Static test of Douglas O-2 and report on static proof test of horizontal tail surfaces (Airplane Section report)

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,..C:ONFIDENTIAI::­File
D 52.1 /Douglas 0.2 / 4 ·
AIR SERVICE INFORMATION CIRCULAR
<AVIATION)
PUBLISHED BY THE CHIEF OF AIR SERVICE, WASHINGTON, D. C.
Vol. VI February I. 1926 No. 554
STATIC TEST OF DOUGLAS 0-2
AND
I
REPORT ON STATIC PROOF TEST OF
HORIZONTAL TAIL SURF ACES
(AIRPLANE SECTION REPORT )
Prepared by E. R. Weaver
Engineering Division, Air Service
McCook Field, Dayton, Ohio
August 7, 1925
WASHINGTON
GOVERNMENf PRINTING OFFICE
1926
Ralph Brown Drmt''fifii}
LIBRARY ~
MAY 2 9 2013
Non·Depoitory
Auburn University
INDEX
FIGURE 1. Plan view. FIGURE 16. Drawing of aileron structure.
2. Front and side elevations. 17. Loading schedule and test results.
3. Plan of upper wing} . 18. Drawing of stabilizer and elevator structure.
4.
Pla n of 1o wer w.m g wmg structures. 19. Loading schedule and test results.
5. Typical rib. 20. Drawing of rudder and fin structure.
6. Typical spar sections. 21. Loading schedule and test resultR.
7. Loading schedule (reverse flight). 22. Drawing of fuselage structure
8. Table of net spar deflections (reverse flight) . 23. Loading schedule and location of loads with
9. Deflection curves (reverse flight). deflections and test results.
10. Loading schedule (low incidence) . 24. Drawing of landing chassis.
11. Table of net spar deflections (low incidence). 25. Landing chassis test set-up and test results.
12. Deflection curves (low incidence). 26. Drawing of tail skid.
13. Loading sched.ule (high incidence). 27. Tail skid test set-up and test results.
14. Table of net spar deflections (high inci- 28. Drawing of interplane struts.
dence). 29. Perspective diagram of control system.
15. Deflection curves (hi![h incidence). 30. Interplane wire diagram.
Static proof test of horizontal control surfaces pages 29 and 30.
INDEX 0F PHOTOGRAPHS
FIGURE 31. Failure in static test of Douglas 0-2 lower
left wing stub tube joint.
FIGURE 37. Douglas 0-2 controls sticks showing bolts
which failed in bending.
32.
33.
34.
35.
36.
Failure in static test of drag truss com­pression
tube in left wing stub of
Douglas 0-2.
Douglas 0-2 wing after failure in static
test at L. F. of 8.
Three-quarter front view of Douglas 0-2
wing after failure in static test.
Front view of Douglas 0-2 wings after
failure in static test.
Failure in static test of lower right longeron
of Douglas 0-2 fuselage at junction of
rear spar truss.
38. Douglas 0-2 front control stick showing
fitting that failed during static test.
39. Failure of Douglas 0-2 wheel in dynamic
landing chassis test.
40. Douglas 0-2 dynamic landing chassis test
set-up.
41. Close-up of Douglas 0 - 2 landing chassis in
dynamic test.
CERTIFICATE: By direction of the Secretary of War the matter contained herein
is published as administrative information and is required for the proper trans­action
of the public business.
(n)
STATIC TEST OF DOUGLAS 0-2 AND REPORT ON STATIC
PROOF TEST OF HORIZONTAL TAIL SURFACES
SUMMARY OF RESULTS
Airplane: Douglas Observation 0-2.
Type: X.
Engine: Liberty 12, 400 horsepower.
Description: A two-seater biplane, steel tube fuse­lage,
wood wings, control surfaces of wood construction,
axleless landing chassis, all surfaces having fabric
covers.
Total weight: 4,620 pounds.
Wing cellule weight: 630 pounds.
Wing area : 413.45 square feet.
Date
May 18, 1925
Do _____ _
Do ___ ___ _
May 19, 1925
Do __ ___ _
Do ______ _
May 26-27,
1925
Do ______ _
May 29, 1925
May 28, 1925
May 26, 1925
May 4, 1925
June 2, 1925
June 3, 1925
June 5, 1925
Part tested Load
required
RESULTS OF TESTS
Pounds per
square foot
or factor
supported
Failed at Weight
Horizontal stabi· 30 pounds 40 pounds --------- - -- -- 0.94 pounds
Jizer. per sq.ft. per sq. ft. per sq.ft.
Elevator _________ _ _____ do ___________ do ____ ___ -- ------- ----- 0.81 pounds
per sq.fl
Elevator control_ ______ do-··--- -1-----do .. _____ -·------------ -------------·
Vertical fin ________ 25 pounds 35 pounds ---------- -·-- 1.6 pounds
per sq. ft. per sq. ft. · per sq. ft.
Rudder_ ___ ___ _________ do __ _____ l _____ do _______ - ------ - ----- - 0.75 pounds
Rudder controL ___ ___ do ___ ______ ___ d0 ____ · ___ -- --- ----- ---- ---~~~-~~:_f\
Ailerons __________ 30 pounds 30 pounds -----------·-- 0.92 pounds
per sq. ft. , per sq. ft. per sq. ft.
Aileron control. ________ do _ _____ 27.5 pounds 30 pounds -------·---- - -
per sq. ft. per sq. ft.
WING CELLULE
Failure
Surface structurally satisfactory.
Do.
Elevator controls satisfactory.
Surface structurally satisfactory.
Do.
Rudder controls satisfactory.
Aileron surfaces are satisfactory, but controls are
faulty ,
Bolt at lower nd of the front control stick failed in
bending. Co)lllecting tube fitting failed.
High incidence. ___ L. F . 8.5 ____ 7. 5 _________ 8. o __ __ ! _____ 1.52 pounds Left lower compression tube and stub wing stub
per sq. ft. . failed.
Low incidence ___ _
Reverse load_. ___ _
Six·foot length of
leading edge.
Fuselage_---------
Tail skid _________ _
Chassis:
5. 5 _________ 5. 5 _ ___ _____ -------------- _____ do ___ ____ Test satisfactory.
3. 5 _____ ____ 3. 5 _______ __ - --- - -·------ - _____ do_._____ Do.
10. o _________ 24. o __ _______ 24. !_ ___ _____ - -----·------ - Structurally satisfactory.
5. 5 __ _____ __ 6. 7 ·--------- ---- --- ---· -- - 317.5 pounds_ Front part of structure loaded to L. F. 6.0; rear part
loaded to 7 .0. Lower longerons showed slight
failure.
3().inch drop_ 42-inch drop_--- -- ------ --- -------------- Satisfactory structurally.
SAtxrlue t_s_._-_-_-_-_-_-_-_- }4 2-inchdrop_ . {Wheel failed at 33-inch drop. 36 by 8 wheels and
51-mch drop_--------- ----- 154 pounds__ tires were fitted, after which no failure occurred.
Shock absorber
Discussion: With the exception of the left rear spar
stub truss, the structure of this airplane passed all
Engineering Division requirements. The aileron con­trol
system should be improved and strengthened.
The landing chassis shock absorbers functioned very
well and are considered very efficient.
WITNESSES
Capt. G. E. Brower, all tests.
Lieut. Harry Sutton, all tests.
Mr. Eric Springer, Douglas Co. representative,
all tests.
Mr. W. E. Savage, all tests.
REPORT OF THE STATIC TEST OF THE Mr. E . R. Weaver, all tests.
DOUGLAS "0-2" AIRPLANE
OBJECT
The Douglas 0-2 was static tested for the purpose
of determining its structural strength. This airplane
was built by the Douglas Co., of Santa Monica,
Calif., and submitted to the Engineering Division for
test in accordance with contract No. 7631, Washing­ton,
D. C., and circular No. 1576-A, dated January
27, 1925.
