Static-test and stress-distribution studies of the Matériel Division 55-foot cantilever all-metal wing (Airplane Branch report)

              .
File D 00.12/122/No. 3501 AIR CORPS TECHNICAL REPORT No. 3501
Vol. VII
CIRCULAR
PUBLISHED BY THE
STATIC-TEST AND STRESS-DISTRIBUTION STUDIES
OF THE MATERIEL DIVISION 55-FOOT
CANTILEVER ALL-METAL WING
(AIRPLANE BRANCH REPORT)
UNITED STATES
GOVERNMENT PRINTING OFFICE
WASHINGTON: 1932
No. 663
Ralph Brown Or ughOn
UBRAR
Non-Depoi or~
Auburn Univ rs1ty
TABLE Page
I. 1
II. 1
III. Summary_____________________________________________________________________________ 1
IV. Description __________ --------__________________________________________________________ 2
V. Computation of test loading ___________________________ .___________ ___ ____________________ 5
VT. Inverted-flight t est_ ____________________________________________________________________ 7
VII. Low-incidence test_ ____ ___________________ -------------------- __ ------------___________ 10
VJ II. High-incidence test ___ __________________________________________________________________ 12
IX. Lead ing-edge test_ _____________________________________________________________________ 16
X. Torsional characteristics_ _______________________________________________________________ 17
X T. Stress-distribution stud ics_ _ _ _ _ _ _ _ _ 20
Description of apparatus ___ ______________ ______________________________________ ~____ 20
Method of loading ____ ___ ________________________ __ _________________________________ 21
Elastic-axis location ________________________________________________________________ 21
Beam-load tests _______ ________________________________________________ ·_____________ 21
Chord-load tests ___ _____________ __ __ ___ ________ ____ ___ ___ _____ __ __________ ___ ______ 25
Torque-load tests __________________________________________________________________ 28
Stres d istribution in sand-load tests __________________________________________________ 30
Computation of . tresses at station!_ _______ __ ____________________ ___ ________ __ _______ 33
XII. Tests to determine the contribution of the leading and trailing edges to the wing rigidity__ _____ ____ 34
XIII. Weight distribution ____________________________________________________________________ _ 41
XIV. Suggestions for improvement_ _________ ------ ---- --- _____ -------_________ ____________ ___ _ 43
LIST OF ILLUSTRATIONS
F ig. No.
1. Upper flange- --------------------------- ------------------------------------------------- 3
2. Lower flange_ ___________________________ _____ ___ ___ ____ _________________________ _____ ____ 3
3. Covering schedul e__ ________ __________ ______ ___________ ____ ____________ ____ _______________ 4
4. Determination of equivalent wing _____________ _________ _____ ___ ____ _________________ _____ __ 4
5. Load distribution_ _________ __ _____________________________________________________________ 6
6. Bending moments due to loads applied at station 17__________________________________________ 8
7. Beam bending-moment curves _________________________ _______________ _________________ _____ 8
8. Loading schedule, inverted flight ___ __________________________________________________ ------ 9
9. Inverted-flight test deflections__ ____________________________________________________________ 10
10. Loading schedule, low incidence_ ___________________________________________________________ 11
11 . Low-incidence deflections_ _________ ______ __________ ________________________________________ 12
12. Distortion of front upper flange angle, high incidence______ ___________________________________ 13
13. Loading sched ule, high incidence_ ____________ __ _____________ ________________________ _______ 15
14. Mean leading-edge deflections, high incidence__ ________________________________________ ______ 15
15. Mean trailing-edge deflections, high incidence___ _____________________________________________ 16
16. Mean tip deflections __ ____________________________________________________________ - _ - - - - - - 17
17. Torsional characteristics ________ _______________________________________________________ - - - - 18
18. Comparison of torsional rigidities, XHB wings __ ____ ______________________ __________ _____ ____ 18
19. Defiection-point locations, torsion test _________ _________________ • ____________ __ ___ __ - ___ - _ - -- - 20
20. Electrical circuit, McCollum-Peters telemeter apparatus_____ ____ ______________________ ______ __ 20
21. Strain-gage locations _____ _______ _______ ___ __ ________ __ ____ ___________ - - _ - - - - - - - - - - - - - - - -- - 21
22. Elastic axis__ ____________________________________________________________________________ 22
23. Section station 47'2 - ________ _____ __________ _______________ ___ _____________ - - - - - - - - - - - - - - - - - 23
24. Stress d istribution, beam-load test, upload _______ ___ __ _________ ________ ___ -- - - - - - - - - - - - - -- - - - 25
25. Stress distribution, beam-load test, download __ __________________________ _____________________ 25
26. Stress distribution, chord-load test, forward load __ ___________________________________________ 27
27. Stress distribution, chord-load test, back load_ ____________ ___ ________________________________ 27
28. Stress distribution, computed and measured stresses____ __ _____ ____ ____ ___________ ____________ 27
29. Chord-load tests, computed neutral-axis locations___ _____ _____________________________________ 27
30. Stresses at station 4% due to torque loads_____ _______ _____ ____ ____________ _____ _____________ 30
31. Stresses at station 14)4 due to torque loads------- --------------------""'- ---------------- --- --- 30
(ill )
JV ..
~~ ~
32. Stress distribution, inverted-flight test_ ____ ___ _______ _________ ____________ _____ ______ ________ 32
33. Stress distribution, low-incidence test____ ____________ __________________________ ___ ____ ____ ___ 32
34. Stress distribution, high-incidence t est___ __ __ ______ ________ __ ___ __ ____ __________ ______ __ _____ 32
35. Section, station L ___________ _______ _ . _ __________ o_ ! _ _ _ _ _ _ _ _ _ _ _ _ 33
36. Torsional properties after higb-incidence test_ _______ _______ ___ ________ ____ ______ _____________ 35
37. Torsional deflectio11s, leading and trailing edge test, 34,388 inch-pounds torque_ ______ ___ ______ __ _ 35
38. T orsional deflections, leading and trailing edge test, 68,775 inch-pounds torque _____ --- -- --------- 36
39. Torsional d eflection s, leading and trailing edge test, 85,970 in ch-pounds torque_ ____________ ______ 36
40. Beam deflection s, leading and t railing edge t est, 500-pound load ______________________________ __ 37
41. Beam deflections, leading and trailin15 edge test, 1,000-pound load ___ _____________ ______ . _ . _ _ _ _ _ 37
42. Suggc ted increase in rib lightening_ _____ ________________________ __________________ ___ __ ____ 41
42a. - ----- - - ---- -- - ---- - -- --- - ---- ------ -- --- - - - - - ---- - -- - - - -- - ---- - ---- - - --- ------ -- - - -- -- 43
43. Bulkhead a ssembly, station 15 _ _ _ _ ___ _ _ _ _ _ _ _ _ __ _ _ _ _ ___ _ _ _ ___ __ _ _ _ __ _ __ _ __ _ __ _ __ _ _ _ __ _ _ 44
44. Fitting attachment, bulkhead, station L _ _ _ _ _ _ 44
45. Details of fitting attachment, station L _ _ _ _ _ _ __ _ _ __ _ _ _ _ __ _ _ ___ __ _ __ _ _ __ _ _ __ __ _ _ _ __ _ _ __ ___ 45
46. Front web -- ----- ---- ---- - -- --- -- - ----- - - - - - - -- ------ - --- - - - ---- - -- - ---- - ------- - -- - ----- 45
47. R ear web ___ _______ ____________________ __________________________________________________ 46
PHOTOGRAPHS
48. Photo I o. 39684, wing prior to installation of lower flange ______ ___________ ____ __ ________ ______ 4 7
49. Photo o. 39681, corrugated upper flange __ ------ --- ------ -- ---------------- - --- - ------- - --- 47
50 . Photo Jo. 39685, typical sheet used on lower flange ___ ________________________________________ 48
51. Photo No. 39687, riveting flat covering to lower corrugated flange_ _____________________________ 48
52. Photo No. 39686, upper flange a t tip_____ ___ ______ __________________________ ____ _____ _______ 49
53. Photo No. 39688, internal wing structure between stations 11and12 ___________ _________________ 49
54. Photo No. 39682, method of splicing corrugated flange materiaL _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ 50
55. P hoto No. 39683, typical nose bulkhead__ __ ______ _______________________________ ____ ________ 50
56. Photo No. 39524, method of applying download at elastic axis in beam-load test_ __________ _______ 51
57. Photo No. 39735, method of applying loads in beam-deflection t est _ _____ ___________ _________ ___ 51
58. Photo Jo. 39736, method of applying dragfoads at wing tip __ _____ ___ ______ __ _________________ 52
59. Photo No. 39734, method of bracing wing agal.nst drag loads____ ___ __ _____ ____ __ ____ ___ _______ 52
60. Photo No. 39526, telemeter strain-gage installation on leading edge__ ___________ ____ __ __ __ ______ 53
61. Photo No. 39527, telemeter strain-gage installation at tip, upper chord ____ ___ _____ __________ ____ 53
62. Photo No. 39523, telemeter strain-gage installation, upper surface, station 4.l4-- - -- - ------------ -- 54
63. Photo No. 39529, telemeter stra in-gage installation on lower chord at root and compression buckles
in beam-loading test___ _____ ___ ______ ______ ______ ___ ____________ _____ __ ____ ___ __________ 54
64. Photo No. 39525, telemeter strain-gage installation on leading-edge torsion test___ _______ ___ ____ __ 55
65. Photo No. 39528, shear wrinkles on lower surface in t or sion t est_ __________________________ ___ __ 55
66. Photo No. 39889, upper surface of wing in chord-loa d t est--chord load, 1,250 pounds, acting back at
station 17---- -- --------- - --- - -- --- -- -------- - - - ------- - ------- - ------- - - --- - - - - - - - - - - - 56
67. Photo No. 39891, set-up for low-incidence t est, 5.5 factors in place __________________ ______ ______ 56
68. Photo No. 39743, 15.5 factors in place after wing failure in h igh-incidence test ___ -- - --------- - --- 57
69. Photo No. 39744, wing supporting 11 factors, high-incidence test ______ _________________________ 57
70. Photo o. 39745, front view at 15.5 factors in place during high-incidence test_ ______ _________ ___ 58
71. Photo No. 39746, failure of nose sean1 and lower flange, high-incidence test __ __ ______ __ __________ 58
72. Photo No. 39749, failure of lower front web, flange angle, and corrugated flange, high incidence____ 59
73. Photo No. 39750, failure at rear web, high incidence__ ____ ____ __________________________ ______ 59
74. Photo No. 39890, failure at station 1 dming high-incidence test __________________ _______________ 60
75. Photo No. 39748, secondary failure of lower skin joint, station 5, high-incidence test ________ ______ 60
76. Photo No. 39753, view of upper t ra iling-edge surface of wing after failure in high incidence, showing
inspection openings and relieving cradle in place _ __ ___________ ___ ___ _______________________ 61
77. Photo No. 39754, elongation found after test in angle irons from which wing was supported ______ __ 61
78. Photo No. 39886, leading-edge test, 5,375 pounds in place between bulkheads 3 to 67'2. inclu sive ____ 62
79. Photo No. 39885, failure of leading edge under load of 5,475 pounds between bulkheads 3 to 6f2__ __ 62
ST A TIC-TEST AND STRESS-DISTRIB
MATERIEL DIVISION SS-FOOT CA
WING
OF THE
L-METAL
(By C. G. Brown, Structures Unit ; Capt. Carl F . Greene, Air Corps, Acting Chief, Airplane Branch; By direc­tion
of Maj. C. W. Howard, Air Corps, Chief, Experimental Engineering Section, Materiel Division, Air
Corps, Wright Field, Dayton, Ohio, June 12, 1931)
I. OBJECT
The object of the tests on the Materiel Division 55
foot cantilever all-metal \Ying was:
1. To determine its ultimate strength in high incidence
for comparison ,1·ith the strength indicated by
the stress analys is.
2. The determination of the torsional characteristics
of the complete wing,
3. The determination of contribution of the leading
and trailing edges to the beam and the torsional
rigidity of the wing.
4. The determination of the distribution of stress in
the fla nges of the wing due to chord, beam, and
t orque loads.
II. I TRODUCTION
The Materiel Division 55-foot wing is the fourth of
a series of 55-foot wings to be tested by the division .
The other th ree were the XHB- 3 two, three, and multi
spar wings. These three wings were %-scale models
of a large monoplane bomber wing. The latter were
of welded steel con truction and would have been
fa bric covered .
T he l\faleriel Division 55-fout ll'ing project 1rnK
ini tiated dming 1928 by Capt. Ca rl F. Greene, with
\\'horn origi11atcd the idea of a ion;ionally rig id canti­lever
wing, embodying a central box beam using cor­rugated
alumillum-alloy sheets for flange members.
