Static test of the Boeing (PW-9) Pursuit airplane (Airplane Section report)

              File D 52.1 /P. W. 9/1
Vol. V
(AVIATION )
PUBLISHED BY THE CHIEF OF AIR SERVICE, WASHINGTON, D. C.
May 15, 1924 No. 466
ST A TIC TEST OF THE BOEING (PW-9)
PURSUIT AIRPLANE
. (AIRPLANE SECTION REPORT )
Prepared by D. B. Weaver
Engineering Division, Air Service
McCook Field, J?ayton, Ohio
February 18, 1924
WASHINGTON
GOVERNMENT PRINTING OFFICE
1924
Ralph Brown On:utllon
LIBRARY
MAY 16 Z013 ·
Non·Depoitory ·
Auburn University
C E RTIFICATE: By direction of the Secretary of War the matter contained
herein is published as a dministrative informa tion and is required for the proper
transaction of the public. business.
(n)
I l
SUIT AIRPLANE
SUMMARY OF RESULTS
Airplane: Boeing Pursuit (PW- 9) .
Type: I.
Description: The PW- 9 is a biplane, wings taper in
both dimensions. Gottingen 436 airfoil. Wings and
stabilizer were wood structures. Fuselage, chassis,
elevator, rudder and fin were welded steel structures.
Actual speed at ground 160.9 miles per hour. Built
by the Boeing Airplane Co. , Seattle, Wash.
Total weight : 2910 pounds.
Wing cellule weight: 408.5 pounds.
Wing area: 241. 7 square feet .
Engine: Curtiss D- 12.
Date
Nov. 10, 1923
Do _____ _
Do ___ __ _
Nov. 12, 1923
Do ____ _
Do __ __ _
Nov. 16, 1923
. · Do __ __ _
Nov. 22, 1923
Nov. 26, 1923
Nov. 19, 1923
Nov. 15, 1923
Part t ested Load
required
RESULTS OF TEST
Pounds per I
s~afa~t~ot Failed at
supported
Weight
Horizontal stabi- 35 pounds 50 · pounds .. -- - - - ------- 1.25 pounds
lizer. per sq. ft. per sq. ft . per sq. ft.
Elevator_ ______________ do ___ ____ _____ do _______ ______________ 1.85 pounds
Elevator controL ______ do ____ ________ do __ ___ __ -------------- per sq. ft.
Vertical fin __ ______ 30 pounds 45 pounds SO pounds 1.04 pounds
per sq. ft. per sq. ft. per sq. rt. per sq. rt.
Rudder_ _______ _____ ___ do _______ 55 pounds 57.5 pounds 1.31 pounds
per sq. ft . per sq. ft. . per sq. ft .
Rudder control__ _______ do _______ ----------- --------------- - - ________ __ _
Ailerons ______ __ ___ 35 pounds 35 pounds ------------ - - 1.3 pounds
per sq. ft. per sq. ft. per sq. ft.
Aileron control__ ____ ___ do ___ ____ 30 pounds ___________________________ _
WING .CELLULE
per sq. ft.
}High incidence ____ {8 ~e~d; 12 }10.5 _________ 11.0 ____ ___ _ {L~ef~~-ngs
Low incidence. ___ 5.5 old; 6.5 5. 5 __________ 5. 5---------- .--------------
new.
Reverse load ______ 3. 5 old; 4.0 4. o __ ________ -------- ------ _____ ________ _
new.
Dec. 27, 1923 Six-foot length of Factor 14 _ _______ _________ U. wing 22.4
leading edge.
None.
Do.
Do.
F ailure
Permanently distorted at SO pounds per square foot,
but did not collapse.
Rudder binges slightly bent.
None.
Do.
Control stick support distorted under load .
Front member of right N strut buckled as the load
was applied.
The fuselage cross tube between longerons conneet­ing
the flying wires pulled in two.
None.
Both satisfactory.