DATE AND PLACE
The component parts of this airplane were all tested
at McCook Field, Dayton, Ohio, on the dates as
shown under "Results of tests."'
GENERAL DESCRIPTION
The Douglas 0-2 is a two-passenger biplane to be
used for observation purposes. This airplane has a
single bay wing cellule with double flying wires and
straight streamline duralumin interplane struts with
streamline incidence wires. The wing panels and all
the control surfaces are of wood construction with
fabric covers.
The fuselage is a chrome-molybdenum tube structure
with . torch-welded joints. The diagonal bracing is
accomplished by round swaged rods.
The landing chassis is a split-axle type chassis with
the shock absorbers mounted in each vertical strut;
(1)
32 by 6 tires and wheels were used on this landing
chassis.
The total weight of the airplane as in flight is 4,620
pounds.
The useful load is 1, 7go pou ncls.
The engine .used is a Liberty 12, 400 horsepower.
Weight per square foot of wing area= 11.2 pounds.
Weight per horsepower= 11.5 pounds.
Airfoil, Clark Y.
Figure 1 is a plan (drawing) of the airplane.
Figure 2 is a side and encl elevation drawing of the
airplane.
WING CELLUJ,E
The Douglas 0-2 wing cellule consists of two upper
panels, jointed on the center line (or at the center of the
airplane), and two lower panels, joined to a stub
center section which extends out 17 inches on either
side of the fuselage.
Four straight cabane struts and four long outer
struts position the wings.
2
The wings were loaded in accordance with the
loading schedule in Figure 1.
The required load factor for the reversed flight
condition is 3.5.
RESUJ,TS
The wings supported a load equivalent to a factor of
3.5 without failure. The wing spars deflected very
little.
Figu.re g is a table of net spar deflections.
Figure 9 is the deflection curves.
CONCLUSIONS
The wing cellule is structurally satisfactory for the
reversed flight condition.
PUOCEDURE FOR TEST--LOW INCIDENCE
The airplane was inverted and reset in the test jig
so that the wings could be loaded on the bottom side.
The angle of inclination (ex) at which the wings
were set for the low-incidence test was go 30' with
The wings are set (or rigged) without stagger.
The wing spars are spruce, routed " I " section
tween panel points.
trailing edge down.
be- The inclination angle was determined as follows:
The wing ribs are a subdivided Warren truss type,
built up of rectangular spruce strips with plywood
gusseted joints.
Each lower panel thickens near the stub center
section, in order that the gasoline tanks may be con­cealed
within the panels.
The airfoil used in the construction of the wing
panels is the Clark Y.
Figure 3 is a plan of the upper wing structure.
Figure 4 is a plan of the lower wing structure.
Figure 5 is a drawing of a typical rib.
Figure 6 is a drawing of the typical spar sections.
Figure 2g is a drawing of the interplane struts.
Figure 29 is a perspective diagram of the control
system.
Figure 30 is a drawing showing the interplane wire
diagram.
The drag bracing is accomplished by round wires in
both upper and lower panels, with the exception of
the inner bays of the lower panels, where a triangular
truss of duralumin tubing is attached to the rear spars
to take care of the drag component at this point.
This means of bracing was executed as a result of
mounting the gasoline tanks in the wing panels between
the spars.
PROCEDURE FOR TEST--REVERSED FLIGHT
The airplane was assembled as for flight and sup­ported
in an especially constructed jig in the right­side-
up position.
The wings were set so that the lower surface was
parallel to the floor, or horizontal. I nclination angle
0°, by authority of the Handbook for Airplane De­signers,
page g1, Section II, Part VII, paragraph 12.
The center of gravity of the load with respect to
the airfoil is located at 31 per cent of the wing cliord,
which corresponds to the center of pressure when
measured from the leading edge.
For the Clark Y airfoil the ratio of net cbord com­ponent
to net beam component is 0.15, and the center
of pressure located at 51 per cent of the wing chord,
measured from the leading edge. See paragraph 15,
under table No. 2, Section II, Part I, page 17, of the
Handbook for Airplane Designers.
See also paragraph 11, Section II, Part VII; pages
g3-g6,
For tan 0.15 and the angle of inclination=g0 30'.
The center of gravity of the applied load corre­sponded
to the center of pressure of the airfoil.
The required load factor for the low-incidence
condition was 5.5, and the load was applied as indicated
in Figure 10.
RESULTS
The wing structure supported the required load
without sign of failure. The wing spar deflections were
very slight.
Figure 11 is a table of net wing spar deflections, and
Figure 12 is the net wing spar deflection curves.
CONCLUSIONS
The wing structure is satisfactory structurally for the
low-incidence condition.
PROCEDURE FOR TEST--HIGH INCIDENCE
The airplane was reset in the test jig for the high­incidence
test.
The wings were inclined to the horizontal at go 21',
with the leading edge down. The angle of inclination
was determined in the following way:
The ratio of net chord component to net beam
component is 0.144 .
• •• 'Y or the inclination angle is the angle whose tangent
is 0.144, or go 12'.
The center of gravity of the load is located at 31 per
cent of the wing chord, measured from the leading
edge. This corresponds to the location of the center
of pressure of the airfoil for the high-incidence condi­t
ion.
The required load factor for the high-incidence test
is 8.5.
The wings were loaded in accordance with the loading
schedule in Figure 13.
RESULTS
The wing structure supported a load equivalent to a
factor of 7.5 very satisfactorily. The failure came as
t he suppor t ing jacks were r eleased with a factor of
8 on the wings. The left lower wing stub truss and
drag strut failed due to a poor weld in the stup t russ,
together with a side bending load clue to the drag
component.
This failure meant the conclusion of t he test.
The net deflections may be seen in Figure 14.
Figure 15 is the spar-deflection curves.
See photographs in Figures 31, 32, 33, 34,
and 35.
DISCUSSION
When the wings were supporting a load
equivalent to a factor of 6 the lift wires
were bearing on the sides of the holes th~ough
the wing covers. T here should be no inter­ference
at these p laces. The vibrating length
of t he lift wires should be unrestricted.
After the fai lure of the rear truss of the
left wing stub the members which failed and
the right stub t russ drag strut were removed
and sent to the Material Section for test to
determine the nature of the material and
the quality of the weld which fai led. T he
results of t hese tests a re included in this
. report.
3
Figure M 1302-1 shows the portion of the bond which
was fairly good. There is no sharp line of demarcation
between the welding material and the chrome-molyb­denum
tubing, which shows that the two materials
were properly fused. Figure M 1302- 2 is a portion of
the weld in which the welding material was not fused
into the tube and was simply stuck on the surface.
It is believed that the failure was due primarily to
an overload, and that it was influenced by a poor
weld. The liability of failq re of this part in future
construction bas been lessened by changing the con­struction
so that there will be no local bending in
the welded joint, and the length of the welded seam
has been increased so that even if a portion of it is
defective the r emaining portion will develop the
strength of the tube.
FIG. M 1302- 1, PLATE 6595-A
Sections of the wing spars were also sent
to the Material Section for test.
CONCLUSION
Magnification, 100 diamete rs. Etching, alcoholic nitric acid. Remarks: Welded
tube from Douglas Observation Airplane 0 - 2. Section at weld-bond is fairly
good. Dark area chrome molybdenum tube; light area low carbon welded metal
The wing structure is structurally unsatis­factory
for the high-incidence condit ion,
having failed where only 94.12 per cent of
the designed load was imposed on the wing
cellule.
MATERIAL SECTION REPORT ON IN­VESTIGATION
OF THE FAILURE OF
THE WELDED JOINT ON TRUSS
MEMBER OF DOUGLAS "0-2"
This report covers the metallographic and
visual examination of the welded joint which
failed on the 0 - 2 airplane during t he sand­load
test.