An extensive program of research in the field of
box beams for wing structures carried out at the
division by Captain Greene and Dr. J. E. Younger
furnished data for the design and guaranteed the
economy and desirability of continuing the research
program with the con truction of a full-sized "'ing.
The wing decided upon was of the same plan form and
span as the XHB- 3 series, in order that the dural­umin
box-beam wing could be directly comparable
with the XHB- 3 series. The project occupied the
period beginning in 1928 and continuing to the elate of
completion of the final report on the tests of the "'ing.
During the period of the preliminary de ign, occupying
the period from September, 1928, to March, 1929,
various individuals contributed valuable suggestions
and crit icisms. T he original stress computations,
drafting, and preliminary tests were carried out largely
by Mr. R.H. Rice, assisted by Mr. E. G. Bruce. The
checking of the final design drawings was carried out
by Captain Greene and Mr. C. G. Brown.
Static tests on the wing were carried out during
October and November, 1930, under the supervision
of Mr. E. R. Weaver, assi ted by Mr. C. G. Brown.
The preparation of the final report was the work of
Mr. Brown, '"ho was assi tecl by Mr. E. H. Schwartz.
III. SUMMARY
The wing supported in the high-incidence condition
a, gross load of 120 pounds per square foot and failed
under a load of 124 pounds per square foot by tearing
of the lower, or ten ion, flange at t he wing attachment
fittings.
The distribution of stresses due to chord and beam
loads was found to be determined at the sections in-
,·estigated by the flexure formula S =~c in spite of the
fact that the webs, or shear members, of the box beam
were trusses.
The moment of inertia that gave the best, and a
conservative, agreement between computed and
measured valnes of tresses due to beam loads was
determi ned IJy the box section alou e, neglectin!!; the
thin covering when it was in compression.
For chord loads !,he moment of inertia of !,he section
about an axis perpendicular to the chord was bes!,
determined by considering the leading edge, the box
section, and half of the trailing edge.
The stresses due to chord and torque loads were
found to be relatively small and need not be considered
unless at some section the stresses due to beam loads
give very low margins. They will, in general, be with­in
the accuracy of calcuJations based on the section
properties mentioned in the forego ing paragraphs.
The leading edge was found to contribute about 11
per cent at the tip and 35 per cent at the root of the
total torsional rigidity of the wing. The contribution
of the trailing edge was about 12 per cent at the tip and
about 5 per cent at the root.
The leading and trailing edges did not contribute
s ufficient to the beam rigidity of the wing as compared
with the box section to warrant their consideration as
structural members. However, they must be designed
to resist appreciable stresses.
(1)
IV. DESCRIPTION
The Materiel Division 55-foot wing is of all dural
(commercial 17ST, Air Corps specification 57- 152- A)
construction, metal covered; employs corrugated com­pression
members extending entirely across the area
usually defined by the spars. Thin trussed webs a re
used instead of spars, and no drag truss is employed,
the covering acting in this capacity.
The general characteristics are:
Span __ ____________________________ feet __
Root chord _______________________ inches __
Tip chord ____ __________________ ___ do ___ _
Root-thickness ratio __ ___________ per cent_ _
Tip-thickness ratio _________________ do __ _ _
Airfoil, Mod. U. S. A., 35 approx.
Unit dead weight ___ pounds per square foot __
Wing loading __ ______ ______ ____ ____ do ___ _
55
131
71
21. 8
13. 5
1. 827
8. 0
The main structural member of the wing is the central
box beam. The flanges of this beam are of corrugated
duraluminum of 2Yz-inches pitch and %-in ch depth,
and vary in thickness on the upper flange from 0.064
inch at the root to 0.020 inch at the tip. The lower
flange employs but two thicknesses, 0.020 inch at the
root, and 0.016 inch at t he tip. Figure 1 shows the
exact variation of these t hicknesses. It will be noted
from this figure that the 0.016 and 0.020 inch sheet has
been applied in two strips spliced along the center of
the box beam in a lap joint. This was made necessary
by the limited available widths of the thin material in
the lengths used. The splices of all the other sheets
were made across the wing in the manner. illustrated
in Figures 49 and 54. Figure 50 shows half of the lower
flange material in place. Figure 49 shows the upper
corrugated flange before the covering has been applied.
The two parallel webs, spaced 45 inches apart, are
made of flat sheet lightened as far as station 13 to form
a double Pratt truss. Beyond station 13 flanged
round lightening holes are employed. The gage of the
front web is 0.051 inch to station 14 and 0.036 inch
from station 14 to t he tip, and that of the rear 0.036
inch throughout. The width of the truss diagonals
varies from 2Yz to 4 ~ inches. The chord members of
the truss are of L shape, 0.101 inch gage, 1 inch width,
and approximately 2 inches depth. The same section
is employed throughout the span of the wing. The
spacing of the vertical truss members of the webs is
such that two bulkheads, or ribs, are included in one
truss bay. The vertical members of the webs are 3/4
by 3/4 inch by 0.051 plain duraluminum 90° angles.
Four of these are placed back to back at alternate
main ribs. At each of the false ribs in the nose portion
a single angle is used, and at alternate main ribs two
plane angles are employed to join the ribs to the webs.
On t he rear web single angles are omitted, stiffeners
being placed only at the main ribs. The stiffeners on
the outside of the box are flanged outward at the top
2
and bottom to clear the L-shaped chore! members.
The same stiffeners were used throughout the span,
there being no decrease in gage toward the wing tips.
The bulkheads were made up in t hree sections, a
nose portion, a center section, and a tail section. The
webs of the bulkheads were of 0.014 flat cluraluminum,
lightened to form a double Pratt truss having a single
vertical member midway between the main webs. This
lyre-shaped vertical member was the only departure
in the construction from a plain angle stiffener . The
bulkhead flange angles were 3/4 by 3/4 inch by 0.051
plain angles throughout the periphery of the wing.
Two bays of double Pratt truss lightening were em­ployed
aft of the rear main web, and round fl anged
lightening holes used to the rear of the Pratt tru ss.
This construction was used to and including rib No.
13. Outboard of station 13 onJy round lightening
holes were used. Figure 48 shows the main and false
ribs in place prior to placing the lower flange.
The nose bulkheads, so clearly shown in Figure 55.
consist of an 0.014-inch web lightened with a large
flanged hole and stiffened with a % by % by 0.020 inch
stiffener. Figure 55 also shows t he method of attach­ing
t he nose portions of the bulkheads to the webs, and
in the lower left-hand corner of the picture is shown the
0.051-inch gusset plate used to join the bulkhead flanges
and the web vertical stiffeners.
The leading-edge covering was unstiffened 0.020-inch
flat stock. All other covering material was 0.014-inch
flat stock. Figure 51 shows the method of applying the
covering to the lower chord of the box portion and Fig­ure
3 shows the covering schedule.
The construction of the wing tip may be seen from
Figures 48 and 50. This tip was not designed to be
flown , but only to carry a sand load.
The wing attachment fittings were of Air Corps
specification 57- 136- 8 steel, heat treated to 100,000
pounds per square inch. These were riveted to the
webs at station 1 by %z dural rivet s, this bulkhead
being reinforced accordingly. One finger plate of the
wing attachment fitting was passed directly through each
of the lower flange angles. Figure 46 shows the front
fitting at bulkhead No. 1 and the disposition of the
rivets securing it to the web. Figure 43 shows the
construction of bulkhead No. 1 at the fittings.
The main fittings at the center section are but 25
inches apart. The dimensions of the wing are those of
a %-scale model of a large wing at one time considered
for a heavy monoplane bomber, which explains the
narrow center section.
The dead weight of the wing, 1.827 pounds per square
foot, is the dead weight of the wing as tested, and docs
not include any gas tanks, wiring, control cables, or
rods, or the aileron structure, nor does it include any
allowance for a protective coating. Such a coating
would, together with the items mentioned above, be
necessary on a similar wing to be placed in service.
3
1
2 178 JT JOZ=r37j~
I
I I I 1 II I] l 'c
II II JI~ II II II
n "" \
J
~
.::.-:::~-:-:: :::_-_ ::::-:-::: ::·:::::;- ~:-::::~ : Ill
~
3
4
Li.
+ 0 1 z .3 5 6 7 tJ ~ 10 11 ' 12 13 11' 15 16 lJ
I
--~ -- ~ r--,f
~- ~ -- _L--- 13 __, -16 J ()(j i..----- 16 - IJLL /i,<uvc£ SPL/CES A.RE fJurr Jo//vrs Exc£PTJO//VTS AA d- tJfJ
OPLI CEJ ( /VOT SJ-tonw) OCCUR Ar £,,;c /'( .R/.8 IHBOARO OF #I z
fi'JG UJtE L- Upper flangc- l\Iatericl Div ision 55-foo t wing
Tu ..... ... .... ........ .. . .
r;I l
7 {} 9 JO
All s_p!c'c:es are !dp joud.s
.FIG URE 2.- Lower llaage--i\l atcriel Divis ion 55-foot wiag
Jf.5
4
.#.Z 138 r7j
I
I II 11 b ti II I I,
I I '
11
I
116 11 I
1111 1W 1115 117
II 11 ii I
M I
I
I
I
Leadi'ng edge qovrzrlng .OZO gage.
All oflzer coveruzg .014 gag_e .
.SJ>!icing .:pui gage off 9p a.rza bo{fo.m are ilz,e .s.uzzrz .
.SmrJ.le nvetea lap ;oints used on covering.
Pi:lch, of corraga ¥[on. ip- Z j
FIGURE 3.-Covering schedule-Materiel Division 55·foot wing
FIGURE 4.-Determination of equivalent wing-l\Iatfaiel Division 55-foot wing
5
V. COMPUTATIONS OF TEST LOADING
Dead weight of wing ______________ __ pounds __ 772
Wing area ________ __ ______ ______ square feet_ _ 422. 6
Design wing loading pounds per square foot_
Design load factors-
High incidence_____________________ 8. 5
Low incidence __ ___ ___ ______________ 5. 5
Inverted flight____________ _______ ___ 3. 5
Center-of-pressure locations-
High incidence __________ per cent C__ 30
Low incidence ___ _____ ______ _ do ____ 50
Inverted flight_ ______ __ ______ do __ __ 30
Thickness ratio root , 1
2
3
8i:3
8
4= 21.8 per cent at station 1.
Thickness ratio, tip, 97~3 =13.5 per cent at station 17.
No data on airfoi l or C/B ratio were available in the
wing-stress analysis. Since the airfoil is close to a
U. S. A. 35 section, a C/B ratio of - 0.160 was used in
high in ciden~e or an inclination of the wing of -9.2.
For low incidence an inclination of 8.6° corresponding
to a C/B ratio of 0.15 was used.
-nit gross wing load= 422.6 square feetX 8 pounds
per square foot = 3,381 pounds.
Unit net wing load = 3,381- 772 = 2,609 pounds.
Consider an equivalent rectangular t ip A- A, as in
Figure 4, so that the same wing area is maintained.
Thickness ratio at equivalent tip A- A
= 13.5- (21.8- 13.5) x ;~g
=13.5-0.875
= 12.6 per cent.
Thickness ratio at station 15
= 13.5 + (21.8-13.5) x ~~g
= 13.5+ 1.25
= 16. 75 per cent.
An equivalent wing will be determined such that
when covered with a uniform load it will give beam
loads that are proportional to the thickness ratio.
Chord at section A- A.
Equivalent chord= ~i:~ X 63 = 36.4 inches.
Chord at section B- B.
Wing chord=80 inches.
Equivalent chord=~~:~~X 80 = 61.5 inches.
A load tapering from Wat station 15 to .SW at the
tip section A-A will be used, W being the loading over
the equivalent wing.
No load wiU be placed on the center section between
stations 1 and 1.
Area section CC-BB:
A 1
13
1.3i 61.5X 238=22,943 square inches.
Area section DD- CC:
A2= 12.5X 131.34=1,642 square inches.
Area section AA- BB:
A3=
6
1.
5! 36
·
4
x 71 .5= 3,500 square inches.
Total area, equivalent wing:
Outboard station 1= 26,443 square inches.
Outboard station 0=28,085 square inches.
Considering a mean loading over the t ip portion of
.9W:
22,943W+0.9 X 3,500W= l,305 pounds.
22,943W+3,150W = 1,305 pounds.
26,093W = 1,305 pounds.
W =0.050poundpersquare
inch.
= 7.2 pounds per square
foot.