Nov. 28, 1923 Fuselage.- ---- - --- 7. O __________ 8. O __________ L. wing 13.9 190 pounds. Several members were badly bent but the fuselage
did not collapse.
Nov. 30, 1923
Dec. 5, 1923
Tail skid __________ 30 inches ____ o ____________ 6 incbes _____
1
9~ inches __ __ Fuselage badly bent; further test useless because of
design not suited to static test conditions.
ChasSstirsu: ts _____ ___ 24-inch drop
full load:
48-inch
39 inches ____ 41-inch drop_ 107 pounds. Right rear strut buckled.
droptload ..
Axle __ ____ __ __ ------ - --------- -- ---- ----- -- --- ------- --- --- --- -- -- --- -
Shock absorber · · I
I
Discussion: This airplane was designed under the I WITNESSES
old load factor requirements, but the recommendations . .
are based on the new load factor requirements. The Lieut. E. W. D1chman ___ Stabilizer and elevator; rud-tail
skid functioned in actual landings, etc., however, der and fin tests.
due to the design. The shock absorber cord was unable Lieut. C. W. Pyle _______ High incidence wing test.
to absorb the shock when given an impact test. Capt. T . E. Tillinghast __ Fuselage test.
OBJECT
This static test was conducted for the purpose of
determining the structural st.rength of the Boeing
PW-9 submitted in accordance with Contract No.
633, also known as the first article of Contract No.
713, dated September 29, 1923.
DATE AND PLACE
All tests were condu<;ted at McCook Field, Dayton,
Ohio, on the dates listed in the preceding table.
Lieut. Geo. C. Kenney __ Low incidence wing test.
Maj. Leslie MacDilL ___ Rudder and fin, and fuse-lage
tests.
C. L. Egtvedt_ _________ Reverse load, t ail skid and
chassis tests.
A. S. Niles_ ___ _________ All wing cellule tests.
W. E. Savage ____ __ ~--- All tests, except tail skid
test.
E. R. Weaver ____ _ _ _ _ _ All tests.
D . B. Weaver_ ___ _____ _ All tests.
(1)
None.
None.
GENERAL RECOMMENDATIONS
STABILIZER AND ELEVATOR
RUDDER AND FIN
WING CELLULE TESTS
2
Reverse load-none.
Low and high incidence conditions. Redesign the
p ortion of the fuselage where the flying wire fittings
are located. Change the position of the pins in the
strut to spar fittings.
AILERONS
Redesign the cable pulley brackets and the control
stick support.
LEADING EDGE
None.
FUSELAGE
None (except change called for in wing tests).
TAIL SKID
R.edesign the assembly.
CHASSIS
Increase the strength of the struts.
GENERAL DESCRIPTION
The PW- 9 is a biplane propelled by a Curtiss D-12
engine, low compression, rated horsepower 375 at 2,000
revolutions per minute. It was designed and built by
the Boeing Airplane Co., Seattle, Wash.
The wing cellule and stabilizer are of wood. construc­tion,
excepting the welded steel interplane N and
cabane struts. The fuselage, chassis, elevator, rudder,
and fin are of the welded steel tube construction.
All surfaces are covered . with cotton fabric.
The control surfaces are actuated by means of flexi­ble
steel cables.
For assembly drawings, see Figures 1 and 2.
The high. speed at the ground is given by the Flight
Test Section as 160.9 miles per hour.
Gross weight __ __ ___________ _____ 2910 pounds.
Disposable load ___________ ·____ __ 920 square feet.
Total wing aTea_ ~ _______ ____ ____ 241 :7 square feet .
Wing loading in pounds per square
foot ___ ________ ________ --· - - - 12.7.
Weight per horsepower_ ___ __ __ ___ 7.8 pounds.
Airfoil __ ___ _____________ c _ _ _ _ _ _ Gottingen 436.
The equipment is according to the specifications in
the contract.
WING CELLULE
DESCRIPTION
The upper wing is built in one continuous piece,
around two box type spars, whose depth tapers both
inboard and outboard from the point where the cabane
struts are attached. The upper and lower wings taper
both in chord and thickness dimensions.