The visual examination indicated that
the weld was improperly made, and t hat
only about 50 per cent of t he seam showed
fu sion between the welding material and
the tube. This was confirmed by metal­logrnphic
examination.
FIG. M 1302-2, PLATE 6596-A
Magnification, 100 diameters. Etching, alcoholic nitric acid. Remarks: Section
1/8 inch from that shown in Figure I. Bad bond at weld. Dark area Cr·Mo
tube; light area low carbon welded metal
MATERIAL SECTION REPORT ON LEFT WING
STUB OF DOUGLAS "0-2" AIRPLANE
PHYSICAL PROPERTIES
Test No., R. NO--- ----------- ------- --- ------
Sample No __________________________________ _
Gauge length, 15 by T ----------------------------­TWhiidctkhn
eosrs d, iianmche tgearu, gOe ._ _D_ _-_-_-_-_-_-__-_-_-_-_-_•_-_-_-_-_-_-_-_-_-_-_--_-_
Yield point, pounds per square inch--------------
Ultimate strength, pounds per square inch _______ _
Specification 10231 _________________ ---------------.
1 Left compression member.
CHEMICAL ANALYSIS
Carbon, 0.30 to 0.31.
8923
0. 510
1.000
• o:34
92, 100
109. 500 I Pass.
2
I 0. 510
1.000
• ffi4
87,200
lffi,100
Pass.
COMPRESSION TEST OF STUB WING DRAG
STRUT OF DOUGLAS 0-2 AIRPLANE BY MA­TERIAL
SECTION.
After the failure of the left stub wing drag strut
which is shown in photograph (fig. 32), the drag strut
from the right wing stub was tested as a pin ended
column in the Olsen testing machine.
At a load of 2,900 pounds the strut failed in com­pression.
MATERIAL SECTION REPORT ON WING BEAMS
AND COMPRESSION RIB FROM DOUGLAS
0-2 AIRPLANE
PURPOSE
To determine the strength properties of the material
in the wing beams and the strength of a compression
rib from the Dquglas 0-2 airplane.
CONCLUSIONS
The material in the wing beams and the rib was
spruce.
The moisture content was normal and the specific
gravity was above the minimum requirement.
The material in the front lower left and right and the
front upper left beams was brash.
The material in the rear upper left and the front and
rear upper right beams was good.
The material in the rear lower right and left beams
was fair.
The compression rib failed under a load of 2,900
pounds.
MATERIAL
One section, approximately 8.J1 feet long, cut from
each wing beam was submitted for test after the air­plane
had been sand tested. The beams were the " I"
type.
4
One compression rib, the design of which is shown in
Plate I , was submitted for test.
METHOD OF PROCEDURE
The upper beams were tested intact in transverse
bending under third point loading over a 96-inch span.
As a check small test specimens, approximately 0.8 by
1.3 by 30 inches, were cut from the webs and flanges of
these beams and were tested in transverse bending
under center loading over a 28-inch span .
The lower beams, being slightly tapered toward one
end, were not tested as long beams but were cut up into
small sp~cimens. These pieces, taken from the flanges
and webs of each beam, were glued up into specimens
approximately 1% by 2 by 48 inches and were tested in
transverse bending under third point loading over a
45-inch span. •
Tests were also made to determine the moisture con­tent,
the specific gravity, and compression parallel to
the grain.
The values for the modulus of rupture and fiber
stress at the elastic limit of the upper beams were
adjusted by a form factor to make them comparable
with those of rectangular beams.
The compression rib was tested in compression as
shown in Plate I. The rib was so located that the
center in the wider face of the rib was approximately
thirty-one thirty-seconds inch from the line of the
applied load.
RESULTS
The compression rib failed at 2,900 pounds by com­pression
in the flange carrying the greater stress. The
failure is shown in Plate I.
The average strength properties of the spruce in the
beams are noted in Table 1.
DISCUSSION OF RESULTS
The strength properties of the spruce in the front
lower right and left beams and in the front upper left
beam were low, with the exception that the value for
the modulus of elasticity compared favorably with
that for spruce of average grade. Th~ low values
resulted from the brash condition of the spruce, which
was not due to deterioration or decay.
The strength propert~es of the spruce in the rest of
the beams compared favorably with those for spruce of
average grade, with the exception that the values for
the modulus of rupture and compression parallel to the
grain were slightly low. The material, however, did
not show any evidence of brashness.
The compression rib failed by crushing in the flange
carrying the greater load. It -is not known what load
the rib was required to carry.
/
5
TABLE !.-Strength properties of spruce in wing beams from Douglas "0-2" airplane
. I Moisture I Location of beam content,
per cent
LowFerr olneftt _:_ _ _______ ___ ______________________ _____ - -- -- -- --
Rear ____ ________ ________________ ---- ____ ---- ____ _____ _
LowFerro rnigt.h. t_:_ _______ ___ ______ ____ ______________ __ ____ ___ __
Rear ---- ------ ---------------------------- - ---- ----- --
1100.. 9877 1
11. 20
10. 31
Upper left:
F ront.. -- - - -- ------- -- -- -- -- - --- - - --- -- - -- - --- - -- - ---- 11.48
Rear ---- ---- - ----- ______ ----- __ _______ ----- - ~- _______ _ 11. 62
Opper right:
Front ---- --- -- ---------------- -------- ---- --- - --------
Rear --- -------------- ----- _________ ----- _____ _____ ---- 11. 92 1
11.12
1 Brash material.
MATERIAL SECTION REPORT ON INTERPLANE
STRUTS OF DOUGLAS "0- 2" AIRPLANE
MATE RIAL
One set, consisting of eight metal struts from the
Douglas 0 - 2 airplane which had been static tested,
was submitted for t est. The struts were manufac­tured
from aluminum alloy tubing formed into a stream-
PLATE I
line section having diameters of 3.35 and 1.5 inches.
The end fi ttings were castings riveted to the stream­lined
tubing. The t hickness of the tube wall was 0.063
inch. (The form of the struts is shown in Plate II.)
Specific
gravity
0. 379
. 419
. 367
.378
.386
. 398
• 461
.408
Modulus of-
Rupture, Elasticity
pounds pounds
per square per square
inch inch
I 6, 720 1, 565,000
8,880 1, 858,000
I 6,860 1, 595,000
8, 810 1, 675, 000
I 7, 970 1, 591, 000
9, 655 1,606,000
11, 025 1, 922, 000
10, 120 1, 682,000
Fiber
stress at
elastic
limit,
pounds
per square
inch
4, 200
5, 600
4, 985
5, 820
5, 835
6,445
6, 830
7, 130
CONCLUSIONS
Com-press
ion
parallel
to grain,
pounds
per square
inch
4,850
5,180
4, 635
5,025
5,070
5,403
6,660
5, 695
Work to
maximum
load in
pounds
per inch '
4. 7
15. 2
8.8
13. 6
3. 9
7. 9
8. 1
8.8
The material used meets the requirements of Air
Service Specification HT-11054.
The long· struts sustained an average maximum load
of 3,800 pounds.
The short struts sustained an average maximum
load of 8,050 pounds.
METHOD OF TEST
The struts were tested as a column in a 20,000-
pound Olsen testing machine. The ends were sup­ported
on bolts through the holes in the fittings .
. RESULTS
The test results are as follows:
Long struts 1, 3,700 pounds; 2, 3,980 pounds; 3
3,700 pounds; 4, 3,820 pounds; average, 3,800 pounds;
no failure.
Short struts 5, 7,750 pounds; 6, 8,250 pounds; 7,
8,700 pounds; 8, 7,500 pounds; average, 8,050 pounds;
end bolt bent and holes somewhat elongated.
DISCUSSION OF RESULTS
The load sustained by the long struts was approxi­mately
the Euler load for these columns.