Considering the dead weight of the wing distributed
over area AA-DD:
772
Wc1=2 x 28.085=0.01373
TABLE 1.- Calculation of shears and moments
Section -----------------I A I B I c I D I E F I G I H I I J I K I
Reference section B-B Reference sec-tion
A-A
Distance from reference station ____ ------- 224 1 ~5 16.5 135 105 75 45 15 53. 5 I 17. 75 Length of section ______________ ---------- 28 30 30 30 30 30 30 30 36. 0 35. 5 Chord, equivalent wing __ _____ 131. 34 127. 2 118.8 109. 9 JOI. J 92. Running load __ _______________ 3 83. 5 74. 7 65. 9 53. 3 42. 0 ------- --- 6. 36
I
5. £4 5. 50 5. 05 4. 61 4. 18 3. i3 3. 29 2. 53 1. 74
Load in section per factor_ ____ ---------- 187 176 163 151 138 125 I 11 98 91 63 Running dead weight_ ___ _____ l.hO l. 75 l. 63 l. 51 l. 39 1. 27 1.15 l. 03 0. £0 0. 73 0. 57 Dead weight in section ____ ____ 54 52 49 45 42 38 34 31 27 26 20
Corrected W• ----------------- 50 48
I
45 42 39 35 31 29 25 24 l
First load, W--- --- -- ------- -- ---------- 139 131 121 112 103 94 82 73 67 45
Other loads per factor_ ________ ------ ---- 187 176 163 151 138 125 111 98 91 63
Shearfload __ --------- _____ -- --- - 967
I
828
I
697
I
576
I
464 I 361 I 267
I
185
I
112
I
45
Other loads. _- -- -------------- 1, 303 1, 116 940 777 626 488 363 252 154 63
86802-32- -2
6
30f. C
I fn vr?rted fli#zi C.P. c :JO % cho.n:i
-=ii• :.e-ll
Lorr i..ricz'dence CE - 501- chord
FIGURE 5
TABLE 2.-Unit shears and moments, high incidence
First 2 factors I Single factor I Half factor
Section Shear (pounds) Moment (inch- Shear (pounds) Moment (inch- Shear Moment (inch-pounds)
pounds) (pounds) pounds)
Beam Chord Beam Chord Beam Chord Beam Chord Beam Chord Beam Chord
--------- ----------
K
106.5 19 1, 882 307 62 10
J
l, 105 li9 3l 5 552 90
262. 0 43 8, 500 1, 380 152 25
I
4,950 804 76 12 2, 475 402
430. 0 70 18, 900 3,060 248 40
H
10, 980 1, 780 124 20 5,4£0 890
621. 0 99 34, 650
0
5,650 358 58 20, 050 3, 250 li9 29 10, 025 1, 625
838.0 136 56,550
F
9, liO 482 78 32, 650 5, 300 241 39 16, 325 2, 650
E
1, 075. 0 175 85, 200 13, 800 618 100 49, 250 7, 975 309 50 24, 625 3, 987
1, 338. 0 217 121, 700 19, 700 767 124
D
69, 900 11, 350 383 62 34, 950 5, 6i5
c 1, 617. 0 262 165, 900 26, 850 927 150 95, 300 15, 480 463 75 47, 650 7, 740
B
I, 920. 0 311 218, 500 35, 400 1, 100 178 122, soo 19, 920 550 89 61, 450 9,960
A
2, 240. 0 363 281, 000 45, 550 1, 287 203 158, 600 25, iOO 643 104 79, 300 12, 850
7
TABLE 3.- Calculation of bending moments
Firs t 2 factors Single factor I L ength -- - -- -
Eection or sec- Moments l\Ioments
tion= x Load in Sb ear section Load Shear w = LW sx wx · wx 2 Total SX - 2- Total
----- --------- --- --- --- -------------I
K 35. 5 108 63
108 0 1, 920 1, 920 63 0 1, 120 l , 120
J 36. 0 158 91
266 3, 994 2, 845 8, 605 154 2, 270 l , 638 5. 028
[ 30. 0 170 98
436 7, 980 2, 550 19, 135 252 4, 620 1, 470 11, 118
11 30. 0 19-1 111
630 13, 090 2, 910 35, 135 363 7, 560 I, 665 20, 343
G I 30. 0 220 125
850 18,900 3, 300 57. 335 488 10, 900 l,875 33, 118
F 30. 0 240 138
1, 090 25, 500 3, 600 86, •135 626 14, 650 2, 080 49, 848
E 30. 0 264 151
1, 354 32, 750 3, 960 123, 145 777 18, 800 2, 265 70, 913
D 30. 0 284 163
1, 638 40, 600 4, 260 168, 005 9-10 23, 300 2, 445 96, 658 c 30. 0 306 176
l , 944 49, 100 4, 590 221, 695 l , 116 25, 200 2, 640 124, 493
B 28. 0 326 187
2, 270 58, 400 4, 890 284, 985 1,303 33, 500 2,805 160, 803
A 12. 5 0 0
TABLE 4.- Bending moments- Beam direclion
HIGH INCIDENC E
St a tion 2 4 6 I 7 8
---
K
1, 882 4, 09'2 6, 302 7, 407 8, 512
J
8, 500 18, 400
I
28, 300 33, 250 38, 200
I[
18, 900 40, 860 62, 820 73, 800 84, 780
a 34, G50 74, 750 114, 850 134, 900 154, 950
F
56, 550 121,850 187, 150 219, 800 252, 450
85, 200 183, 700
E
232, 200 33 1, 450 380, 700
D
121, 700 261, 500 401, 300 471, 200 541, 100
c 165, 900 356, 500 547, 100 642, 400 737, 700
218, 500 464, 300
B
710, 100 833, 000 955, 900
281, 000 598, 200
A
915, 400 1, 074, 000 1, 232, 600
TABLE 5.- Bending moments at station 1 due to loads
at station 17
[Dis t a nce between station 1 and station 17= 280 in ches. Dis t ance be­tween
torque cables=229.5 inches]
Chord-load test Torque due to
platform loads
Beam -load tes t
I Torque
Load a t / M Pla t-s
t a tion oment a t form
17 s t a tion 1 load
250 70, 000 500 140, 000 75 17, 194
500 140, 000 1,000 280, 000 150 34, 388
750 210, coo 1, 500 420, coo 225 51, 581
l, 000 280, coo 2,000 560, 000 300 68, 775
1, 250 350, 000 -------- - --------- -- 375 85, 969
VI. INVERTED FLIGHT
During all the tests on the 55-foot wing it was sup­ported
from two 12-inch steel I beams. Two 4-inch
strnctural steel angles about 2}'z feet in length were
milled to slip into the wing fittings, to which they were
secured by %-inch aircraft bolts. The strnctural angles
8. 5 9. 0 11 10 15
9, 064 9, 616 11, 840 10, 720 16, 245
40, 675 43, 150 53, 130 48, 100 72, 850
90, 270 95, 760 117, 860 106, 740 161, 640
164, 975 175, 000 215, 340 195, 000 295, 300
268, 775 285, 100 350, 620 317, 500 481, 000
405, 325 429, 950 529, 050 479, 200 725, 450
576, 050 611, 000 751 , 810 680,~ 1, 030, 400
785, 350 833, 000 1, 024, 960 928, 300 1, 404, 800
1, 017, 350 1, 078, 800 l, 326, 200 1, 201, 700 1, 816, 200
1, 3ll, 900 J, 391, 200 1, 710, 600 1, 549, 800 2, 342, 800
were secured to the I beams by four %-inch cold-rolled
steel bolts, as shown in Figure 71. Four of the %-inch
bolts carried the entire wing load, acting in double
sllear.
In the inverted-flight test the wing was mounted with
the lower surface down and horizontal, i. e., C/B=O.
Figure 64 shows the timber framework used to support
the I beams. The loading was applied in two sections,
as shown in Figures 5 and 8. The center of pressurn was
at 30 per cent chord throughout the span.
Only the design load of 3.5 factors was applied in
this test, which was more of the nature of a proof test
than a strength test.
Strain-gage readings were taken at each of the
flange angles and at a station approximately midway
between the webs.
The deflections of the leading and trailing edges
were taken. With t he design load of 3.5 factors on the
wing the leading-edge lower surface buckled notice­ably.
The buckles practically all disappeared upon
removal of the load.
8
I I
'
r
I'>
~
~ ~
r f' , ,. t-;
..... ~ f--
,l :T 1
n ll't- ' I ,.._I
N.. tt.:~
_.,...
I
k I· ~~ 'i: ~ ~ iD
t?
FIGURE 6
a~ . ~ ~ 11 - ,, .
t-t p. r•
1- l -· __,. ~ tS
~
'
-1- ~ - - μ: - -
-
~
~
- ~
I
~
' ' '
~
' ~~ -f' · ~ I+ ~
- -
~
~ ' ' I
. , ~~ - I f'i' "' :±: ~"' ·- ~
"!- ~~ "' J
' ff ~~ B: I I _:;:: I I!
I ' I '
l' ~v~ I+ J ~ r~ ~ - I ~ f'~
~
FIGURE 7
9
TABLE 6.-J nverted flight test-Beam deflections
LE.\DTNG EDGE
I
Station
Load
factor Left wing flight winrr
-,I 'rip- I 1-7 -13 -9 I 6 3 3 6 9 13 17 'J'ip I --------- --- --- --- --- -
o o o o o I 0 0 0 0 0
I
0 0 0
~00 1.00 .~ . ~ .w . 22 . 18 . IO . 15
I
. 25 . SS . iO .~5
3. oo I I. 75 l. 31 I. 07 . 62 . ~8 . 29 . 20 . 25 . 60 .Pl 1.40 I. 40
~oo ~u 1.00 1.m .~ . 43 .30 . 20 . ~o . 65 . gq I. 19 I. fi2
-- J__ ---
' On elastic axis.
'rR AILlXG EDGE
0 ------- - 0 0 0 0 0 0 0 0
I
0
I
0
2. 00 ------ -- . 82 . 55 . 30 . 15 . 05 -.05 . 05 . 15 (' ) . 70 --·---
3. 00 ------·- l. 50 . 92 .51 . 25 -. 10 - .O!i . JO . 28 (' ) I. 24 1-·---- 3. 50 ------- - l. 81 l. 13 . 59 . 27 -.IO . 00 . 14 . 31 (2) I I. 45 --- ---
' Scale knocked off.
TABLE 7.- Inverted-flight test-!Ylean beam deflections
LEADlKG EDGE I '!'RAILING EDGE
- -
Load Station factor
I I I I 9 I 13 I 3 G 9 13 17 3 G 17
- - - -------- - -- --------
0 0 0 0 0 0
I
0 0 0 I 0 0
2.0 .140 . 185 . 320 . 59 . 795 0 • LOO . 225 ( ') . 7G
3. 0 . 245 . 315 . 610 .91 l. :155 -. 075 . l75 . 395 (') I. 37
3. 5 . 250 . 365 . 695 1.13 1.695 - . 050 I . 205 . 450 ( ') I. 63 I • I
t Scale on r ight win g k nocked off.
LOADS
Load F R }' R F R I F R _:_I_:_ F R _:_!_:_ I F R F ~'. :___:___/ factor
-------- ----,---- - -- --
2 163 163 153 153 142 142 132 132
120 I 120 llO 110 97 97 85 85 79 79 54 54
I 3 93 93 88 88 82
82 1
75 75 69 69 63 63 55 55 49 49 46 46 31 31
I 3. 5 47 47 44 44 41 41 38 38 34 34 32 32 27 27 25 25 23 23 15 15 i
FrGURE 8.-Loading schedule, inverted flight, J\faieriel Divis ion 55-foot wing
10
f'
FIGURE 9
VII. LOW-INCIDENCE TEST
In the low-incidence t est the wing was mounted
with the upper surface down, the lower chord inclined,
and the trailing edge down 8.5°, giving a C/B ratio
of 0.15.
The I beams were again supported by the trestles
of heavy timber.
The center of pressure was located in this test at 50
per cent chord throughout the span.
A 2-section loading was again employed, being ap­plied
in accordance with the loading schedule shown in
Figure 10. Only the design load of 5.5 factors was
applied to the wing.
The deflections of the leading and trailing edges of
the wing were taken for each load application, using
suspended scales and a Y level to read them.
At 5.5 factors the only visible wrinkles were on the
upper (down) surface of the wing just aft of the rear
web. These were apparently shear wrinkles caused
by the chord loads and ran diagonally between ribs
from one rib at the rear web toward the trailing edge
at the adjacent inner rib. These, while visible, were
not at all serious.