Figure 5 is a drawing of the interplane strut.
Figure 6 is a drawing of the cabane struts.
A pair of %-inch diameter streamline flying wires
are employed, located in the plane of the front spars.
In this same truss one -&-inch diameter streamline
landing wire is positioned between the pair of flying
wires. The truss of the rear spars does not contain
any wire bracing.
Figure 7 gives the typical wing ribs and compression
ribs.
The lower compression ribs between the spars are
made of a solid piece of spruce with many large lighten­ing
holes. The spars are box type, having a spruce
flange top and bottom, connected by a three-ply veneer
web on both sides.
Figures 8 and 9 show the cross section of the spars.
Figure 10 is a sketch showing the strut to spar
fitting , and how the strut is attached to the ~par.
Weight of upper wing ___ __________ 232 pounds.
Weight of the two lower wing panels_ 111 pounds.
Area of the two lower wing panels_ 81.12 square feet.
Area of upper wing_ ____ ___ _______ 160.6 square feet.
Structural weight of upper wing
per square foot _ _ _ _ _ _ _ _ _ _ _ _ 1.45 pounds.
Structural weight of lower. wing
per square foot _____ ___ _______ _ 1.37 pounds.
Weight of wires, housing, inter-plane
and ca bane struts___ ____ __ _ 40Yz pounds.
Weight of complete wing cellule ___ 415 pounds.
Structural weight of wing cellule
per square foot __ · __ __ ____ ____ ___ 1.72 pounds.
The lower wing has a positive stagger of 11 inches.
PROCEDURE FOR TEST (REVERSE FLIGHT CONDITION)
The airplane was assembled as for flight. It was sup­ported
through the fuselage so that the mean chord
made an angle of 14 degrees with the horizontal, trail­ing
edge down.
The center of gravity of the load was located 25 per
cent of the chord length back of the leading edge of
the wing.
The load was applied according to the loading sched­ule
in Figure 11. After a load had been put on and
supported for five minutes, deflection readings at
points indicated along the $.par (fig. 12) were taken.
The wing tip retreat was measured as indicated in the
same figure.
RESULTS
The required load factor of 4 was supported satis­factorily
without any actual or indicated failures.
Tabulated results are given on Figure 12 and lower
wing spar deflection curves on Figure 13.
DISCUSSION
None.
CONCLUSION
The wing cellule answered the requirements satis­The
lower wing is made in two parts, a right and a factorily.
left panel. Each panel is connected to the fuselage
by two Yz-inch diameter bolts and to the upper wing
PROCEDURE .FOR TEST (LOW INCIDENCE CONDITION)
by means of an N-type interplane strut. See Figures , The airplane was set up in an inverted position so
3 and 4 for general layout of the upper and lower I that the angle ·v between the wing chord and the
wings. horizontal was 7° 15', trailing edge down.
1
1
The angle of incidence a of the wing at low incidence,
and f3 the angle between the vertical and resultant air
force, are determined from wind tunnel data on the
Gottingen 436 airfoil.
L.a = -2° 30'
L.f3 =cot- I L/D =cot-I 12 =4° 45'
L-Y = L.f3- L.a =4° 45'-( -2° 30') =7° 15'
The center of gravity of the load was placed at 60
per cent of the chord from the leading edge of the
wing.
The load was applied according to the loading
schedule (fig . 14). At the points indicated (fig. 15)
deflection and retreat readings were taken after .each
increment had been supported for five minutes .
RESULTS
3
PROCEDURE FOR TEST (HIGH INCIDENCE )
The airplane was reset so that the angle between
the wing chord and horizontal represented by -y was
7°, leading edge down.
L.a = 12° 24'
L.f3 =cot-I L/D =cot- I 11.0 =5° 12'
L.-y= L.a- L.f3=12° 24' -5° 12' =7° 12'
The center of gravity of the load was located at 30
per cent of the wing chord.
