling for test speeds to 64 mph (load factor
just under 2) but at approximately 70 mph wing buckling occurred with
a load factor slightly higher than 2.

5. The longitudinal instability noted for the higher load factor
conditions was apparently the result of increased wing incidence inboard
due to growth and stretch in the nylon fabric.

4. Wing buckling which occurred as a column type failure inboard
approximately midway between the fuselage and wing-guy-cable attachment
points resulted in the inboard wing sections folding upward in a manner
to bring them within the propeller disk area. When wing puncture did
not occur due to contact with the engine (without propeller), wing
recovery from a buckled condition was instantaneous with load reduction
and without apparent damage.

5. The addition of one wing guy cable attached on the lower wing
surface at the point of initial buckling appreciably reduced the static
longitudinal instability at higher speeds and allowed the airplane to
reach stall at 70 mph before buckling occurred. Buckling followed the
wing oscillations produced by stall. Maximum lift loads reached
approximately 1,400 pounds (load factor approximately 2.5) before stall.






NACA RM L58E09


6. For all configurations studied, the wing behavior following
buckling was such that an operating propeller would have struck and
destroyed the wing.

7. Tests made with airplane inflation pressure reduced and air
speeds considered minimum for maintaining level flight indicate that
flight should be possible in an emergency for inflation pressures less
than one-half the normal inflation pressure.

8. Elevator and aileron control characteristics were modified some-
what by changes in speed due to flexibility in the structure and con-
trol system; however, adequate control should be maintained throughout
the design speed range.


Langley Aeronautical Laboratory,
National Advisory Committee for Aeronautics,
Langley Field, Va., April 17, 1958.




NACA RM L58E09


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NACA RM L58E09


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(b) Main support yoke.


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Figure 5.- Concluded.


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16 NACA RM L58E09








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NACA RM L58E09


Total lift, lb


(a) V = 41 mph; qv = 4.06 lb/sq ft.

Figure 8.- Variation of wing-guy-cable loads with airplane total lift for
several wind velocities. Original configuration; normal inflation
pressure (7 lb/sq in.); controls neutral.






NACA RM L58E09


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Left front
SRight Front
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Total lift, lb

(b) V = 54 mph; qav = 7.07 lb/sq ft.


Figure 8.- Continued.






NACA RM L58E09


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Total lift, lb

(c) V = 64 mph; qav = 10.15 Ib/sq ft.


Figure 8.- Continued.


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NACA IM L58E09


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(d) V = 71 mph; qa = 12.2 lb/sq ft.


Figure 8.- Concluded.


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NACA RM L58E09


So -4o0 600 o00 /0oo
Total lift, lb

(a) V = 54 mph; qv = 7.07 lb/sq ft.

Figure 9.- Variation of wing-guy-cable loads with airplane total lift for
several wind velocities. Additional wing guy cables installed; normal
inflation pressure (7 lb/sq in.); controls neutral.






NACA RM L58E09


Cable
0 Left front
U Right front
D Left rear
SRight rear
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(b) V = 64 mph; q = 10.15 lb/sq ft.


Figure 9.- Continued.


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NACA RM L58E09


Cable
Left front
Right front
Left rear
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Figure 9.- Concluded.


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NACA RM L58E09


Upper camera; a = -35; L = 562 lb.


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Lower camera; a = -3o; L = 562 lb. L-58-1655

Figure 10.- Deflection study photographs for original configuration.
V = 54 mph; qav = 7.07 lb/sq ft.






NACA RM L58E09


Upper camera; a. = -1.2o; L = 676 lb.


Lower camera; a. = -1.2o; L = 676 lb.


L-58-1634


Figure 10.- Continued.




NACA RM L58E09


Upper camera; a = 0.7 ; L = 7.88 lb.


Lower camera; a = 0.70; L = 788 lb.
Figure 10.- Continued.


L-58-16355


`





NACA RM L58E09


Upper camera; a = 1.7 ; L = 830 lb.


Lower camera; a = 1.70; L = 830 lb.


Figure 10.- Concluded.