T ABLE 8.-Low-incidence test- Beam deflections
LEADING EDGE
Station
Left wing R ight. wing
Load
~~ I '
Tip 1 11 15 12
1
9 6 I ' ; ' • • " " I n ' '""
I LE FW
-o- o o o _ o _,_o--o -j,-0--0 -.-0 - o _ o _l_o _ o - 0--0--0 -
2.0 0. 89 .69 .54 . 34 1 . 22 . 14 .02 .02 .JO . OS .JO .12 .34 .50 .64 .86
3. 0 1. 72 1. 40 l, 17 . 81 . 54 . 37 . 21 . 14 . 10 . 10 . 12 . 21 . 44 . 65 . 84 1. 16
4. 0 2. 38 1. 87 1. 51 1. 05 , . 69 . 45 . 26 . 17 . 14 . oo . 19 . 31 . 60 ' 93 1. 19 1. 67
4.5 2.65 2. 03 1.70 1.16 1 .74 . 47 .28 .16 . 15 . 12 .20 .39 . 72 1.13 1 1.47 1.96
5. 0 2.£0 2. 25 1.84 1.24 .79 .48 .29 .16 .1 5 . 12 .25 .46 . 84 1.31 1.69 2. 30
~5 ~~ ~il 1.00 1.D _g . 54 .W .IB . 16 .IB .W .M .00 1.M l.H ~M
1 On elas tic axis.
Load
factor
Tip 17 15
11
TABLE 9.-Low-incidence test-Beam deflections
'l' RAILING EDGE
Station
Left wing Right wing
---------------------·
12 9 6 3 6 9 12 15 17 Tip I
0
TE RW --1---------------1 0 --- - ----- 0 0 0 0 0 0 0 0 0 0 0 0 0 0 ------- -
2. 0 ----- ---- .82 .71 .54 .37 .31 .17 .14 .09 .20 .29 .37 .50 . 68 .80 --------
3.0 1. 61 1.40 1.11 .82 .62 .43 .34 .15 .31 .41 .53 .70 .95 1.12 - ------· ·
4.0 --------- 1 2.21 J.90 J.49 1.08 .80 .56 .44 .19 .45 .60 .81 1.07 1.38 1.60 --------
4.5 - -------- 2.47 2.12 1.64 J.21 .89 . 63 .48 .24 .51 . 70 .93 1.25 1.63 1.90 -- - --- --
5.0 --------- 2.76 2.30 1.80 1.32 1.00 .70 .56 .26 .60 .81 1.09 1.44 1.88 2.20 -- --- ---
5.5 ---- - ---- 2.92 2.52 1.97 1.44 1.08 .76 .63 .29 .70 .92 1.24 1.68 2. 16 2. 52 - ----- --
Load
factor
'
- ----~
TABLE 10.- Low incidence- l\1ean beam deflections
LEADING EDGE TRAILING EDGE
Station
__ _ o_ \_3_ __6 _ _ 9_ _1 2_~ 1_1_ __o _ __3 _ __6 _ __9 _1 2_ ~ _11_
0 0
. 02 I 0
2. 0 . 050
3. 0 . 14 I . 155
4. 0 . 17 .180
4.. 5 . 16 . 200
5.0 .16 . 205
5. 5 . 18 . 235
I
Load F R l" !actor
------
2 163 163 153
3 93 93 88
4 93 93 88
4. 5 47 47 44
5. 0 47 47 44
5. 5 47 47 44
0 0 0 0 0 0 0 0 0 0
. 120 .170 . 340 . 520 . 665 . 14 .185 . 300 . 370 . 520
. 245 . 375 . 625 .920 1. 120 . 34 . 370 . 515 . 675 . 905
. 320 .500 . 825 ]. 220 1. 530 . 44 . 505 . 700 .945 I. 280
. 335 .565 . 940 1. 415 1. 775 . 48 . 570 . 795 1. 070 I. 445
. 365 .625 I. 040 1. 575 J. 970 .56 . 650 . 905 J.205 l. 620
. 420 I . 695 1. 155 I. 750 2. 175 . 63 . 730 1. 000 1.340 ]. 825
LOADS
I ~' R I F I R I \_:_ R F R 1'' H F f n F
------------------ -
153 142 142 132 132 120 120 110 llO 97 97 85 85 79
88 82 82 75 75 69 69 63 63 55 55 49 49 46
88 82 82 75 75 69 69 63 63 55 55 49 49 46
44 41 4.l 38 38 34 34 32 32 27 27 25 25 23
44 41 41 38 38 34 34 32 32 27 27 25 25 23
44 41 41 38 38 34 34 32 32 27 27 25 25 23
FIG ORE 10.- Loading schedule, low incidence, Materiel Division 55-foot wing
0
. 695
1.175
1. 640
1. 875
2. 090
2. 340
R
0
.810
1. 365
1.905
2.085
2.480
2. 720
1+.f
211·
4
21.L"
4
F H
------
79 54 54
46 31 31
46 31 31
23 15 15
23 15 15
23 15 15
12
t±
I
I
μ
t-l--+--!-+-
1+~.;-+-"~'-1-t-i-+-IHJff
i+
\I I
~j I -- - ~
it j -
1-
H:tttt+t=l+t=f++:tu~T2~-~~~T~H~ ~ j-:;_+ ~ ~+-+,-,~..~~IH-1H-1-+-.1J_· .._..l~~r-+-rH
,_J_,-, '' ,,- ,1"1'
=+
I• I !-"""'' I
(T I
11 I
fi4 I 'L
't'
_J_
Hi! '
,. ! _J ~
FIGUHE 11
VIII. HIGH-INCIDENCE TEST
In the high-incidence test the wing was supported by
the two 12-ineh I be!1ms as in the previous tests; the I
beams in turn were mounted on heavy timber trestles.
The wing was inclined, leading edge down and lower
chord up, so that the angle between the chord and the
horizontal was 9.1°, giving a chord-to-beam ratio of
0.160; that is, 9.1°=tan-1 0.160.
The loading was applied in three sections throughout
the span . The division of the load is shown in Figure 5.
In order to speed up the loading, the loads were applied
in units of two factors up to six factors, then in units
of single factors up to eight factors. Above eight fac­tors
the load was applied a half factor at a time. The
loading schedule is given in Figure 13.
Deflections of the leading and trailing edge were
noted by means of telescopes, and scales were suspended
from the '\'ling. The location of these scales \Yas sym­metrical
about the longitudinal axis of the wing in
order that any rotation of the wing about its supports
could readilv be detected.
Jacks we~e placed beneath the webs at frequent
intervals and the load applied to the wing by lowering
~j~ks. ·
During the application of the loads nothing of inter­est
developed up to six factors. At this load it wa
noticed that the leading edge was buckling slightly.
When lowering the jacks with seven factors in the
wing, a loud "pop" was heard coming from th_e vicinity
of the center section of the wing. No failure was
apparent, and the wing supported the load without
difficulty.
At 8.5 factors, the design load for the wing, the only
signs of high stress were the wrinkles along the compres­sion
(true upper) side of the leading edge.
While taking deflection readings with nine factors on
the wing another "pop" was heard, apparently com­ing
from the same region as before. The wing settled
slightly, making it necessary to read the deflections
over, but supported the nine factors. At th is load the
leading edge was still buckling.
At nine factors the extreme trailing edge was appar­ently
carrying an appreciable tension load. On the
upper surface of the wing in the vicinity of the root
section compression buckles could be seen just aft
of the rear web. These buckles merged gradually
into the taut-tension stressed skin toward the trailing
edge.
While lowering the jacks with 11.5 factors on the
wing, the right front fitting holding the structural
angle to ·the I beam failed by stripping the threads on
one of the %-inch tie bolts. This failure was due to
fa ulty threads on the tie bolt.
The jig failure occurred on Saturday afternoon. The
entire load was left on the wing until the following
Monday morning. After its removal no permanent
set could be detected by sightin g along the maximum
ordinate line of the upper surface. The only signs of
distress noted after the load had been removed were
in the leading edge and in the 0.014-inch flat covering
of the box portion of the wing, both being on the lower
or tension surface of the wing. This metal was visibly,
but not as yet seriously, stretched near the root section
of the wing.
13
While the load was removed, splice sheets ofJ the
trailing-edge strip at station 1 were reinforced with
additional rivets. Also additional Kalon screws were
placed in the flat skin splice at station 9, along the
lower chord. At the same time two inspection open­ings
were cut in the upper surface of the trailing-edge
portion of the wing at approximately the point where
the stress changed from tension to compression. These
openings were placed between stations 0 and 1 on either
side of the center line and permitted the observation
of the fittings and structural members in that vicinity.
New tie bolts were made and installed, and 11 factors
replaced in three divisions of load, each being approxi­mately
3% factors.
No deflections were taken during this loading, except
the tip deflections at 11 factors, and each successive half
factor above 11 factors.
With 12 factors on the wing pronounced shear
wrinkles could be seen extending diagonally across the
trailing-edge portion of the wing on the upper surface.
At this load the flat skin over the compression flange of
the box portion of the wing was buckling between rivet
heads over the entire wing, being most pronounced at
the root section . The leading edge was continuing to
buckle but was not failing. At this loading an inspec­tion
of the internal structure at the root section revealed
nothing of importance.
With 12.5 factors on the wing the rear vertical truss
members of bulkheads 0 and 1, right and left, were
observed to be buckling. Also the flat material of the
web was buckling away from this vertical member,
having buckled to such an extent that it contributed
nothing to the strengtJ,i of the vertical member. ·
At this load a relieving cradle was built and placed
on jacks beneath the center section of the wing. An
analysis of the fitting showed them to be good for 12
factors, and rather than chance a fitting failure the
relieving cradle was put in place to insure a main struc­tural
failure. This cradle is shown in Figure 76.
At 12.5 factors the leading-edge wrinkles were quite
deep and were inclined between the ribs, going from the
front web toward the leading edge at the next rib
inboard, at the root. This condition became Jess ap­parent
toward the wing tip.
Between 12.5 and 15 factors nothing of importance
developed. The leading edge continued to develop
wrinkles but did not fail.
While lowering the jacks with 15.5 factors on the
wing, it failed at station 1 on the tension flange, appar­ently
by the failure of one of the lower flange angles
in tension at the section where the fittings are attached.
Whether or not the front fitting actually failed fu-st is
not definitely known. The failure proceeded across
the lower corrugated flange to the center of the wing,
thence along the seam in the flat skin midway between
the flanges to a transverse seam, along which it passed to
the rear web. At the same time the leading-edge seam
at station 1 pulled open on the tension side of the wing.
This failure is shown in Figures 71 to 74, inclusive.
86802-32--3
At failure the wing dropped to the jacks, which were
nearly clear at the time, and rested on them.
After the failure the leading edge was nearly as free
from buckles as it was prior to test, all signs of stress
having disappeared. They were unfortunately not
photographed prior to failure, though some conception
of their appearance may be had from Figure 63, which
shows mild buckles that formed during the beam-load
stress-distribution study.
A secondary failure also occurred at station 5, where
a transverse seam on the tension flange of the trailing­edge
portion failed. This failure is shown in Figure 75.
After the failure the load was removed from the wing,
and the leading and trailing-edge covering stripped
from the wing in the vicinity of station 1, so that the
condition of the internal structure could be determined.
FIGURE 12.- Distortion of front upper flange
angle, high incidence
An inspection revealed that the front upper flange
angle had started to fail along the leg tb which the
web was riveted.
This failure was in the vicinity of stations 3 and 4
and was limited to the internals between nose ribs, and
may have been due to the wing dropping to the jacks.
Such a failure was noted only on the failed side of the
wing. The compression member of bulkhead No. 2
midway between the webs was badly bowed, having
failed as a column. This failure was probably second­ary,
and would not have occurred due to air load8, where
the greater part of the load would have been applied to
the upper surface of the wing. Several of the gussets
used to attach the bulkheads to the webs showed signs
of having been highly stressed. At the upper flange
angles the gussets were stressed in compression, and at
the lower flange angles some of the rib webs either
started to tear or showed signs of fairly high tensile
stresses.
A few of the rib vertical members in the trailing-edge
portion at the center section buckled. It was noticed
prior to failure that these members were at about their
point of failure.
Some of the failures just mentioned were probably
secondary in nature, having occurred as a result of the
right-wing panel settling suddenly on the jacks after
the main failure occurred.