The loads were placed a ccording to the loading
schedule, Figure 17. Deflection and retreat readings
were taken at points indicated in Figure 18 after each
load increment had been supported for five minutes.
RESULTS
The required factor (old specification) of 5.5 was Required load factor (old specifications) was 8.5;
put on and after having supported the load for three new specifications 12. At a factor of 7 the N struts
minutes, the cross tubes between the lower longerons were bent very noticeably . The bending increased
connecting the flying wires, pulled in two. See Figure with each added load increment. Af a factor of 8 the
32 for _photograph of failure; Figure 15 for tabulated test was delayed by the other side df the cross tubes
deflect10n and retreat readings, and for wing spar . between ther lower longerons pulling out, a similar
deflection curves, Figure 15. I failure to the one durirrg .low incidence. test.
DISCUSSION As the jacks were let down at a load factor of 11 ,
Although the rear spars showed a generous amount
of deflection, further test for low incidence condition
was deemed unnecessary .
Due to the t ype of fitting between the interplane
struts and the spar, the spar deflection caused art "S"
shape bend in the interplane struts to an amount of
one-half inch from the true centerline.
CONCLUSION
With the fu selage made strong enough and the strut
to spar fitting pins relocated, so that they will be
parallel to the wing ribs, the wing cellule will be
satisfactory for the old requirements.
RECOMMENDATIONS
the front member of the right end strut failed , as
shown in Figure 33. For tabulated deflection and
retreat reading, see Figures 18 and 19. For wjng sj:lar
deftectfoiI. curves-, see· Figures 20 and 21.
DISCUSSION
Th:e distortion and additional load imposed on the
N strut by the spar deflection was an important fact or
in causing failure .
At a load factor of 8 the other. side of the fuselage
cross tubes pulled out. These tubes were rewelded
and in addition a steel strap was connected from one
side of the fuselage to the other side, carrying the
stresses around the welds,' instead of through them.
This held satisfactorily.
Redesign the portion · of the fuselage where the The following test results were obtained by the
flying wire fittings are located. I material section from four wing spars, the right upper
Change the position of the pins in the strut to spar front and rear, and the left lower front and rear.
fittings. Test specimens were taken from the spar flanges .
Strength properties of flang es from Boeing pursuit wing beams
Location of- Modulus of- Fiber Compres- Work to
Moisture stress sion maximum
content, Specific elastic parallel load,
per cent gravity Rupture, El asticity, limit, to grain, Beam Flange lb. per l ,OOOlb . lb. per lb. per lb. per sq. in. per sq. in. sq. in. sq . in. sq . in.
Upper right:
7. 8
FronL - - ---- -- - - - -
{Top __ __ __ 8. 63 0. 395 11, 455 l , 790 8, 535 8, 190
Bottom __ _ 9. 35 . 506 15, 865 2, 390 10,830 10, 140 14. 3
Rear ___ ___ _____ ___ {Top ______ 8.80 . 431 13, 250 2, 060 9, 425 8, 815 11. 6
Bottom ___ 9. 16 . 443 13, 490 2, 159 10, 150 8, 730 8. 8
Lowe.r left:
. Front . ___ _____ ____ {Top _ _____ 9. 54 . 438 13, 210 2, 140 9, 855 8, 915 10. 3
Bottom ___ 9. 61 . 402 11, 500 1, 762 8, 625 7, 755 i . 5
Rear ___ ______ _____ {Top __ _____ 9. 68 . 4.29 12, 050 1, 938 10, 200 8, 630 10. 7
Bottom ___ 9. 84 . 402 10, 090 1, 832 10, 090 7, 960 2. 9
DISCUSSION OF RESULTS
The spruce in the flanges gave gooQ. and consistent
strength properties. With the exception of work to
maximum load for the bottom flange of the lower
left rear beam, the strength properties of the flanges
were slightly h!gher than those for spruce of average
grade. The small size of the test specimens may have
had a tendency to increase the strength properties
slightly.