L-58-1656


E!3





NACA RM L58E09


Upper camera; a = -5.8o; L = 557 lb.


Lower camera; a = -5.80; L = 557 lb. L-58-1637
Figure 11.- Deflection study photographs for original configuration.
V = 64 mph; qv = 10.15 lb/sq ft.


~L~L~c


'Fir .L:-.- ,? .-IJ
LJ f- '' .






NACA RM L58E09


Upper camera; a. = -4.90; L = 639 lb.


Lower camera; a = -4.90; L = 639 lb.


Figure 11.- Continued.


L-58-1658






NACA RM L58E09


Upper camera; a = -40; L = 765 lb.


Lower camera; a = -4 ; L = 765 lb.


L-58-1639


Figure 11.- Continued.






NACA RM L58E09


Upper camera; a = -3.1 ; L = 904 lb.


Lower camera; a = -5.10; L = 904 lb.


Figure 11.- Concluded.


L-58-1640






NACA RM L58E09


Upper camera; a = -6.80; L = 616 lb.


Lover camera; a = -6.80; L = 616 lb. L-58-1641

Figure 12.- Deflection study photographs for original configuration.
V = 71 mph; q =a 12.2 Ib/sq ft.





NACA RM L58E09


Upper camera; a. = -6o; L = 850 lb.


Lower camera; a = -60; L = 850 lb.
Figure 12.- Continued.


L-58-1642


~nc*~(-~


L~,Z~f~:





NACA RM L58E09


Upper camera; a = -5.750; L = 890 lb.


Upper camera; a = -5.50; L = 958 lb.


L-58-1645


Figure 12.- Continued.






NACA RM L58E09


Upper camera; a = -5.550; L = 1,011 lb.


Lower camera; a = -5.10; L = 1,100 lb.


Figure 12.- Continued.


L-58-1644






40 NACA RM L58E09

































Upper camera; a = -4.75. L-58-1645


Figure 12.- Concluded.






NACA RM L58E09


Upper camera; a = -4.90; L = 650 lb.


Lower camera; a = -4.90; L = 650 lb. L-58-1646

Figure 15.- Deflection study photographs with additional wing guy cables
installed. V = 64 mph; qav = 10.15 lb/sq ft.






NACA RM L58E09


Upper camera; a = -3.1; L = 849 lb.


Lower camera; a = -3.1; L = 849 lb.

Figure 15.- Continued.


L-58-1647





NACA RM L58E09


Upper camera; a. = -2.20; L = 971 lb.


Lower camera; a = -2.2 ; L = 971 lb.


Figure 15.- Continued.


L-58-1648


1~ ""` 1'7'"~4:;~
'~~"






NACA RM L58E09


Upper camera; a = -1.20; L = 1,086 lb.


Lower camera; a = -1.2o; L = 1,086 lb.


L-58-1649


Figure 13.- Continued.






NACA RM L58EO9


Upper camera; a = -0.50; L = 1,126 lb.


Lower camera; a = -0.3 ; L = 1,126 lb.


Figure 13.- Concluded.


L-58-1650






NACA RM L58E09


Upper camera; a = -6.80; L = 554 lb.


Lower camera; a = -6.8; L = 554 lb. L-58-1651

Figure 14.- Deflection study photographs with additional wing guy cables
installed. V = 71 mph; qav = 12.4 lb/sq ft.






NACA RM L58E09


Upper camera; a = -4.90; L = 889 lb.


Lower camera; a = -4.90; L = 889 lb.


Figure 14.- Continued.


L-58-1652





NACA 1W L58E09


Upper camera; a = -4.10; L = 1,089 lb.


Lower camera; a = -4.10; L = 1,089 lb.


L-58-1653


Figure 14.- Continued.






NACA RM L58E09


Upper camera; a = -5.2; L = 1,500 lb.


Lower camera; a = -3.2o; L = 1,500 lb.


L-58-1654


Figure 14.- Continued.


NACA Langley Field. V.






NACA IM L58E09


Lower camera; a = -3


Upper camera; a = -.

Figure 14.- Concluded.


L-58-1655







NACA RM L58E09


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