I
I
I
Load
factor
0
2.0
4. 0
6. 0
7.0
8.0
8. 5
9. 0
9. 5
Load
factor
r['ip 1
- --
0
. 745
1.610
2.640
2. 940
3. 510
3.800
3. 640
4. 015
Tip
17 15
------
0 0
.650 . 545
l. 430 1.185
2.661 1.915
-------- 2. 185
------ -- 2. 595
-- -- ---- 2. 795
-------- 2. 695
-- ------ -- ------
17 15
14
TABLE 1 l.- High-incidence test-Beam deflection~
LEADING EDGE
Station
Left wing I Right wing
0
12 9 6 3 3 6 9
LE FW
------------- ----------
0 0 0 0 0 0 0 0 0
• 390 . 205 .080 . 100 .095 .07 . 15 .20 .30
• 730 .500 .240 • 280 .215 .17 . 30 .50 . 72
1.355 .870 .480 . 400 . 345 . 23 . 47 . 71 1. 08
l. 550 . 995 .585 . 510 . 485 . 32 . 65 1.00 1. 47
1.840 1.190 . 700 .580 . 500 • 37 . 75 1. 10 1.61
1. 980 1. 295 . 775 .630 . 565 .42 . 77 1.17 I. 73
1.920 1. 285 . 780 .6 0 .665 . 44 . 95 1.41 2. 08
2.125 1. 410 . 875 . 720 . 695 (') . 95 l. 41 2.10
1 On elastic axis. ' Scale fell down.
TABLE 12.- High-incidence test- Beam deflections
'I'RAILING EDGE
Station
12
--
0
. 97
1.07
I. 62
2. 12
2.17
2.52
3. 02
3. 07
Left wing Right wing
0
12 9 3 9 12
RW TE
---
15 17 'J'ip I
-------
0 0 0
. 70 . 71 .91
l. 54 1. 75 2. 07
2. 30 2. 70 3.12
2. 95 3. 46 3. 93
3. 34 3. 89 4. 47
3. 55 4. 14 4. 74
4.15 4. 85 5. 51
4. 25 5.00 5. 65
15 17 Tip
----- ---
__ ,_ --- - ---------- - ---------
0 0 0 0 0 0 0 0 0 0 0 0 0 0 0
2.0 . 620 . 490 .33 . 205 .100 .065 . 05 .050 . 10 . 16 . 28 . 48 .63 . 75
4. 0 I. 345 1. 060 . 72 . 425 . 215 . 145 . 11 .100 . 22 .36 . 63 1. 00 1.44 1. 70
6. 0 2. 165 f: ~~ I 1.19 . 729 . 405 . 245 . 20 . 175 .31 . 54 . 91 1. 49 2. 10 2.58
7. 0 2. 370 1. 2b . 735 . 395 . 215 . 25 .170 . 36 . 67 1.17 1. 88 2. 65 3. 20
.o 2. 860 2. 270
1. 53 1
. 915 . 495 . 245 . 30 .175 . 40 . 75 I. 29 2. 09
3.00 I 3. 62
8.5 ::::::::1 3.060 2.420 I l. 65 . 980 . 535 . 275 . 30 . 220 . 41 . 79 1. 36 2. 20 3.14 3. 87
9.0 2.900 2. 260 l. 49 . 855 .440 . 215 . 32 . 190 . 47 .92 1. 61 2.58 3. 68 4. 50
9. 5 3. 210 2. 510 1. 67 . 985 . 510 . 260 . 35 . 215 . 47 . 92 l. 63 2. 62 3. 80 4. 65
'._
TABLE 13.- High incidence-J\!lean beam deflections
--
LEADING EDGE I TRAILl ;-\G EDGE
- Load
lac· Station
tor - -
I
0 3 6 9 12 15 17 0 3 6 9 12 15 17
--- - - --- --------------------- ------------ ---
0 0 0 0 0 0 0 0 0 0 o I o 0 0
2.0 . 095 . 125 . 140 . 253 .680 . 623 .680 .050 . 083 . 130 . 243 . 405 . 560 . 685
4. 0 . 215 . 290 . 370 . 610 .900 J. 363 1. 590 . 100 . 183 . 288 I . 52M .860 l. 250 l. 523
6. 0 . 345 . 4.35 . 595 . 975 l. 488 2.108 2. 681 .175 . 278 . 473 .820 I. 340 l. 928 2. 373
7. 0 . 485 . 580 . 793 1. 233 1.835 2. 568 -- -- -- -- .170 . 288 . 533 . 953 I. 565 2. 270 2. 785
8.0 . 500 . 665 . 900 I. 400 2.005 2.968 ----- --- .1 75 . 323 .623 1.103 I. 810 2. 635 3. 240
8. 5 . 565 . 700 . 973 I. 513 2. 250 3.173 -------- .220 . 343 .663 1. 170 I. 925 2. 780 3. 465
9.0 . 665 . 815 1. 095 1. 683 ~:~~ 1--~~~- .190 . 343 . 680 l. 233 2. 035 2. 970 3. 700
9. 5 . 695 . 835 1. 143 I. 755 . 215 . 365 . 715 I. 308 2. 145 3. 155 3. 930
15
M I R F IM ~) R ]' I l\I
I
F IM
_:_I Load F l\I R F .M R F M R F R F R F ?vl R F .1\1 R
factor I
---- - --- - -- - - - ----- - - - ---
2 163 82 l 153 77 76 142 71 71 132 66 66 120 60 60 110 55 55 97 1 49 48 85 43 42 79 40 39 54 27 27
4 186 93 93 176 88 88 164 82 82 150 ;s 75 138 69 69 125 63 63 110 55 55 98 49 49 92 46 46 62 31 31 I
6 186 93 93 176 88 88 164 82 82 150 75 75 138 69 69 125 63 63 110 55 55 98 49 49 92 46 46 62 31 31
7 93 47 46 88 44 44 82 41 41 75 38 37 69 35 34 63 32 31 55 28 27 49 25 24 46 23 23 31 16 15
8 93 47 46 88 44 44 82 41 4 l 75 38 39 69 35 34 63 32 31 55 28 27 49 25 24 46 23 23 31 16 15
8. 5 47 u 23 44 22 22 41 21 20 38 19 19 34 17 17 32 16 JG 27 ! 14 13 25 13 12 23 12 LI l6 8 s
I
9-15.5 ,, Same as above
FIGURE 13.-Loading schedule, bigh incidence, Materiel Division 55-foot w ing
FIGURE 14
16
1-1-~-!-4-+-'--+++++++++-1-H--H--j-+-l +t-H-H--t++-H-+-±Lol1 A+cl.+++•1-+-+-hH-.i++-+++r+~o~~·-~~~-
l?'-j., ,,~ 11 ~, , ~ t-t-t-1-trl-t-H
~ -
!' t-1-t-H-+-+-++
H--
FIGURE 15
TABLE 14.- Ti]J deflections on elastic axis- Sand-load
tests
INVERTED FLIGHT
Load Left Right Mean
factor wing wing deflection
--1I- ---- -
0 c 0 0
2. 0 1. 00 . 85 . 925
3. 0 l. 75 1. 40 1. 575
3. 5 2. 14 ]. 62 I. 880
LOW INCIDENCE
0 0 0
I
0
2.0 .89 .86 .875 3. 0 I. 72 1.16 1.440 I
4. 0 2. 38 1. 67 2. 025
4.5 2. 65 1. 96 2. 305
5. 0 I 2.90 2.30 2. 600 5. 5 3. 12 2.65 2.885 I
TABLE 15.-Tip deflections on elastic axis- High
incidence
FIRST 'l' E ST
Load Left Right
1
Mean j
factor wing wing deflection
- ----- ---.---
0 0 0 0
2.0 . 745 . 91 .828
4.0 1.610 2.07 1.840
6. 0 2. 640 3.12 2. 880
7. 0 2. 940 3. 93
I
3. 435
8.0 3. 510 4. 47 3. 990
8.5 3.800 4. 74 4. 270
9. 0 3. 640 5. 51 4. 575
9. 5 4. 015 5. 65 4.833
10. 0 4.100 6.04 5. 070
10. 5 4.300 6.40 5.350
11. 0 4.500 6. 80 5.650
I
TABLE 15.-Tip deflections on elastic axis-High
incidence- Continued
I Load
factor
0
11. 0
11. 5
12. 0
12. 5
13. 0
13. 5
14. 0
14. 5
15. 0
SECOND T EST
Left Right Mean
wing wing deflection
0 0
6. 01 5. 24
6. 47 5. 43
6. 77 5. 66
7. 05 5. 93
7. 32 6. 07
~: g! NI 0 ref 1rs1.
8. 54 7. 67
0
5. 625
5. 950
6. 215
6. 490
6.695
7. 315
7. 690
8.105
IX. LEADING-EDGE TEST
This test was the last to be made. The wing was
placed on the floor, with the lower chord up and hori­zontal.
A framework was built in front of the leading
edge to support a sheet of duralumin which acted as
a guide plate during the test. The set-up is shown in
Figure 78.
The section selected included stations 3 to 6~ and
was 54 inches in length. Seven nose ribs were included
in this length.
In applying the load to the leading edge four layers
of 25-pound shot bags were first laid in place, and on
these the remainder of the load was built up with 50-
pound pigs of lead.
The total supported load was 5,425 pounds. This
would correspond to a load of 100.2 pounds per inch, or
775 pounds per nose rib.
The action of the nose-rib vertical stiffening angles
was watched during the loading. About 50 or 100
17
: ~ - I- - ~~1
j ,
I t't ~ :z . l~,
::} I .. ~ I ' 7
I
,, I
I
- I =;i
J1J'.1 ,, f
- .
rt - -
- '
~ - IL ''~ d '
I ~, l
~ -~
.,
-· .. t' f -
' -
' -
-
-
~ - - - - - - -
-
FIGURE 16
pounds before the Jcading edge failed several of t hese
were observed to bow and flatten out slightly at their
centers. The web where the stiffeners were riveted to
it were observed to be highly stressed at this time.
The fin al failure was progressive, starting at the outer­most
rib and proceeding across the test section. This
failure is shown in Figure 79. There was no noticeable
deflection of the -leading edge prior to failure.
Using the handbook expression for computing
leading-edge loads
where
w= running load per inch per factor.
w0 = gross running load per inch = 7.84 pounds
per inch, average.
Lr= location of front spar in per cent of chord=
19.5 at station 5.
n = number of factors supported.
Using instead of w, W equal to the gross load
supported,
WX 20
n= w0 L1
_ 100.2X 20
-7.84X 19.5
=13.15
A pursuit wing would be required to support 14
factors on the leading edge, an observation plane 10,
and a bomber 6. The factor just determined, however,
is based on the wing loading of 8 pounds per square
foot, which would be low for the foregoing types of
airplanes.
X. TORSIONAL CHARACTERISTICS
In order to determine the torsional characteristics
of the wing, it was necessary to obtain a curve of the
twist of each section of the wing for a known applied
torque. This was done by applying a known torque
at station 17, near the wing tip, and obtaining leading
and trailing edge deflections by the use of suspended
scales and a Y level, as in other tests. In order to
obtain a check on the test results, deflections for several
known torques were obtained. The method of apply­ing
this torque to the wing has been described on page
21 of this report. Figures 56 and 58 show the torque
arms in place and the suspended scales. Figure 65
also shows the manner of deformation of the skin due
to torque loads.
The deflection curves are shown on Figure 17.
The coefficient of torsional rigidity was computed by
the following expression and plotted against the dis­tance
from the tip on the same figure.
where
dL
CTn= M dO
CTn = coefficient of torsional rigidity in pound­inches2.
dO=twist in degrees in section of length dL
inches.
.M=applied torque causing deflection do, in
inch-pounds.
Since the torque acted so as to cause the trailing edge
to move in one direction and the leading edge in the
opposite direction, these deflections must be added to
18
0 ·-
I\
\ 111d· 'eri~ ,,1 'vis cm j;5 ft. Fl t>rzg
~
~ ,L- ,,,._._
""
' ---
." c ~ ~- --
"" "<t 1• •v
a "" ~
"
s
K"'? " I'.. '\ ~-
1• vv
·- '
~ I" "o
~-
......
!"'-. - t...-:° LM
\:ii'. "' ~ ~ ... I..-" ~ v v
I~ "' ~ L..---
. ..,
~ ""' '-. ....- __..
Ill ~ ?s.;1• f"'-.-.._ ~ ! L/1 ~ ·- ....__ f""-...
......._ I..-" ~ '"'""
I'---. !'--.. i"--..... :---...... ~ I..-- ,.._.
~ ~" r---.... ~ I'---.. • ...... ~ i>< ...... "'
~ s-o"" I'---.. 0
,__ 1...--' i-. ~ "" ·- i--- t> ::::::;:: --... i---I o --...i--- '" -- r--,__ --- ...__ i--- i---
i..--i-- -1-- -ro- ---I--._ r--• r--1-- ~ i--,.. i..-- ,..__,,_ ~ r-- -...__ 1-- ..' - ""' ~ er - r-;.. r-- -~ I--I-- r-- I--1-- f::; t::-1-- . - r-: ~
-4~ 60 oo .11 •o ./. 'O ./!(.() .lM .1, 'O Z<Po z 0 ,i:'KJ .Z.PO z 'O _, ~o
n.: l ,t~- -- r_ j ' _ , L .1 - i
FIGURE 17
I
,. - -
- -
I
- +
I
LL
I
dl.i:t. r, .. 1-l• ~:):11"
I " ,_
l - I ,.,, . I' - U,~ I
+' -+ I ~ u+ I ~fl.Jr I
1-Y ~
FlGURE 18
19
obtain the total deflection of the leading edge with
respect to the trailing edge.