4
An examination of the small test specimens disclosed
a few small compression failures in the bottom flange of
the lower left rear beam, which may have resulted from I
the static test. In cross bending, this specimen failed in
brash tension, giving the extremely low value in work to
maximum load.
Furthermore, a slope in grain of approximately 1: 12
was found in a portion of the bottom flange of the lower
front left beam. This defect should have been noted
before construction of the flange.
CONCLUSION
The strength properties of the spruce flanges were
good and consistent.
The moisture content was satisfactory and uniform.
A few small compression failures· were noted in the
bottom flange of the lower left rear beam and the slope
of the .grain of 1: 12 in the lower front left beam.
Excepting the cross tubes ·of the fuselage in the plane
of the flying wires, the wing cellule held sati sfactorily
the load for which it was designed. For the new re­quirement
some redesigning is necessary before the
wing cellule will be entirely satisfactory.
RECOMMENDATIONS
Redesign the portion of the fuselage supporting the
flying wire fittings. Redesign the end strut spar
fittings.
AILERON TEST
DES.CRIPTION
A spring balance attached to the control stick was
used to determine the pounds pull on the stick for each
load increment added. Load . increments of 5 pounds
per square foot were put on, up to and including 25
pounds per square foot, after which 27"2 pounds incre­men.
ts were added for the remainder of the test.
RESUf, TS
Required load- 35. pounds per square foot.
T railing I
Load in Pounds edge of
pounds pull on ailerons per control below Remarks
sq. rt. stick neutral
position
- -- - - -
5 10 0
LO 20 J4
. 15 30 f.
20 40 ft
25 50 l Aluminum control stick support
27J,2 60 1)4 started to deform, also cahle pulley
30 70 IV, bracket located in t.he wing began
32~·1 7.5 1Ys to bend.
35 85 2Ys
See Figure 34 showing how the pulley brackets bent
under load. Part of the bracket from the other half
of the wing is also shown in this figure.
Figure 35 shows how the control stick was bent.
DISCUSSION
In the pulley bracket, the location of the bolt holes
does not permit the bracket to keep its original form
and support the required load. The control stick sup­port
bent, due to being previously distorted by the
·end of the control stick ,_ which bent the four angular
legs, weakening them for the aileron load, \Vhich, when
put on; cau.sed the bracket to distoTt and shift the
stick from its original position.
CONCLUSION
The ailerons were satisfactory, but their control
system should be strengthened.
RECOMMENDATIONS
The two ailerons were located in the trailing edge _Redesign the cable pulley brackets and the control
of the upper wing, each supported by four hinges. I stick support.
They were controlled by means of flexible steel cables, LEADING EDGE STATIC TEST
placed inside the wing on the back side of the rear spar.
From the upper wing the cable extended to a pulley DESCRIPTION
on the lower longeron and then to the control stick. A 5-foot section of the upper wing leading edge
The aileron was built about a box-type spar. The was selected for the test. The leading edge extended
rib webs were three-ply veneer, poplar core with 1 forward of the vertical centerline of the front spar, an
mahogany sides. The web was capped with spruce average distance equal t o 16.23 per cent of the averc
strips, and terminated at the trailing edge, which was age chord for the section selected.
a ti-inch flexible steel cable. For an assembly draw- A 6-foot section was selected from the lower wing.
ing of the aileron, see Figure 22. The leading edge forward of the vertical centerline of
Area of each aileron ____ ___ ____ __ __ 7. 77 square feet.
Weight of each aileron __ _____ _____ _ l0.00 pounds.
Structural weight per square foot of
area ~ ___ ___ ______ _____ ____ __ __ . 1.3 pounds.
PROCEDURE FOR TEST
The aileron was assembled on the wing as for flight.
With the control system in the neutral position, initial
.readings were taken.
-the front spar amounted to 15.09 per cent of the aver­age
wing chord of the section.