Due to the fact that the bolts in the two ends of the
shorter struts were in different planes, one end was
slightly fixed, which resulted in a total load somewhat
higher than if the end bolts had been in the same plane,
as was the case with the longer struts.
The results of the physical tests and chemical
analysis of the strut material are as follows:
Physical tests:
Ultimate strength, 58,000 pounds per square
inch.
Yield point, 46,300 pounds per square inch.
Elongation in 2 inches, 19 per cent.
Elongation in 4 inches, 14.75 per cent.
Chemical analysis:
Copper, 3.40 per cent.
Silicon, 0.36 per cent.
Iron, 0.48 per cent.
Magnesium, 0.48 per cent.
Manganese, 0.52 per cent. ·
Aluminum, difference.
AILERONS
DESCRIPTION
The Douglas 0 - 2 airplane has four ailerons, two on
the upper wing and two on the lower wing, with a con­tinuous
cable control system. The ailerons them­selves
are built on one large box spar, to which the
rib and hinges are fastened. The trailing edge is a
PLATE II
V-shaped duralumin channel. The ailerons are fabric
covered.
Figure 16 is a drawing of the aileron structure.
The area of each aileron is 12.5 square feet.
The weight of each aileron is 11.5 pounds.
The weight per square foot of surface is 0.92 pound.
PROC~DURE FOR TEST
The upper and lower left aileron were made ready
for test with the airplane in flying position. A spring
balance was coupled to the control stick to register
6
the force necessary to move the surfaces while under
load.
The required load the 0 - 2 ailerons must support
without failure is 30 pounds per square foot.
RESULTS
Figure 17 gives the loading schedule and the test
data.
The aileron surfaces supported the required load,
although the control cable stretched out so that the
loads would slide off the ailerons. This was in part
due to the masts on the ailerons being too short.
The ailerons were tested by using both front and rear
control sticks.
The fir t failure in the front control stick assembly
was that of the five-sixteenths bolt in the control
stick yoke. This bolt failed in bending.
Test was again resumed and the load was trans­ferred
to the rear control stick.
At a load of 20 pounds per square foot the two con­trol
sticks were out of alignment 7.Yl! inches at the
pilot's grip. After examination it was noted that the
fitting on the front control stick at the torque tube
had failed. Figure 38 is a photograph of this failure.
Figure 31 is a photograph of the control stick assembly
showing bolts which failed.
In order to load 30 pounds per quare foot on the
ailerons it was necessary to again transfer the load to
the front control stick. The two control sticks were
then 11 inches out of alignment.
RECOMMENDATIONS
Use heavier bolts in the control stick yokes and
eliminate stretch in control system. The control
system should be designed a.ad built so that the aile­rons
on one side can be operated independent of the
ailerons on the other side of the airplane.
Make control stick to torque tube fitting strong
enough to take care of a 30 pounds per square 'foot.
load on the control surfaces.
ELEVATOR AND STABILIZER
DESCRIPTION
The elevator structure is a combination of wood and
duralumin. The main tube or spar is a duralumin
tube to which the plywood ribs are fastened by means
of special fittings riveted to both ribs and spar. Each
fitting fo rms a socket for a rib.
The trailing edge is a dural channel.
The stabilizer is a wood structure built on two spruce
spars routed " I " section.
The ribs are plywood, with spruce cap strips, and
the drag bracing is accomplished by diagonal wires.
The elevator is hinged to the rear spar on micarta
hinges which surround the main spar tube of the
elevator.
Figure 18 is a drawing of the stabilizer and elevator
structure.
PROCEDURE FOR TEST
The elevator and stabilizer were assembled on the
fuselage as for flight and the ·urfaces loaded according
to the load schedule in Figure 19.
The center of gravity of the elevator load was
located at five-twelfths of the chord measured from the
hinge center, due to the fact that the load at the
trailing edge was one-third of the load at the hinge.
Deflections were measured as indicated in Figure 19.
A spring balance was coupled in the control system
to measure the force required to move the elevator
when loaded.
The required load the horizontal tail surfaces must
support without failure is 30 pounds per square foot.
The area of the stabilizer= 33.6 square feet.I
The area of the elevator=21.4 square feet. both
The weight of the stabilizer=31.72 pounds. sides.
The weight of the elevator= 17 .28 pounds.
The weight per square foot of stabilizer= 0.94
pounds.
The weight per sq uare foot of elevator = 0.81 pounds.
RESULTS
The horizontal tail surfaces supported the required
load and an additional 33 per cent overload without
failure.
After the test the elevator control system was care­fully
inspected and it was discovered that the bolt
holding the lower end of the rear control stick failed
in bending.
Figure I!) shows the deflections.
Rl~COM MENDATIONS
Use heav ier bolt in rear control-stick yoke.
CONCULSlON
The horizontal tail surfaces arc structurally satis­factory.
RUDDER AND FIN
DESCRIPTION
The rudder construction is the same as that of the
elevator. The only difference in the surfaces is the
general shape, the size, and the length of the masts.
The fin is a wood structure, with the exception of
the leading edge, which is a curved steel tube. The
ribs are plywood and have spruce cap strips.
Figure 20 i& a drawing of the rudder and fin stru cture.
The area of the rudder= 10.3 square feet.
The area of the fin=7.9 square feet.
The weight of the rudder = 7.8 pounds.
The weight of the fin =8.4 pounds.
The weight per square foot of surface of rudder = 0.75
pounds.
The weight per sq uare foot of surface of fin = 1.60
pounds.
The required load for the vertical tail su rfaces is 25
po11 nds per sq uarc foot.
PROCEDURE FOR TEST
The rudder and fin were assembled on the fuselage,
and the fuselage supported on its side. A spring bal­ance
was coupled to the rudder bar to measure the
force necessary to move the rudder with the load on.
The load was applied as indicated in the load schedule
in Figure 21.
79026-26t--2
7
The center of gravity of the rudder load is located at
five-twelfths of the chord, measured from the hinge
center line. This is due to the fact that the load at
the trailing edge is one-third of the load at the hinge.
RESULTS
The rudder and fin supported 35 pounds per square
foot, or a 40 per cent overload, without failure. Both
rudder bar and pedals supported the required load
without failure.
In Figure 21 may be seen the deflections.
CONCLUSION
The rudder and fin and the rudder controls are struc­turally
satisfactory.
RECOMMENDATION
None.
FUSELAGE
DESCRIPTION
The Douglas 0-2 fuselage is a chrome-molybdenum
steel tube structure with welded joints. The engine
mounting is also a tube structure and is detachable.
The bracing of the structure is accomplished by
wires throughout.
Figure 22 is a drawing of the fuselage structure.
PROCEDURE FOR TEST
The fuselage, with the wings mounted on, was set
in an especially constructed test jig and loaded as indi­cated
in Figure 2:l, which gives the location and mag­nitude
of the various loads imposed.
The required load factor for this fuselage is ~.5.
RESULTS
The engine mount supported a load equivalent to a
factor of 6 without failure. The rear portion of the
structure supported a load factor of 6.5, and when the
equivalent of a load factor of 7 was on the lower fon­gerons
in the bay immediately to the rear of the rear
spar truss showed signs of failure in compression. -
Figure 23 gives all of the test results.
Figure 36 is a picture of the failure.
CONCLUSION
The fuselage is structurally satisfactory.
LANDING CHASSIS, DYNAMIC TEST
DESCRIPTION
The landing chassis of the Douglas 0-2 airplane is
a molybdenum steel axleless landing chassis wi,th two­unit,
of double-wound shock absorbers built in · each
outer front strut.
A fairing protects and neatly streamlines each
shock-absorber unit.
The weight of the landing chassis complete with
wheels, but without fairing, is 154 pounds.
The size of the wheels is 32 by 6, and the weight of
the two wheels and tires is 68 pounds.