Table 16 gives the torsional deflections and the com­putation
of CTR·
Since the angle of twist was at all times very small,
the angle of twist in degrees can be accurately computed
by the expression
Figure 18 shows the comparative rigidities of the
55-foot wing and the XHB-3 wings, which were of
the same span but of lighter loading. The gain in
torsional rigidity is immediately apparent.
0 dLE+dTEX 57.3
Chord
TABLE 16.-MaUriel Division 55-foot wing
FIRST TORSIO 1'EST
17 15 13 11 Fittings
Platform load
F R F R F R F R F R F R F R F R F R FF RF
--------------------------------------
o __ ------------------------ -------- 12.13 8. 77 6. 96 8.00 4. 53 7. 72 1. 76 5. 84 6. 79 7. 91 6. 00 6. 30 4. 20 2. 61 4.11 2.16 4. 93 7.19 0. 61 0. 53
75_ - - -- - -- - - - -- - - - - - - - - -- - - - - - - - - - - 12. 06 8.88 6. 91 8.09 4. 50 7.80 1. 71 5. 90 6. 75 150 ___ _______ _____ ________________ _ 7. 96 5. 99 6. 31 4.19 2. 65 4.10 2.19 4. 91 7. 20 . 61 . 53
11. 95 8. 97 6.82 8.14 4. 41 7. 86 1. 68 5. 94 6. 71 8.00 5. 95 6. 36 4.16 2.68 ~091 225 _______ ____ _______ ___ ___ ________ 2. 20 4. 90 7. 20 • 61 . 53
ll.83 9.05 6. 74 8.21 4. 36 7. 92 I. 61 5.99 6. 681 8. 02 5. 921
6. 40 4.15 2. 70 4.05 2. 21 4. 90 7. 22 . 61 . 53
300. - - - - - - - - - --- - - - - -- - - - -- - - - - - - - - 11. 70 9. ll 6. 65 8. 29 4. 30 7. 99 I. 57 6.00 6. 63 8. 09 5. 90 6. 41 4. 11 2. 72 4. 05 2. 24 4. 90 7. 27 . 61 . 53
375. - - --------- -- ---- --- - -- -------- 11. 58 9. 20 6. 55 8. 361 4. 22 8. 04 l. 50 6.08 6. 60 8.12 5. 87 6. 47 4.10 2. 75 4. 02 2. 28 4. 89 7. 28 . 61 . 53
DE!o'LECTIONS
75_ - - --- - - - - - - - --- ------- --- - - - - - - - 0. 07 o. ul 0.05 0.09 0.03 0. 08 0.05 0.06 0.04 0.05 0.01 0.01 o. 01 0. 04 0. 011 0.03 0. 02 0. 01 --- -- -----
150. - --- ----- --- - - - - - - - ----------- - .18
. 201
.14 .14 .12 . 14 .08 .10 . 08 .09 . 05 . 06 . 04 . 07
. 021
. 04 .03 . 01 ----- -- ---
225 . --- - --------- --- - - ----- -- - - - --- . 30 28 . 22 .21 .17 . 20 . 15 .15 .11 .11 .08 .10 . 05 . 09 . 06 . 05 . 03 • 03 ----------
300_ - -- - --- - - -- - - ---- -- - - -- - - - - ---- . 43 • 34 .31 .29 . 23 . 27 . 19 .16 .16 .18 . ·10 .11 .08 . 11 .06 . 08 .03 . 07
375. - -- - - ---- - - - --- -- - - -- -- - - -- - - - - . 55 . 43 . 41 . 36 . 31 . 32 . 26 . 24 .19 . 21 .13 .17 . JO . 14 .09 .121 • 04 .09 ----------
I
FRONT DEFLECTIONS PLUS REAR DEFLECTIONS .
75_ - ---- - ---- ------------- --- - - - --- o. 18 0.14 O. ll 0.11 0. 09 o. 02 0.05 0. 04 0.03 --- ----- ---- 150_____ _______ ___ _________________ . 38 .28 . 26 .18 .17 .11 . ll .06 . 04 -- - ---------
225. - - --- -------------------------- . 58 . 43 . 37 .30 . 22 .18 .14 . 11 . 06 -- ---------- 300________ _____________ ___________ . 77 .60 . 50 .35 . 34 . 21 .18 .14 . JO -- ----- ----- 375_____ ________ ___________________ . 98 . 77 . 63 . 50 • 40 . 30 . 24 .21 .13 -------- ----
Distancebetweenstations_ ________ 40.37 42. 00 36.25 36.25 31.50 32.00 30.00 30.00 1----------------
r~~38~.~oooc---l..--c7~8~.3~7,--~I~ 1~20~.~37o--71~1~576.6~2=--~1-0109~~~8~7-'I ~22~4 .3""7,..-~1~2~5~6.~37=--~,~2""8"6"."3~7-l,--;3c1~6~.3"-7,..-71.---_---_-_ _-_
Distance from tip ..... . .......... . 58. 00 I 99. 00 I 138. 50 I 174. 75 I 208. 37 I 240. 00 I 271. 00 I 3ill.oo--c= .... :-.: ___ _
Chord .... . __ .•.• -- .. . --... _ •... _ .. 12.00 I 80.50 I 89.50 I 97.50
1
105.5 I 112.oo
1
119.oo
1
125.50
1
132.00
1
_____ ___ _
75. ------------------ - - - --------- - -
212550 _. _-_-_--__--__- -__-_-_--__-_--__-_- -__-_-_-_-_-_--_-__-_- -__- -_
300. - - - -- - - - - -- - - -- - - - - -- ---- --- - --
375. - - - - - - - - - -- - - - - -- -- - - - - - - - - - - - -
0.143
.295
. 541
. 613
. 780
75 ________ __ __ ________ _____________ 0. 031
125205__ _-_-_-_-_- _-_-_-_-_--_-_-__-_-_-_-_--__-_-_-_--__--_-__-_-_--_ .. 104742
300________________ ________________ .184
375__ ______ ________________________ . 222
75. - - - - --- ------- - - ------- - - ------- 1, 290
150. - --- - - --- --------- - - --- -------- 565 225_ ----- ------ --- ----------------- 278
300_ --------- -- -- - ------- --- ---- -- - 217
375_ - ---- -- - -- -- - - - - ----- - - - -- -- - -- 180
17,194
34,388
51,581
68,775
85,969
22. 2
18. 7
14. 3
14.8
15. 4
0.100
.200
. 306
.428
. 549
0.023
.058
.092
.135
.170
1,827
725
457
315
247
ROTATION (degrees)
0. 0704
.1664
. 2370
.3200
. 4030
0. 0648
.1060
.1768
. 2060
. 2940
dB (degrees)
0. Ol5 0. Oll
. 037 .030
. 054 . 040
. 073 .052
.1 02 . 072
dL/dB
2,420 3, 270
977 1,200
670 900
495 690
354 500
0. 0488
.0923
.1194
. 1845
. 2170
0.008
. 021
. 030
. 038
. 049
3, 980
1,520
1, 032
838
650
J.I dL/d8 (millions pound-inches•)
31. 4 41. 5 56. 2 68. 5
24. 3 32. 7 40.0 50.8
23. 6 3~5 45. 4 51. 0
21. 6 33. 9 47. 3 57. 5
21. 2 30.4 j2.8 55. 8
0.0102
.0563
. 0921
. 1075
. 1533
0. 007
. 018
. 028
. 033
. 042
4, 580
l, 780
1, 150
970
760
78. 8
59. 5
59. 2
65. 5
65. 2
0. ()?.Al I . 0530
. 0675
. 0867
.1156 1
0.007
. 014
. 022
. 030
. 035
4,300
2, 150
1,370
1, 000
857
73. 8
71. 8
70. 5
68.5
73. 7
0. 0183
.0284
. 0502
. 0640
. 0960
0.006
~01 2
. 021
.028
. 032
4,970
2,480
1,420
1. 067
935
85. 5
83.0
73. 0
73. 2
80. 5
0.01 30 -
•• 00127640 1
. 0434
. 0564
---- ------------
-------- --------
--- ------------
-- --------------
------------- ---
----------------
--- - - - --- -- - - -
---- ------------
----------------
----------------
----------------
--------- -------
-- --------------
--------------- -
--- -------------
20
tJCfLECTloN Po1NT LOCATIONS
Totr.slON TE~ r
J'lt'li£RIEL /JIY/.5/01'>' o,f'!.O"h!JNG
I Bea.zns
" " " "' ~"' " " ~ " .. ~ ~ ~ ~
~ ~ ~ ' I
' I
ll) _, _, I:)... ""i to \()
~
, I
0) ' J - I \!_
_,
~ CJ) ~ ~ 'O
llweh
~ {")
I<) ~ "-i
~
l'-..
"-!
,'<'<:) ... ...
F IGURE 19
XI. STRESS-DISTRIBUTION STUDIES
In order to determine the distribution of the stresses
due to varying conditions of loading a McCollum-Peters
electrical strain-gage, or telemeter, outfit was employed.
The strain gages proper consist of two carbon stacks,
the pressure on which is varied by the elongation in the
gage length of the material to which it is attached,
when that metal is stressed. The gage length of these
instruments was 8 inches.
These two carbon piles form two resistances of a
Wheatstone-bridge circuit, and any change in balance
of the circuit caused
by a change in pres­sure
on the carbon
piles is measured by
a millivoltmeter.
Two carbon piles
are used in order that
the resistance of the
FIGURE 20.-Electrical circuit, McCol· strain gage shall vary
lum·Peters telemeter apparatus approximately line­arly
with elongation.
The strain gages are attached by long wires to a panel
board on which is mounted a voltmeter for maintaining
a constant voltage across the entire circuit, the milli­voltmeter
already mentioned, a series of s"itches for
connecting individual strain-gage elements to the milli­voltmeter,
and resistances for balancing each circuit.
This panel boai:d accommodated the six available strain
gages.
The change in balance of the system per 0.001 inch
elongation was such that the millivolt reading was ap-proximately
1 millivolt, or by changing a resistance
plug could be made to read 1 millivolt per 0.0005 inch,
approximately.
A calibration curve for each strain gage was drawn
after the tests were completed, on which millivolt read­ings
were plotted against known elongations. These
curves were later used in converting elongations to
stresses.
The expression of the elongation of any material be­low
its elastic limit is
where
t:,.=PL
AE
.6.=elongation in length L=S inches.
P=applied load over section of area A.
E=modulus of elasticity, taken at 10,000,000
pounds per square inch for duralumin.
From the above expression
p E
A=L A= 1,250,000D.
where D. is in inches, for an 8-inch gage length.
The graduations on the millivoltmeter were such
that it was necessary to estimate anything under Yz
millivolt. Tenths could be picked off readily, but
anything under a tenth is liable to error.
The strain gages are subject to hysteresis, and will
not give accurate information for other than an
increasing load.
Upon removal of the load, in practically every test
it was found that the strain-gage circuits were unbal­anced
from 0.1 to 0.5 millivolt.
21
The strain-gage records are not considered to be
without error. Due to hysteresis, possible slippage
(small) and other errors, the results may be from 10 to
20 per cent in error, and in cases where records were
made on thin metal and buckling occurred, even more.
The computed stresses being dependent upon E, are
subject to errors due to its variation.
In the beam tests, where the load was applied by
means of a jack and platform scales, it was found to be
very difficult to obtain consistent readings by repeated
applications of load.
The stress limit of the strain gages was about 10,000
pounds per square inch, corresponding roughly to 8
millivolts with an 8-inch gage length.
Strain-gage readings were taken at as high as 24
points around the periphery of the wing during some
of the tests. In order to do this, it was
necessary to locate the six available gages
four times around the periphery of the
wing.
The strain gages were located around
the leading edge at station 57~ and around
the remainder of the wing at section 4}'2.
When the gages were installed, it was
found that two nose bulkheads, 5 and
5%, were just 8 inches apart, and hence
provided a convenient anchorage for the
gages. The remainder of the gages would
have been anchored into a splice had they
been placed at section 5~; so they were
moved inboard slightly to clear this splice.
At station 147'2 the gages were all located
in the same section throughout the periph-ery
of the wing. The installation of the
strain gages is shown in Figures 60, 61,
62, and 63.
Jn the trailing-edge portion of station
4% the strain gages were attached to
%6-inch round-head bolts inserted in the
covering. At all other points the strain
gages were attached by tapping into the
metal, regardless of the thickness.
The strain gages located in the leading
and trailing edges of the wing often gave
erroneous indications of stress due to the
formation of buckles between the gage
points.