Figure 23 gives the dimensions of the sections tested,
locating them in respect to their original position when
a part of the wing. ·
PROCEDURE FOR TEST
Each section was supported in an inverted position
with points of support at the spars between the rib
c11pstrips. ·
..
5
The load on the leading edge was counterbalanced
by shot bags placed on the section along the rear spar.
The factor of failure was computed as follows: For a
leading edge having a length equal to 10 percent of
the wing chord, one-half of the normal· load per run­ning
foot on the entire wing, from which sectiqn was
taken, was considered a load factor of one.
The load factor for the upper wing leading edge was
-.1:-6I2Q3 X the 1o a d f actor i. ust found, for the lower wm. g
leading edge it is ·~~~9 x the load factor.
Dividing the new factor into the load per foot run
supported by the leading edge, gives the load factor at,
which the leading edge failed.
RESULTS
A factor of .14 required.
01623 x 2495
.10 2 . 1· 9·.353 6 34.8 basic factor, upper wing.
3900
- 5 - = 22-4 load factor at which the leading edge of
34.8 upper wing failed.
·1509 x 2495 · ·664 40.8 basic factor, lower wing.
.10 2 . 30.58
3365
- 6- = 13.9 load factor at which failure occurred.
40.8
DISCUSSION
None.
CONCLUSION
Both leading edges were satisfactory.
RECOMMENDATIONS
None.
STABILIZER AND ELEVATOR TESTS
DESCRIPTION
The stabilizer was built around two spruce spars, the
front spar having an "I" cross section, was glued up of
six thin spruce pieces to a similar curvature of the
leading edge. The rear spar was· made .from one piece
of spruce, routed on the stabilizer side, which provided
clearance for the elevator spar. The leading edge was
co.vered with -fi-inchi three-ply veneer, mahogany side
with poplar core. The stabilizer was adjustable,
through the front spar, being hinged about the tubular
rear spar of the stabilizer was connected to the fuselage.
The elevator spar being positioned by means of a
bearing on the tail post of the fuselage.
The elevator was of the steel tube type, with elec­trically
welded joints, and fabric covered. Control
was obtained by two pairs of flexible s'teel cables from
the control stick to the control masts located on both
sides of the fuselage.
See Figure 24 for an assembly drawing, and Figure 37
for a photograph.
Weight of stabilizer_ ______________ 23.5 pounds.
Weight of elevator __ ,_ __________ _ 18.5 pounds.
Area of stabilizer_ _-____ ___________ 18.75 square feet.
Area of elevator_ ___ ______________ 9.9 square feet.
PROCEDURE FOR TEST
The surfaces and control system were completely
·assembled as for flight. 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. Scales were suspended at
points indicated in Figure 26, from which deflection
readings were taken by means of a wye level.
The load was applied in increments of 5 pounds per
square foot up to and including 25 pounds per square
foot, and then in 272 pounds increments until failure
' resulted.
The average load per square foot on the elevator
was assumed to be two-thirds the .average load on the
stabilizer. The stabilizer load was assumed to be
uniform, and the elevator load as varying from a
maximum at the hinge to one-third maximum at the
trailing edge, which results in a center of pressure
location for unbalanced surfaces at +, of the mean
chord.
RESULTS
Tabulated results are given in Figure 26. The
stabilizer adjustment worked very well until a load of
25 pounds per square foot when the adjustment of the
surfaces required all one hand could pull. At 30
pounds per square foot the adjustment mechanism
could no longer be operated. After a load of 60
pounds per square foot was supported without failure,
further testing was decided unnecessary for fear of
damaging the fuselage.
DISCUS,SIO -
None.
CONCLUSION
The tail surfaces and their control system was
elevator spar. There were two streamline wires per
side, connecting the li.inge, located 2972 inches from
the vertical centerline of the airplane, with the lower satisfactory.
1 ongeron and the fin mast.
For a general layout of the stabilizer, see Figure 24.
For a photograph of the stabilizer structure, see
Figure 36.