Figure 24 is a drawing of the landing-chassis struc­ture.
\
PROCEDURE FOR TEST
The landing chassis was mounted on the fuselage as
for flight, and the rear portion of the fuselage elevated
so that the center of gravity of the load would be
vertically over the axle center line.
The wheels were placed on a platform which was
incli ne::! to one side go 27' .
In testing the landing chassis this way a condition
of combined down load and side thrust was simulated.
The fuselage and engine mount were loaded so that
the reaction at the tires was 2,135 pounds, or one-half
the weight of the airplane (loaded for flight), in the
tail high position. The tire pressure was 75 pounds.
The required distance of drop for this landing chassis
with one-half load is 42 inches.
Figures 40 and 41 are photographs of the landing
chassis test set-up.
RESULTS
The test was begun by dropping the chassis 6 inches,
then 12, 18, 24, etc. After being dropped 33 inches
the hub was torn out of the right wheel, and heavier
wheels and 36 by 8 tires had to be fitted. The air
pressure in the tires was 75 pounds.
Dropping the chassis co ntinued until the maximum
clistance was 51 inches, afLer which the test was clis­continued.
The chassis structm e showed no sign of fa ilure and
the shock absorbers functioued very satisfactorily .
Figure 25 gives all the da.ta of the landing-chassis
tests. Figure 39 is a photograph of the wheel failure.
CONCI,USION
The landing-chassis structure is structurally satis­factory.
TAIL SKID
DESCRIPTION
The tail skid is a steel tube, reiiiforced where nec­essary,
shackled to the shock-absorber elastic at the
upper end and fulcrumed at the middle.
From the swivel on down to the tail-skid shoe the
tube is curved, while the part above the swivel is
straight.
Figure 26 is a drawing of the tail skid .
PROCEDURE FOR TEST-DYNAMIC TEST
The tail skid was assembled in the fuselage as for
flight. A hoist and trip was attached to the structure
a11d the tail skid allowed to come to rest after each
drop, on a platform inclined to one side at go 27'.
This was done to simulate a combined down load and
8
side thrust. The rear portion of the fuselage was
loaded so that the reaction at the tail-skid shoe was
563 pounds, which is the load on the tail skid in the
landing position.
The required height of drop for this tail skid is 30
inches.
RESULTS
The Douglas 0-2 tail skid withstood a <;lrop through
a distance of 42 inches without failure or damage to
the fu selage structure.
CONCLUSION
The tail. skid and tail-skid mounting in the fuselage
is structurally satisfactory.
LEADING EDGE
PROCEDURE FOR TEST
A 6-foot wing section was cut from the upper panel
and supported on the spars for test. It was necessary
to make only one test, because both upper and lower
panels were the same airfoil and had the same chord .
The test constituted loading the leading edge, the
portion just ahead of the front spar, with lead shot unti l
fa ilure, and then computing to find the factor at
which the st.ructure failed.
RESULTS
The total load on leading edge at t ime of failure was
4,150 pounds.
The factor at which failure occurred was computed
in the following way :
Weight of airplane in flight=4,620 pounds.
Span =37. 75-feet.
Leading edge in terms of chord =8.825 per cent.
Part of load carried by upper wing=53 per cent.
W = load on 1 foot run of test section per load factor.
=4,620X .53X
37.75 8 . 82 5
20
=28.6 pounds.
W = 172 pounds or load considered as one factor.
The load on the leading edge at time of failure was
4,150 pounds.
Therefore, the load factor at which the leading edge
failed is 4,150 =
172 24 .1 .
The required load factor for the leading edge is 10.
CONCLUSION
The leading edge is structurally satisfactory, and is
141 per cent stronger than the requirement.
9
Fm. !.-Plan view
JSR41t:f9! 1oj-#-------'
FIG. 2.-Front and s1· de vi.e ws
10
Jaax. 039· VPPER J.i. QD.X. 038 LOWER
FIG. 3.:..T:: pP!'r wing
--~--'<+--~/08~~~-
1/ o.n.x. 052 .DORAL Tl/BE
I~ QD. X. 03C3 UPPER_
4 . a .D.X. 03.9 LOWER
F IG. 4.- Lower wing
~zjsPRUCE
TYP.ICAL RIB
WhYG POS/TID
li'O()T
UPPE,P ll!DDLE
TIP
FI G. 5.- 'J'ypical r ih
JWNG PtJJ/Tltl
HOOT
GUSSE7:5z8PLY MSHOGANY
E4CE POPLAR CORE
LOWEil If/DOLE
..._~-J:=:..J>==i--'-==-'-
T IP
F IG. 6.-T ypical spar seetions (full size)
,
12
FROJY. £:f!tY2. \
I
L,,,.,l-1-45;!-~ ,,,,,i+ 43 ,i-~ 43: '~"
UPPER W/IYG
L OADlNG SCH EDULE IN POUNDS
Load
Lower wing I Gppcr wing Lower wing I Upper wing
Total
factor Front Rear I Front--Re;;;:- Front - Rear I Front Rear load -m 980 -----;o--;o------;,;- -;o-1--;o ------
2.0 860 7, 350
2. 5 I, 240 I, 245 I I, 095 I, 095 I, i40 I, 245 I I, 095 I, 095 9, 350
3.0 I, 500 I, 510 I, 330 I, 330 I, 500 I, 510 I, 330 I, 330 II, 350
3. 5 I, 770 I, 775 , I, 565 I, 565 1 I, 770 I. 775 I, 565 I, 565 13, 350
FIG. 7.-Reverse flight
F
--___.........+----+---·-L--l--------- ~I ,M .N
6/ ----60 __, ___ 1,+ 5w /.j.
TABLE OF NET SPAR DEFLECTIONS
Load
hctor A B c
Deflections in
G H
inches
..
I J
--
o. 3 0. 4
.5 . 7
.8 . 9
1. 0 1. 1
2o Q.3 ~ -0.S o.-;-~ ~ o.610:6-- 2. 5 . 5 . 7 1.0 .4 .6 1. 2 . 8 . 7
3. 0 . 8 1. 0 1. 3 . 5 . 7 I. 4 1. 0 0. 8
~. 5 . 9 I. 2 I. 6 . 6 . 8 I. 6 !. 1 I. 0
K L
----
0. 6 o. 2
.8 .3
I.I A
1. 3 . 5
FIG. 8.-Reverse flight test
REAR SPAR I
LQA/J FAClVR
~­.
,2.5-
L
r
E
M
M
--
0. 5
. 5
. 7
. 7
N
- -
0.8
1.0
1.1
1. 4
F'
FIG. 9.- R everse flight test deflection curves
0
- -
0. 6
. 7
.8
.9
p
--
0.5
.6
• 7
.8
G
Tip travel
Left Right
Upper Lower Upper Lower
--------
+0.2
+.2
+.1
+. l
+0. 1 -0.2 -0. I
+.1 - .3 - . 1
+. 2 - .4 .0
+. l - .4 .0
P REFEREM:E Lll'h
LOAD FACTOR
13
' ,,.