Three types of loading were stu9ied in
the stress-distribution tests. These were
beam loads, applied approximately at the elastic axis
of the tip section, chord loads, forward and back,
applied at station 17, and a torque load, also applied
at station 17.
The elastic axis was found by applying a concentrated
load at points along the wing chord until a point was
found that would give equal deflections of the leading
and trailing edges at that section. Having found the
elastic-axis location, upbeam loads were applied by
means of a jack and platform scales, as shown in Figure
57, and downbeam loads by means of pig lead symmet­rically
placed about the elastic axis, as shown in Figure
56. During these tests the tip not being loaded was sup­ported
by a trestle and sand piled upon it to act as a
86802-32--4
counterbalance. Pig lead in sufficient quantities to
weight the wing down was piled on the I beams to
resist up loads.
Figures 58 and 59 show the method of applying chord
loads at station 17 and the bracing necessary to resist
them. Both wing tips were equally and simultaneously
loaded in order that there would be no twist of the
wing about its supporting fittings.
Torsional loads were applied as a couple, there being
no accompanying beam loads. A block was cut to the
contour of the airfoil at station 17 and padded with
felt to prevent damage to the wing. Two 4 by 4 inch
beams were clamped to this block, the one on the lower
side extending toward the trailing edge and the upper
one extending toward the leading edge. A platform
was hung from the leading-edge beam, and at the trail-
.MllT@EL /J!V!SICN ..f'S'-0" W/1'16
.5rRAIN 6ME LOCATION
Secli'orz 14j
.S"<zcfz'o.n 4 j
FIGURE 21
ing edge a platform was suspended from a pulley so
that it pulled up just as much as the other pulled down.
Figure 56 shows the set-up for the torsion tests;
BEAM-LOAD TEST
Figures 24 and 25 reveal the nature of the stresses
indicated by the strain gages. The extreme variance
of the stresses in the leading and trailing edges can be
readily attributed to the buckling of the thin metal of
these sections under load.
The lower corrugated flange sheets were overlapped
one corrugation at the center of the box beam, which
would indicate the possibility of a slight drop in stress
at that point. It does not account, however, for the
22
scattered points on the lower-flange stress curve. In
fact, there is a drop in stress at point 3 of the upper
flange, but the upper flange is continuous from web to
web, there being no longitudinal splice at this section.
A part of the variance of the lower-flange stresses
may be due to the installation of the strain gages.
These were secured to the corrugated material by bolts
screwed into tapped holes in the corrugations. The
thickness of the lower flange was but 0.020 inch, with
a covering of 0.014-inch flat sheet. Just how securely
the gages were attached to this thin material is ques­tionable.
Whether or not the 0.014-inch sheet and the
0.020-inch sheet always acted together is also question­able,
for the flat sheet was secured to the corrugated
sheet by rivets spaced about 2J/i inches. It is quite
possible that the flat sheet could have been buckled
slightly between rivets so that at low loads it might
not be carrying its full share of load.
By plotting the stresses at the web flange angles on
the sketch of the wing, using the webs as zero stress
that is almost identical with that found graphically.
The stresses computed by the above expression are
shown on Figures 24 and 25 for 1,000 and 2,000 pound
loads. It is seen that they are in excess of the measured
stresses but are essentially correct in shape. To be
entirely consistent, the lower skin should be neglected
in computing the section properties for the case shown
in Figure 25 when the lower skin is in compression.
This would result in a shift of the neutral axis of about
an inch and some change in moment of inertia. How­ever,
the agreement between the computed and meas­ured
stre ses is even better for this case than the other
loading, where the upper skin was neglected.
It is probable that the use of the moment at station
4}'2 in computing stresses is slightly severe, for the webs
are trusses, and the nearest panel is at station 5, about
7 or 8 inches outboard of the centers of the strain
gages. For a 2,000-pound load at the tip the moment
at station 5 is about 440,000 inch-pounds, 12,000 inch­pounds
less than at station 4J/i. This difference is
Mtltfr&Rt£L c>1n.310N ss~o" W11VG
E.t..4.STIC A:x!V LOCATION
I I I -I .,1- +- + I );",.l,..,.nt' tw@ t -+ -t-~I I
OZ' I I I
1
FIGURE 22
lines, the neutral aids was located by joining corre­sponding
flange-angle stress points. This was done for
both up and down loads at the tip. The stress lines
for the upload are quite consistent in crossing the webs
at about 13 inches above the lower chord. For the
download the higher stresses indicate the same neutral
axis, though the first load does not.
The most consistent agreement between measured
and computed stresses was obtained by computing the
stresses according to the beam theory, using the flexure
formula
The'moment was the bending moment at the section
where the strain gages were located. The moment of
inertia I , the distance to the outermost fiber c, and the
location of the neutral axis was obtained by considering
only the box portion of the wing and by neglecting the
flat skin on the upper surface. This gave a neutral-aids
location of 12.9 inches above the lower chord, a location
but 2J/i per cent, which is small in comparison to either
the variation in stresses or the differences between
computed and measured values.
By locating the neutral ax is as on Figures 24 and 25
it appears to be 13 inches from the lower flange.
Having the neutral axis location, the moments, and the
stresses, it is possible to solve for I.
I
_MXc
- p
A
Solving this expression, using M =452,000 for 2,000
pounds upload,
At rear upper flange angle _______ J = 1,200 in.4
At front upper flange angle ______ J = 1,300 in.4
Center, between webs __ ________ J = l,150 in.4
Lower flange (mean) __ ____ ___ __ J = 1,175 in.4
Mean ________ _____ ____ __ J = l,180 in.4
The value of I computed for the box portion only
was 1,020 inches4• (See p. 23.) The difference
-
23
ZZ.'30
2:J.Z8 zaza
t ~
FIGURE 23.-Section 4 ~-Materiel Division 55-foot wing.
between these values is probably due to the material
of the leading edge, for that part of the leading edge
that is acting in tension gave a good stress record for
the download considered above. The contribution
of the trailing edge seems to be negligible.
TABLE 17
SECTION PROPERTIES, SECTION 4)1
Description Area Lever Moment
------------- --- - - - ---- ----
Upper corrugations ___ ____ -- -- - ---- ---_. __
Lower corrugations ___ ________ __ ___ ----- - -
FRreoanr tu uppppeer rf lfalanngge e_ _____ --_-_-_-__-_--_-_-__-_--__--_-_-__-_--_-_
Front and rear lower flange ___ __________ _
Lower surface sheet_ ____ __ ___ _______ _____ _
Front and rear lower web portions __ _____ _
Front upper web portions ____ ______ __ ___ _
Rear upper web portions __ --------- ---- - -
2. 590
1.138
. 287
. 287
. 574
. 630
. 229
. 128
. 090
TotaL _______ ----------- -- - - - - - -- -- 5. 953
22. 300
. 375
20. 285
23. 285
. 715
.007
!. 576
22. 424
19. 424
L oca t1. 0n o Cn eutraI lW.S 756..995783 -12.93 above lower surface.
MOMENT OF INERTIA
Description Area c•
57. 900
. 428
6. 570
6. 680
. 410
.003
. 360
2.880
!. 747
76. 978
leg
--------------1--~ --- ------ -
Upper corrugations______ ___ __ ___ __ __ 2. 590 129. O 334. O
Lower corrugations_______ ___ ____ ____ 2. 570 158. 2 407. O
Rear upper flange__ _______________ __ . 287 42. 3 12. 2
Frontupperflange_ ___ __________ ___ .287 96.8 27. 8
Front lower flange____ ___ ____ ___ __ __ . 287 148. 5 42. 7
Rear lower flange______ ______ _______ . 287 148. 5 42. 7 -- -- -- -
Lower surface sheet___ ___ ____ _______ .630 167.0 105.3 _____ _ _
Front and rear lower web portions__ . 229 118. 8 34. 2 0. 132
Front upper web portions ___ ___ ___ __ I· . 128 82. 2 10. 5 }
Rearupperwebportlons________ __ __ .090 36.3 3.3 · 113
. 'l'otaL ___ __________________ __ =-:-:-:-:-=-:-:-:-:- 1,019. 7 ~
. 245
I 1, 019. 945
Moment oC inertia= approximately 1,020 inches•.
STRESSES AT SECTION 4)1 D UE TO BEAM LO.A.DB
(Neglecting chord loads)
Moment
452, 000
inch­pounds
Lbs.fsq.in.
Stress at front upper flange angle ____ __ Y=ll. L_ 4, 930
Stress at rear upper flange angle ___ ___ Y= 8. L _ 3, 600
Stress at maximum ordinate ___ ___ ___ _____ 12. 3__ 5, 450
Stress at lower flange _____ __ _______ __ __ ___ 12. 9__ 5, 720
Bending moment, st ation 4)1:
5 factors, low inciden ce ___________________ in ch-pounds __
3 !actors, inverted fligh t_ __ ______________ _____ __ __ do ___ _
11 factors, high incidence __ ______________ ___ __ ____ do __ _ _
I of box, section 4)1 ____ ____ ___ ________ ____ ____ ___ inches• __ _
Y , front upper flange angle _________ ______ ____ ____ inches _ Y, rear upper flange angle ___ ___ __ ___ ______ ________ __ do __ Y, maximum ordinate ___ ____ __ ___ __ __ ___________ ____ do ___ _
Lower flange---- ---- ----- ------ - -- ------ ------------do ___ _
I NVERTED F U GHT
M/I-280,000_
. 1,020 3 factors.
Stresses:
Front upper flange angle ___ ___ pounds per square inch __
RMeaaxri umpupmer o frladningaet ae n__g_l_e__ _-_-_-_-__-_- -__-_--__--_-__- -__--__- -_-__-_-_- -_dd oo ______ __
Lower flange _________ _____ ___ - -- - - - -- - -- - - __ _____ do __ _ M/I-470.000_
. 1,020
Stresses:
Low INCIDENCE
5 factors.
Front upper flange angle ____ __ pounds per square inch __
Rear upper flange angle ______ ______ __ ____ _____ __ do ___ Maximum ordinate _____ ____ ___ __ __ __ ___ _____ ___ do ___ _
Lower flange ___ __ __ _____ __ __ ____ ___ ______ ___ __ ___ do ___ Stresses:
HIGH I NCIDENCE
M/J=l,800,0CO
1,020
Front upper flange angle _________ _______ _ _
Rear upper flange angle __________ ________ _
Maximum ordinate ______________ __ ___ ___ _
Lower flange _________ _____ ___ ____ ___ ___ __ _
11 fac tors
Lbs.fsq. in.
19, 600
14, 300
21, 700
22, 800
226, 000
inch­pounds
Lbs.fsq. in.
2, 465
1,800
2, 725
2, 860
470, 000
280, 000
1, 800, 000
1, 020
11. l
8.1
12. 3
12. 9
3, 050
2, 220
3, 380
3, 540
5, 120
3, 730
5,660
5, 950
4 factors
Lbs./sq. in .