The ribs were made· from A -inch mahogany sides,
None.
RECOMMENDATIONS
. RUDDER AND FIN TESTS
DESCRIPTION
poplar core veneer, reinforced and capped with rec- The structure of the rudder and fin was welded steel
tangular section spruce strips. tube construction, ·fabric covered. Three hinges con -
For typical rib drawing, see Figure 25. nected the rudder to the fuselage and fin. One of the
T~e stabilizer and. elevator were connected ?Y means [ three hinges was located on the tail post of the fuselage,
of six band type hmges. Through these hmges the the other two were on the fin m·ast. The fin mast
telescoped into the fuselage tail post for about an inch,
being held there by the coμipression from the two wires
connecting the fin mast to the elevator stabilizer hinges.
See Figure 27 for an assembly drawing of the rudder
and fin .
The rudder was controlled by means of a -!.-inch
diameter flexible steel cable.
For a photograph of the vertical tail surfaces with
the fabric removed, see Figure 38.
Rudder Fin
WeighL . .... .. ..... . . ..... .. .... .... 6 lbs .... .. .... I\ lbs.
Area .. ... .... ... .... ... ... ... ..... . .. 5.78 sq. ft ...... 4.58 sq . ft.
Structural weight per square foot . .... 1.31 lbs. .... .. . 1.04 lbs.
PROCEDURE FOR TEST
The rudder and fin were assembled to the fuselage as
for flight. To one end of the rudder bar the control
cable was connected, to the other end of the rudder bar
a spring balance with block and tackle was connected
by means of which the force in p ounds necessary to
actuate the rudder under load was obtained. The
fuselage was firmly supported on its side with the
rudder and fin in a horizontal position. At the points
indicated in Figure 28, scales were suspended in order
that the deflection readings could be read by means of
a wye level. The loading and taking of readings was
the same as for the test of the stabilizer and elevators.
RESULTS
Required load- 30 pounds per square foot.
Figure 28 gives the deflection readings and pounds
pull on the rudder bar per load increment.
A load of 50 pounds per square foot caused a perma­nent
set in ·the fin ribs. When 57.5 pounds per square
foot were supported, the load was about to fall off.
No more load was added.
DISCUSSION
Figure 38 does not show how much the ribs were
distorted, however; the ribs could be straightened to
their original shape.
The addition of a small bolt through the lower end
6
PROCEDURE FOR TEST
The fuselage was supported in a t est jig by the upper
wing through the cabane struts and flying wires
until they had supported their required portion of
the load, after which the lower wing bearing points
were allowed to support the remainder of the load
applied.
The load was applied at four points, A to D, inclusive;
Figure 30, according to the loading schedule. Scales
were suspended from four points, E, F , G, and H,
from which deflection readings were taken by means
of a wye level. The load increments were supporte'.i
for five minutes before readings were taken.
RESULTS
Factor of 7 required.
As . the load was applied at a factor cf 5 the rear
spar cabane strut fitting, left side, pulled away from
the spar, see Figure 39. At 6.5 the engine bearer
strut, right side, began to buckle. At a factor of 8
the lower longerons, left side, buckled in t he third
bay forw ard of the tail post, see Figure 40. In the
third and fourth bays from the tail post the vertical
tubes were buckled.
For tabulated deflection readings, see Figure 30.
DISCUSSION
When the fabric was removed from the wing a rivet
was found sheared off, due to shifting of the flying
wire and strut to spar fitting, see Figure 41. This
failure is credited to the fuselage test.
At a factor of 8 no load was put on the engine
bearers.
When the cross tubes supporting the flying wire
fittings have been made strong enough, the fuselage
structure will be satisfactory.
CONCLUSION
The fuselage held the required factor satisfactorily.
RECOMMENDATIONS
None (except change called for in wing test).