,~ I FRO/'fT L04IJ
~
' FRO/'II' LOAD '
+- , ,~ ~ \ .. REAR !LOAD
'
,.____ -
-- 1tj_ - ~1it-l-41t"f{- l-4111- -4~(+ -41;1_._j
~ "REAR LOAD V
[,;,+45,.{+~{ +45,[-+ -"';{~~-
LOWER. M'JYG UPPER W//'fG
LOADLNG SCHEDULE LN POUNDS
·- ---------
Lower wing Upper wing Lower wing Upper wing [,oad Total I
factor load
F ront Rear Front Rear Front Rear F ront Rear
-- --- --------- ------- --· - -----
3. 0 1, 330 1,335 1, 500 1, 505 1, 330 I, 335 1, 500 1, 505 II , 340
4. 0 1,800 1, 805 2, 030 2,035 1,800 1,805 2, 030 2, 035 15, 340
4. ~ 2,035 2, 040 2, 295 2, 300 2, 035 2, 040 2, 295 2, 300 17, 340
5. 0 2,270 2, 275 2, 560 2, 565 2, 270 2, 275 2, 560 2, 565 19, 340
5. 5 2, 505 2, 510 2,825 2, 830 2, 505 2, 510 2, 825 2, 830 21, 340
FIG. 10.-Low incidence
T ABLE OF NET SPAR DEFLECTIONS
Dcfleclions in inches Tip travel
L,oad I
I I I I Left Right
f 1ctor
A B c D E l' G ll l J K L M N 0 p Q R s T ~ I ~ ;;; ;; I " " ~1~ "'
c. I c. 0 c. "0 ' I o ....i p ...:i - -- --- --- - -- -- -- - - -- - -------- --------- ---
3. 0 'T" o. 5 0. l 0. l o. 4 0. 5 0. 6 l. 0 0.9 1. 0 o. 2 0. 2 0. 2 o. 3 l. l I.I 1. 3 0.0 + 1. 1 0.0 + o. 6
4. 0 . 8 . 7 .5 . 2 . I . 0 . 3 .6 . 8 I. 0 I. 5 I. 2 l. 3 . 4 . 2 . 4 .4 I. 6 I. 6 l. 9 .o + 1. 4 .0 + 1. 1
4. 5 . 9 . 7 .6 .~ •
. l . 0 . 3 . 7 I. 0 I. 2 I. 7 I. 4 I. 4 . 4 .2 . 4 . 5 l. 8 l. 8 2. 3 +. I + 1. 8 - . I + 1. 7
.';. 0 . 9 . 8 . 7 . I .o . 4 . 8 I. I I. 4 2. 0 I. 5 t. ll . 5, .2 . 4 . 6 2. 0 2.0 2. 5 +. I +2. 0 - .I + 1. 7
.5. 5 l. I . 9 .8 . 3 . 2 . I . 4 . 9 I. 2 I. 5 2. 2 1. 8 1. 8 . 5 . 2 . 4 . 6 2. 2 2. 3 2.8 +. I +2. 3 -. 2 + 1. 7
0 after
test._. . o I . I . I . 0 . I . o I .2 . I . 2 . 2 . 2 . 2 . 2 . 1 / . I .2 .2 . 4 . 3 . 4 . 0 +. 2 .0 +. I I I '
FIG. 11.- Low incidence test
FRDNrSPARA
LCIAD FACTOR
REARSPAR K
LQ41J FACTOR
3.-
• ' 1\il/
'~~f ~
+
B c
L M
F.RCJZYT LOAD
-RE-A-R £0AD
14
H
R
FIG. 12.-Low incidence test deflection curves
I
s
J REFERENCE LINE
LQ4D FACTOR
3.
T £EFER£NCE LllYE
£ a4IJ FYJCTOR
0.
4.
-4.5
-5.
-.5.6
FRONT LOAD
.REAR
''" - 111C1.!:fl-41d41.&r_j__41§.l. 414[_ /6 16 /6 /6 /.G
LO/v7i'R WING
Load
factor
- -
4. 0
5. 0
6. 0
7. 0
7. ;,
8. 0
8. 5
Lower wing
Front Rear
--- - --
1,800 1,805
2, 270 2, 275
2, 740 2, i45
:1, 210 3, 215
3, 445 a, 4.r,o
3, 680 3, (i8[)
3, 915 3, U20
UPPER W'/NG
LOADING SCHED ULE IN POUNDS
Upper wing Lower wing
Front Rear Front Rear
--- ------ - --
2.mo 2. 03f1 1,800 I, 805
2, "60 2. ;->IJ!l 2, 270 2, 27!;
3, 000 3. 0% 2, 740 2, 74:j
3, 620 3, H25 :1, 21U ~i . 21 !)
3,88) 3, 890 ;;, 415 3, 450
4, 1:.0 4, 15!l 3, 080 :l, (i8!i
4, 415 4, 420 3, 915 3, 920
FIG. 13.-High incidence
Upper wiug
I Front
1
Jlear
--- - - -
2.mo 2, (l:j;;
2, !"l60 2, f65
3, O<JO 3. 0\)5
3, n20 :3. fi2fi
3,88r. 3, \J<Jll
4, "'° 4; l !i.1
4, 415 4,420
/'
T otal
load
---
l f1, J40
rn, 340
2:1, :140
27. 340
29, 340
3 1. 340
;;3,340
15
c D ,F rc·
-~ --l- - ~- , J ,K ,L -h--t- ,N . 0
88 -+--57 --1-- 68
TABLE 9F NET SPAR DEFLECTIONS
Deflections in inches
Load I
fa ctor I
A B c D E F G H I J K L M N 0 p
---- -- ---- ----- --- - -----i- --------
4. 0 I. 4 I.I I. 0 0. 3 0. 4 I.I 1. 1 I. 4 1.3 I.I u I o:i 0.5 0.8
0 I 0.9
.5. 0 1.9 I. 5 1. 3 . 5 . 5 I. 4 I. 5 I. 9 1. 6 I. 4 .6 I. 2 2. 2 1.3
6. 0 2. 3 l. 8 1. 5 .5 .6 l. 7 I. 9 2. 4 2. 0 I. 6 l. 5 . 7 . 7 I. 5 2. 7 1. 8
7. 0 2. 7 2.1 I. 8 .6 . 7 2. I 2. 4 3. 0 2. 3 l . 9 I. 7 . 8 .8 1.9 3.1 I. 8
7. 5 3. I 2. 3 I. 9 . 6 . 8 2. 3 2. 6 3. 3 2. 6 2. I I. 8 . 8 .8 1. 9 3. 3 2. 6
8. iJ Left rear low er spar tube failed in tension and bending, due to buckling of compression.
8. 5 Tube in drag truss.
FRON7"SR4R A
LQ<ID FACTCX!
REAR SPAR
LOAD fl'ICTOR 7.-
7.5-
B
79066-26-- 3
FIG. 14.-High incidence
c D E F G
FIG. 15.-Higb incidence t~st de!lcction curves
Tip travel
Left Right
Upper Lower Upper Lower
----- - --
-0.4
-.5
- . 7
-. 7
- . 7
- 0. 5 + 0.2 o. 0
-.7 +. 2 - . 2
- .9 + . 2 - .4
-1. 2 +.I - .5
-1.4 +. 1 - . 8
H REFERENCE LlJYE
LOAD FACTOR
REFERENCE l/1'1£
FACroR
16
1EOLT
~--
Load
Pounds Pounds
per square on each
loot aileron
0.0
5. 0
10. 0
. Ci
5. 0
10. 0
.0
15.0
20.0
22. 5
25. 0
27. 5
. 0
5. 0
10.0
15. 0
20.0
22. 5
25. 0
27. 5
30. 0
. 0
. 0
5. 0
10.0
15. 0
20.0
20. 0
25. 0
27. 5
30. 0
30. 0
. 0
-o
~2
124
0
62
124
0
248
279
310
341
0
62
124
186
248
279
310
341
372
0
0
~ -.JPLY ~HOGANY FACES POPLAR CORE
j-8 PLY SPRUCE .lJ£4PHRAi'1SAT RIBS
SOLID SRRUCE BLOCKS AT HINGES
Pullon
stick,
Deflections
FIG. 16.-Aileron
pounds Lower lert Upper left Lower right Upper right
o o __________ o ____ ______ ------------ _________ __ _
15 lYs down __ IVs down _ ------------------------
450 -o- -_ _- -_-__-_-_-_- -_-_- o--_ -__-_-_-_--_-_-__- ------------------------ -__-_--_-_-_-_-__-_-_- -_
25 Ys down ___ % down ___ - -------------- - --------
40 3 down ____ 2!){ down __ ---------------- - --- --- -
IJEFLECTJOIYS
TAKE/f HERE
Rema rks
0 3!){ up _____ 3!){ up _____ % down __ -- --- - - - ---- With stick hard over to left.