DYNAMIC 'I;EST OF TAIL SKID
of the fin mast and the fuselage tail post, at the place DESCRIPTION
where the fo~mer t e'.escopes with the latter, would be Photograph, Figure 42, shows the general design of
very helpful m holdmg the fin to the fuselage in case I the tail skid and shock absorbing unit. Foot control
one of the diagonal wires should break. of the tail skid was provided by means of flexible steel
CONCLUSION
The surfaces held the requir~<l'\:iad satisfactorih.- ' and the rudder control worked very well.
RECOMMENDATIONS
None.
FUSELAGE TEST
DESCRIPTION
The fuselage was made from mild steel tubes,
electrically welded together. For a general layout
and tube sizes, see Figure 29.
Weight of fuselage- 190 pounds.
cables connecting the rudder bar to the tail skid
control mast. In each cable a steel spring with a fixed
amount of travel was mounted , so that the shock and
jerking of the tail skid when taxying will not be trans­ferred
so severely to the rudder bar.
Weight of tail skid assembly- 9_Y2 pounds.
PROCEDURE FOR TEST
The chassis was intact with the fuselage and tail
skid. The wheels of the chassis were placed on a plat­form
inclined to the horizontal at an angle of 9_Y2
degrees in the plane of the axle.
The load was placed on the tail post and at the
· vertical strut location along the fuselage, as shown in
,_
Figure 42. The load was made equal to the tail skid I
reaction when the airplane was in a landing position,
which amounted to 450 pounds.
The first drop was made from a 6-inch height and
each successive drop from a 3 inch greater height than
the last drop.
RESULTS.
Required height of drop-30 inches.
The first six-inch drop distorted the rear part of the
fuselage. This concluded the test. Figure .42 is a
photograph of the failure.
DISCUSSION
The design of the tail skid and the shock absorbing
unit was such that when the impact occurred in a
landing position the shock absorber unit was ineffective,
for it would not function. The only force that would
cause the rubber to function , was the horizontal drag
force present when the airplane was in motion and the
tail skid sliding along on the ground. The resultant
force from the drag and impact forces, which would
actuate the tail skid and extend the shock absorber
unit, was not present in a plain drop as used for .the
standard impact test and which frequently occurs in
tiervice.
CONCLUSION
The tail skid was not satisfactory.
RECOMMENDATIONS
Redesign the tail skid assembly.
LANDING CHASSIS
DESCRIPTION
The landing chassis was all steel with the exception of
the axle fairing, which was wood framework with three­ply
veneer covering. This fairing has a thick airfoil
cross section, 18 inch chord, which gives 6.75 square
feet of lifting surface as well as being the crosstie
between the two struts.
97159- 24t- - 2
The struts are formed from ,sheet steel and acetylene
welded at the trailing edge.
The shock absorber was wound with .Yz-inch diameter
cord.
Weight of chassis, including wheels-107 pounds.
For an assembly drawing of chassis, see Figure 31.
PROCEDURE FOR TEST
With the cha·ssis assembled to the fu selage and the
tires inflated to an air pressure of 60 pounds per square
inch, the tail end of the fu selage was connected
through a hinge coupling to the test jig, so that the
center of gravity of the airplane was vertically above
the point of tangency of the wheels.
To simulate the combination of direct load and side
thrust, an auxiliary platform was placed under the
wheels, built at an angle of 9.Yz degrees to the hor­izontal.
When the impact occurred the ratio of side
load to down load was 1 to 6.
The reaction at the wheels with the fuselage in flying
position was 2,650 potmds. In order to insure safety to
everything concerned while testing, one-half t he load
was used dropped from twice the height. Height of
first drop 6 inches.
RESULTS
Required height of drop 24 inches with full load, or
48 inches with half load, which for all practical purposes
are the same when the height of drop is small.
The chassis held for a 39 inch drop; when dropped
from a 41 inch height, the · right rear strut began to
buckle, as shown in Figure 43.
The shock absorber units functioned satisfactorily.
DISCUSSION
None.
CONCLUSION
The chassis was not strong enough.
RECOMMENDATIONS
Increase the strength of the chassis struts.
8
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