70 2H down __ , 2)4 down __ I ~ down __ - - - ---- -----
4)4 down __ 3~ down __ 1% down __ -- -------- --
5Vs down __ 5 down ____ l down ____ ------ - - - - - -
6)4 down __ 5~ down __ H down _ _ --- --- - - ---- 3n up _____ 3% up _____ 3)4 down __ ---- - --- ---- Load slid off. Deflections show vositions without load.
O - - -- -- -- ---- ------------ 0---------- 0---- - - ---- Front control s tick used .
25 ------------ ------------ IYs down __ !,\down _
45 - ----------- ---- -------- 2)4 down __ 2yg down__ ,
105 ------------ ------- - -- -- 5)4 down __ 4tt d own __
g~ :::::::::::: :::::::::::: ~~ ~~:~:: ~~ ~~:~::!
115 - • ---------- ---------- - - 21'.; down __ l Ys down __ Load slid off. Stick pulled hard o,·er and fast ened. D e- l flections taken a lter reloading.
---------- - -- -- ------- -------- -- -- 2ti down __ 2h down __
--------- - ------------ ---- ------- - 3Ys down __ 2,i. down __ f.-holt heading in front control stick .
------ --- - ------------ ------ -- ---- H down __ _ n down __ _ Load off .
O -- ---------- -------- -- -- o ______ ____ 0----- -- --- Rear control stick used . (13)4" from longeron.)
25 ------------ ------------1l ~ down __ l~ down __
40 ------------ ------------ 2n down __ 2H down __
60 ----------- - ------ - ----- 3Ys down __ 4n down __
80 ------------ --------- - -- - --------- - - - - -------- --
% up_____ _ ! yg up ____ _
,\up ____ _ Vs up _____ _
,\down _ _ J4 up ___ __ n down __ _ o _____ ____ _
5Ys down __ 4% down __
!. up ___ __ _ Vs up _____ _
FIG. 17.-Aileron tests
Control sti ck 7.5 inches out or alignment at pilots grip.
Load slid off. (Fitting slipped on front control s ti<'k
torque tube.)
All slack ta ken up by turnbuckles. Load trans ferred t.o
front stick. Control stick 11 inches out of alignment .
R equired load .
Stick 16~ inches from longeron .
Load off. Stick 16 ~ inches from longeron .
17
,__----------------~-- 85;z -~f SPRVCE
-9 _,__.9--9,/:-f----9 -i-- 9 -r----.9 -r--- 9J - :Zr8 - -i~ .JPLYWE:lJ
l*K~~~- t===tl===l==t~-~~*=~-~-~~==i==*=t===*'~~i_i /
/)I ~I'--.. /Vii ~
ll=At -/ ~tv #=4F=O===tl===*=I=:= -r "7 <~1 ~\
-----<>---<\\...__-""~-'7! ~' 6' Ill ,,, " -2 \// II' ~
FIG. 18.-Stabilizer and elevator
Pounds
per
squaro
feet
Loads in pounds
Stabilizer Elevator
I
A
L
H
B I c
~j ill ~~ ~ 0
:I I
0j 22. 5 i76 360 0 . 2 I . G
25. 0 870 400 0 . 2 . 7
30. 0 1058 480 0 . 4 . 9
A
G F E
De!lections in inches
D 1,;
- -I -
F I G
-- - - - - - - --
0. 6 0. ·I - 0. l 0. 3 0.5
I. 5 l. 3 .(j LO l. 4
2. 8 2.3 l. 4 I. 8 2. 7
3. 6 3. 3 2. I 2 5 3. 5
4. 3 3.8 2. 5 2. 8 3. u
4. 9 4. 4 2. u 3. I 4. 3
5. 6 5. I 3. 4 3. p 4.& r.. 2 5. 7 3. 9 3. 9 5. 3
T
--- - --
0. 7 o. 2 0. 0
I. 7 . 4 .1
3. 0 . 6 . 3
4. 0 .8 .4
4. 4 . 9 . 3
4. 9 l. 0 .5
5. 5 l. 0 .5
6. 0 LI . 6
~=
- - ---
0. 0 20
. 1 .A5
. I 70
.1 95
. I LIO
.2 120
. 2 145
I I . 2 155 27. 5 964 410 o . 3 I . 8
--'-~---'---'------'-----'----'-----'---'-----'---'----'---~~
:J2. 5. 35.0, 37.5, and 40.0 held 0 . K. μuJI ou sL ick 170, 180, 19:,, anti 22.1, rcspecti\·cly .
FIG. 19.-Stahilizcr and ele>ator
18
FIG. 20.-Fin and rudder
D
c
E
B
A
LOADING SCHEDULE AND TEST RESULTS
Pull on
rudder Remarks
Pounds
per
square
feet
Loads in poundsl-----De_tiec_t_io_n_s_i_n_i_nch_e_s ____ ,
Fin Rudder A B C D E bar
5. 0 49 42 -0. 2 0. 0 0. 0 0. 0 0. I
10. 0 97 84 •. 2 . 6 . 7 . 3 • 2
15.0 146 126 . 4 I. I 1. 3 .3 .2
20.0 194 168 . 7 1.6 1.9 .6 .3
22. 5 218 189 . 8 1. 9 2. 2 . 8 . 2
25. 0 242 210 -- - - --- - -- --- -- - ---- -- -- -- - - -- -- - - --- - - -
30. 0 Held 0. K .
35. 0 Reid, sagging at C.
FIG. 21.-Fin_and rudder~test
50
90
135
180
20.5 Required load . Hoth .
230 Rudder bar and peda Is.
Held required load 0. K.
19
.-l.---==;~~~~~~t"""""~~~~~~~:-JU 7VP PLAN
""ih~-+---= -,,..~:--':?i'~-~~--?"!k:--- ---:r---,--,,.~- -.<!---?<;--- ~
Fm. 22.-Fuselage
--- - - 69------------------ 208 -----------
! A B c
Loads, in pounds Deflections, in inches
Load factor Remarks w, w. w, w. w, w, A B c D
------------ - - ------------ ---
2. 0 4,080 955 808 902 27,7 502 o. 6 0. 3 0.3 0.2
3.0 6,1:.\(l 1,432 1,264 1,393 416 784 .8 . 5 . 5 .6
4. 0 8,160 1, 909 1, 720 J,884 555 1,066 1.0 .6 .6 . 8 .
4. 5 9, 180 2, 147 1,948 2, 130 625 1, 207 .9 . 7 . 7 1.0
5. 0 JO, 200 1, 385 2, 176 2,376 695 1,348 .9 . 75 . 9 1.3
5. 5 11, 220 1, 623 2,404 2,622 765 1,489 .9 . 8 1.0 1. 5 Required load .
Load factors 6.0, 6.5, and 7.0 carried 0 . K. (No more weight was placed on the engine bearers.)
F10. 23.- Fuselage test data
20
:;1j
_L_ ·--~~
............. -- ------~ ---· --r-
66 FREE LENGTH
~--------43
l.//THOUT WEIGHT O/Y WHEELS
FIG. 24.-Landing gear
.213S"
Alf< PRESSURE 72.5'
STRUTS M7JIOl/T F'A!RllYG 80 #
WHEELS 68 '
ACTUAL WE/61fr 1.54 fl/#
SHOCK ABSORBER DEFLECTIONS
Drop Right I Left Remarks Drop Right Left Remarks
--- - -- --- ---
6 ir. tt 36 4-i\ 4Ys 36 by 8 wheels were fitted. T est
was continued.
12 2,\ l~ 39 4i\ 4Ys
18 2% 2% 42 4~ 4n Required drop.
24 3Ys 2~ 45 4'• 4-h
30 3% 3~ 48 4tt 4%
33 4)/g -------- Right wheel failed. Hub tore out. 51 5. 0 4 ~
FIG. 25 .-Landing chassis test set up and test results
21
o+
SECTIO!f A-A
f.pOLT
3 ff Q.D. ,r. 062 !v'AlL
SECT/O/'f B-B
_g;f QD. X. 062 lv'ALL
Fm. 26.-Tail skid
Distance Failure Remarks· Distance of drop of drop Failure Remarks
6 None. 30 None. Required drop.
12 None. 33 None.
18 None. 36 None.
21 None. 39 None.
24 None. 42 None.
'n None.
Fm. 27 .-Tail skid set up--dynamlc test
22
' J.2-S RIVET'S
1-------~.5 jf, _____ _.
19+19
32 JZ
~ ;f,-
NOTE-THE
~TERVU:. IN THE STRUTS
IS _DURALUN/JY
Fm. 28.-Wlng struts (Cull size)
FIG. 29.-Coutrol syatem
tfI I
23
JV/RES
W/RES
5
,Z-iG STREAHLl!iE WIRES
FIG. 30.-Interplane wire diagram
FIG. 31
24
FIG. 35
' I
27
FIG. 38
FIG. 39
F IG. 41
•
r
I
I ~
PROOF STATIC TEST OF THE DOUGLAS X0-2 HORIZONTAL CONTROL
SURFACES, AIRPLANE NO. P-374
OBJECT
These proof static tests were conducted for the
purpose of determining whether or not the control
surfaces were structurally satisfactory and safe to
run the performance tests.
DATE AND PLACE
November 26, 1924, stabilizer and elevator (first
' test).
November 26, 1924, ailerons.
November 26, 1924, stabilizer and elevator (second
test).
WITNESSES
Mr. Donald Douglas. Mr. W. E. Savage.
Cap~. G. E. Brower. Mr. D. R. Weaver.
Mr. D. B. Weaver.
DESCRIPTION
The elevators are attached to the stabilizer by
means of two hinges each and positioned laterally
by a tgrust bearing. The hinges are made of micarta
blocks and capped with a piece of sheet steel. To
the cast aluminum control masts the Ys-inch flexible
steel cable extends into the fuselage to the control
mechanism.
The stabilizer is adjusted at the rear spar and turns
· about the front spar. The two streamline wires extend
from an outboard point of the front stabilizer spar to
the front end of the fin and to the rear fin spar which
supports the rudder. 1;3elow the stabilizer one stream­line
wire extends froqi. the same outboard point on the
front stabilizer spar to the fuselage lower longeron.
The rear spar has a tubular strut per side for a brace,
which undergoes the same adjustment as the stabilizer.
None of the control surfaces have balanced area.
Area of stabilizers, 35 square feet.
Area of elevators, 21.2 square feet.
PROCEDURE
The fuselage was supported so that the thrust line
was horizontal. A spring balance with block and
tackle attached to the control stick was used to meas­ure
the pull required to actuate the elevators under
load. The usual procedure of suspending scales from
several points and taking deflection readings by means
of a Wye level was dispensed with because of workmen
busy with the finishing up of the assembly details
while the proof static test was being made.
RESULTS
Design requirement is an average loading of 30
pounds per square foot. Proof test requirement, 15
pounds per square foot without any sign of weakness
or failure.
Load in Load in pounds on
pounds Pounds per pull on control Remarks
square Stabi- Elevator stick foot lizer
------ ---
5.0 100 40 29 rtabilizer adjustment locked
10.0 201 81 63 in position loaded and was
12. 5 251 101 87 immovable throughout tbe
15. 0 301 121 117 j test.
The surfaces held the required load satisfactorily.
Mr. D. Douglas, with his mechanics, checked the
amount of clearance in the quadruple-thread adjust­ment
nut together with a change in design, which con­sisted
of the addition of a tubular brace to both the
right and left side of the adjusting nut which carried
the brace-strut fittings. The function of these added
braces are to position the adjustment nut so that it
can not be ro.cked or tilted by an unbalanced load on
the surfaces. An unbalanced load jammed the nut on
the screw and prevented stabilizer adjustment.
It is the writer's belief that the horizontal force com­ponents
of the diagonal struts tend to compress and
give the nut more of an elliptical shape, causing the
nut to bear heavily on opposite sides of the adjustment
screw, this distortion increasing with the load.
A second loading was deemed necessary, and at 272
pounds per square foot the stabilizer was adjustable,
but locked with a load of 5 pounds per square foot.
DISCUSSION
Due to the design of the adjustment mechanism, the
mechanical advantage was low, which is partly the
cause of the difficulty encountered when attempting
to adjust the stabilizer under a load higher than 272
pounds per square foot.
CONCLUSION
The horizontal tail surfaces are structurally satisfac­tory
for flight testing.
RECOMMENDATION
Improve the stabilizer adjustment mechanism or re­design
it so that adjustment is possible when surfaces
are supporting at least one-half of the load for which
they were designed.
(29)
•
•
AILERON TEST
DESCRIPTION
There are four ailerons in the wing cellule, one being
mounted in the outboard trailing edge of each wing
panel. They are actuated through a tandem control
hook-up, whereby the cable stretch throughout the
system is common to all ailerons. A Ys-inch diameter
flexible steel cable is used.
The control cable mast appeared to be of wood.
The hinge-pin type of hinge is used, three per aileron,
located on the lower edge of the wing.
Area of each aileron, 12 square feet.
The writer is prevented from giving a more detailed
description of the horizontal control surfaces on account
of not dissembling the surfaces at the conclusion of the
tests.
PROCEDURE
The airplane was completely assembled, and with the
thrust line horizontal the airplane was given the neces­sary
support to insure the wing cellule against any
excessive strains during the test. A block and tackle
with spring balance was attached to the control stick to
measure the force required to actuate the ailerons on
one side of the wing cellule when loaded.
The distance that the inboard tip at the trailing edge
of the aileron fell short of returning to its neutral posi­tion
at zero load was measured when the control stick
was returned to vertical position for each load increment
supported.
RESULTS
Designed for load, 30 pounds per square foot.
Required static-test proof load, 15 pounds per square
foot.
Drop of ailerons on
No. CfJ Pull on
left side of wing
on aile- the con- wing cellule ·
rons trol stick
Upper Lower
---- - - - - - - --
Pou'lllU Inche& Inch ea
5 15 2'4 3
30
Due to excessive cable stretch, it was necessary to
stop the test and go over the control system.
A second test gave the following results:
Drop of ailerons on
No. [ j J Pull on the left side of
on aile- the con- wing cellulc
rons t rol stick
Upper Lower
--- --- ------
Pounds Inches Inches
5 29 1)4 1)1
10 58 2)1 2)1
12~ 77 3Ys 3%
15 97 4 4Ys
The surfaces and their controls held the proof load
with excessive deflection.
DISCUSSION
A Ys-inch flexible steel cable is strong enough, but has,
a generous amount of stretch. Whether stressing the
cables to their safe working loads prior to their installa­tion
in the airplane will lessen the amount of stretch
obtained should be investigated. If not a larger size
cable should be used, or the use of a separate control
system for the ailerons on each side of the wing cellule,
thus preventing the cable stretch on one side showing up
at the other side. ·
The aileron tip deflections recorded are the combina­tion
of aileron torsional deflections at point of measure­ment
and cable stretch, the ·former amounting to
approximately 25 per cent of the reading.
CONCLUSION
Structurally the ailerons and their control system
supported the proof loading without failure.
RECOMMENDATIONS
Eliminate at least 50 per cent of the cable stretch .