4, 080
2, 980
4, 525
4, 750
'
Gage location ____ ______
Gage No _____________ _
Load
500- --- - - ----- -- -- {
MV
ll
P/A
1,000 __ ____ __ __ ___ { MV
ll
P/A
1,500- - - ·- - - - -- ---{
MV
ll
P/A
2,000 _______ ______ { MV
ll
P/A
Gage location ____ ____ __
Gage No _________ _____
Load I MV .. __ l ll
P/A
1,000 _________ ____ { MV
ll
P/A
1,500 ____ __ ___ ____ { MV
ll
P/A
2,000 _____ ___ ___ __ { MV
ll
P/A
Gage location __________
Gage No ____ ___ ____ ___
Load
500 __ ___ ___ _______ { MV
ll
P/A
1,000- -- - - -- -- - --- {
MV
ll
P/A
1,500 ____ ____ ____ _ { MV
ll
P/A
Gage location ____ __ ____
Gage No ___ ____ ____ __ __
Load
500 __ ______ ___ ____ { MV
ll
P/A
1,000 ___ ___ _______ { MV
ll
P/A
l,500 _____ ___ _____ { MV
ll
P/A
--
l
·--
93
-0. 90
- . 0008
-1,000
-1.80
-. 0017
-2, 188
-2. 70
- . 0024
-3, 063
-3. 50
- . 0032
-4, 000
13
93
- - -
1. 00
. 001
l , 250
2. 00
. 0020
2, 500
3.10
. 0030
3, 750
4.10
.0038
24
TABLE 18.- Stress-distribiition study
BENDING 'l'ES'l'
UPLOAD ON Tll'
2 3 4 5 6 7
------ ------------
90 88 87 86 85 93
- - - ---------------
-1.00 -1.00 -0.85 -0.85 -0. 55 +o. 10
-.0008 -.0008 - . 0009 -. 0007 -.0005 .0007
-1,063 -1, 000 -1, 125 -938 -688 875
-2.20 -2.00 -1.70 -1. 75 -1.10 0.65
-.0018 -.0016 -.0017 -.0015 -.0011 .0006
-2, 250 -2, 000 - 2, 188 -1, 938 -1, 375 813
-3.40 -3.10 -2. 70 -2.80 -1. 75 0.60
-.0027 -.0024 - .0028 -.0024 -.0017 .0006
-3, 438 -3,000 - 3,500 - 3, 000 -2, 188 750
-4.50 -4.30 -3. 70 -3.90 -2.45 0.50
-.0036 -. 0034 -. 0037 - . 0033 -.0024 .0005
-4,500 -4,250 -4,625 -4, 125 - 3, 063 625
14 15 I 16 17 18 19
90
-8-S - ,- 8-7 -
86 85 93
---- --------------
1. 20 1. 00 1. 00 1.30 1. 00 3. 90
. 0011 .0009 .0011 .0012 .0010 . 0038
1, 375 1, 125 1,375 l, 538 1,313 4, 750
2.30 1. 90 1. 90 2.50 2. 00 5. 25
. 0021 . 0018 . 0020 . 0024 .0021 . 0048
2, 688 2,250 2,500 3,000 2,625 6,060
3. 40 2. 90 2. 70 3. 50 3.00 6.40
. 0031 . 0027 . 0031 . 0034 .0031 .0058
3,940 3, 375 3,875 4, 250 3, 938 7,250
4. 50 3. 90 3.40 4.50 4.00 7.40
. 0041 .0037 .0036 .0044 .0041 .0066
4,82s I 5, 125 4, 625 4, 500 5,500 5, 125 8,315
DOWN LOAD ON TIP
l 2 3 4 5 6 7
-- ---------------
93 90 88 87 86 85 93
---------------- --
.85 1.10 1.00 0.85 0. 80 o. 55 0.125
.0008 .0010 .0009 .0009 .0007 .0006 .0001
1, 063 1, 250 1, 125 I, 125 938 750 163
1. 70 2.20 2. 00 1.65 1.65 1.125 o. 25
. 0016 .0020 . 0018 .0018 .0015 .0012 .0002
2,063 2,560 2, 250 2, 250 I, 938 I, 500 313
2. 60 3.375 3. 00 2. 45 2. 45 1. 70 0. 20
. 0025 . 0031 .0028 . 0026 .0023 .0018 .0002
3, 125 3, 875 3, 500 3, 250 2,875 2, 250 250
13 14 15 16 17 18 19
93 90 88 87 86 85 93
- - - - ------------ - -
-0. 575 -0.60 -0.60 - 0. 75 -0.625 - 0.50 - 2.50
-.0005 -.0005 -. 0004 -.0008 -.0005 -. 0005 -.0023
-688 -625 -563 -1,000 -725 -625 -2,875
-1.90 -2.35 -2. 05 - 2.30 - 2. 10 -1. 75 -4.50
- .0018 -. 0020 -.0016 - .0023 - . 0018 -.0017 -.0040
-2, 250 -2, 500 -2, 000 -2,875 -2, 250 -2, 185 -5,000
-3. 20 -4. 10 -3. 80 -3. 80 -3.65 -2.85 -5.80
-.0029 -.0032 -. 0030 - . 0038 -. 0031 - . 0028 -. 0052
-3, 625 -4, 060 -3, 750 -4, 750 -3,875 -3,500 -6,500
All signs plus unless marked otherwise.
8 9 10 11 12
--- ------------
90 88 87 86 85
---- - - - --------
-0.15 0.125 0 0 -0.15
-. 0001 .0001 0 0 -. 0001
-150 150 0 0 -188
-0.125 o. 125 0 0 -0.125
- .0001 . 0001 0 0 - . 0001
-125 150 0 0 -163
-0.25 o. 625 0 -0. 25 0.45
-.0002 .0006 0 - . 0002 . 0004
-250 750 0 -250 563
-0.425 0. 750 -0. 05 0 0.85
-.0003 .0007 - . 00005 0 . 0009
-438 875 -63 0 l, 125 1
20 21 22 23
- ----- _ 24 1
90 88 87 86 85
--- ---------- --
0. 525 1. 00 0. 50 -0.175 -0.475
.0004 .0009 . 0005 -.0001 - . 0004
563 1, 125 625 -188 - 563
l. 50 1.90 0. 95 -0. 35 - 0. 95
.0014 .0018 . 0010 -.0003 - . 0009
1, 750 2,250 1, 250 - 375 -1, 188
2.60 2.90 1. 30 -0.60 -1.40
.0024 .0027 . 0014 - . 0005 -.0014
3, 000 3,375 I, 750 -688 -17, 750
3. 75 3. 70 1. 70 -0. 95 -1.40
.0034 . 0035 . 0018 - . 0008 -.0014
4,310 4, 375 2, 250 -1,063 -1, 750
8 9 JO 11 12
---------------
90 88 87 86 85
- - - ------ ------
- 0.05 0 0 0 - 0. 45
-.00005 0 0 0 -.0004
-63 0 0 0 -563
-0.125 -0.10 0 0.025 -0. 70
-. 0001 -.00008 0 • 00002 -. 0007
- 125 -100 0 31 -875
-0.25 -0.40 -0.20 0 -0.85
-.0002 - . 0003 - . 0002 0 -.0008
-250 -375 -250 0 -1,063
20 21 I 22 I 23 24
---------
90 88 87 86 85
- - - -- ---;---- --
- 1.00 -1. 05 -0.55 0. 10 0. 60
-.0008 -.0008 -.0006 . 0001 .0006
-1,063 -1, 063 -750 125 750
-5.60 -2. 15 - 1.20 0.20 1.25
- . 0044 -.0017 - . 0012 .0002 . 0013
-5,500 -2, 125 -1, 500 250 1, 625
-7.25 -4.00 -1.90 0. 30 l. 90
-. 0057 - . 0032 -.0019 . 0003 .0020
-7, 180 - 4,000 -2. 375 I 375 2,500
25
CHORD-LOAD STRESSES
In resisting stresses due to chord loads, the flanges of
the box beam become webs, but the web flange angles
only act as flanges, since the true webs are trusses. The
leading edge would be expected to resist compressive
loads due to a forward-acting chord load, but due to its
flat, unsupported nature its resistance to stress becomes
small and variable.
Figures 26 and 27 show the measured stresses at
Station 4 ~ for chord loads applied at Station 17 and
acting both forward and rearward. That these stresses
are determined by beam action is immediately apparent
from the straight-line variation of stress across the box
FIGURE 24
section. Buckling of the leading edge is also quite
evident from t hese curves.
In determining the stresses due to chord loads no
measurements were made in the trailing-edge section
due to the previously obtained erratic results in this
region.
It may be noted from Figure 26 that the zero stress
line (that is, the line forming the points where the stress
curves pass through zero) is not perpendicular to the
chord of the wing, and further, that this line changes
position with a change in direction of load. The change
in position may be expected, for when the chord load is
acting rearward the trailing edge is in compression and
is ineffect ive. When the chord load acts forward, the
trailing edge is stressed in tension and is effective,
tending to draw the neutral axis rearward.
This slanting apparent neutral axis makes it ex­tremely
difficult to obtain agreement between measured
and computed stresses, unless each of the corrugated
flanges is treated separately. Since the stresses due to
chord loads are so much smaller than those due to
beam loads, this refinement of analysis is hardly
warranted.
When the chord load was acting forward, stressing the
trailing edge in tension, the best agreement between
computed and measured stresses was obtained by con­sidering
the box section, the leading edge, and one-half
FIGUltE 25
of the trailing edge as effective. This results in a con­servative
indication of stress for the lower corrugated
flange, but the computed stresses for the upper flange
are lower than the measured stress, though in closer
agreement. The discrepancies at the flange angles are
about 20 per cent for the lower, 30 per cent conserva­tive,
and 15 per cent dangerous for the upper flange.
The moment at St.ation 1 due to a 2,000-pound chord
load was 560,000 inch-pounds. The chord-load moment
at Station 1 in high incidence at 15 factors was but
374,800. The shears wouldjbe:2,ooo pounds and about
3,000 pounds, respectively, so the chord-load test
stressed the wing more severely in bending than the
high-incidence chord loads at failure.
Gage location_- - --- --- --- -- ---- 1 2 3 4
Gage No ____ ___________________ 93 90 88 87
--- - ----
Load { MV - 0. 60 - 0.30 -0.10 0.125
500 __ -- -------- -- - - ---- -- t. -. 0006 -. 0002 -. 0001 . 0001
P/A -750 -313 -125 156
MV -1.15 - 0. 625 -0. 25 o. 225 """-------· ---------·--1 t. -. 0011 - . 0005 -. 0002 . 0002
P/A -1, 375 -688 -250 282
MV -L40 -0.80 -0.30 o. 25
1, 250_ - ---------------- -- t. -. 0013 -. 0006 -. 0002 . 0002
P/A -1, 625 -813 -313 313
1, 500_ - ------------------ {
MV -1. 60 - LOO -0. 375 0. 25
t. -. 0015 - . 0008 -. 0003 . 0002
P/A -1,875 -1. 000 -375 313
- --
I 250-- ---~~~~----------- -{ MV o. 25 0.10 0 -0. IO
t. .0002 . 0001 0 -. 0001
P/A 313 125 0 -125
500---------------------- {
MV 0. 45 0. 20 0 -0. 25
t. . 0004 . 0002 0 -. 0002
I P/A 594 250 0 -313
750- --------------------- {
MV 0. 65 o. 275 0 -0. 375
t. . 0006 . 0002 0 -. 0004
P/A 813 313 0 -500
1, 000-------------------- {
MV 0.90 0. 375 0 -0. 50
t. .0009 . 0003 0 - . 0005
P/A 1, 125 438 0 -625
l, 250-- ----- ---- - ------- - {
MV 1.125 o. 50 0 -0. 70
t. . 0011 .0004 0 - . 0007
P/A 1, 375 594 0 -875
All signs plus unless marked otherwise.
TABLE 19.-Stress-distribution stud),
OIIORD-LOAD TEST
CHORD-LOAD FORWARD AT STATION 17
5 6 13 14 15
_ 1_6_,_11_ ------
86 85 93 90 88 87 86
---------------------
o. 30 0.50 0. 35 0. 30 0. 10 -0.05 -0. 20
.0003 . 0005 . 0003 .0003 . 0001 -. 00005 .0002
375 625 438 375 125 -63· -250
0. 55 LOO o. 65 o. 50 0.175 -0.20 -0.40
.0005 . 0010 . 0006 . 0005 . 0001 - . 0002 - . 0004
625 1, 250 813 625 213 -250 -500
0. 65 1. 20 o. 80 o. 575 o. 20 -0.25 -0. 555
.0006 . 0012 . 0008 .0005 .0002 -. 0002 -. 0005
750 1, 563 1, 000 688 250 -313 -625
o. 75 !. 375 1. 00 0. 70 o. 25 -0.30 -0. 70
. 0007 . 0014 . 001 . 0006 . 0002 -. 0003 -. 0006
75 1, 750 1, 250 813 313 -375 -750
C f!OUD-LOAD BACK AT STATION 17
-0. 275 - 0.35 -0. 20 -0.15 -0. 05 0. 05 0. 175
-. 0002 -. 0003 -. 0002 -. 0001 - . 00002 . 00005 . 0001
-313 -438 -250 -157 -32 63 219
-0.50 -0. 675 -0. 45 -0. 325 -0.15 0.10 0. 30
-. 0004 -. 0007 -.0004 -. 0002 -. 0001 . 0001 . 0002
-563 -875 - 500 -344 -144 125 344
-0. 75 -0. 975 -0. 725 -0.50 - 0. 25 o. 15 0.45
-.0006 -.0009 -. 0006 -.0004 -. 0002 . 0001 . 0004
-844 -1, 219 -844 -532 -250 188 532
-1. 025 -1.30 - 1.00 -0. 75 -0. 35 0. 20 o. 525
-.0009 -. 0013 - . 0009 -.0006 -. 0002 .0002 . 0005
-l, 125 -1, 625 -l, 157 -782 -357 250 625
-L 30 -1. 575 -1.35 -1.00 - 0. 50 o. 25 0. 65
-. 0011 -. 0015 -. 0012 -. 0008 - . 0004 .0002 . 0006
-1, 375 -1, 969 -1, 532 -1, 063 -500 313 769
- -
18 19 20 21 22 23 I 24
___ I ___
85 93 90 88 87 86 85
--- -------------------
-0.30 -1.30 -0. 75 -0.85 - 0. 75 -0. 70 -0.50
. 0003 -. 0012 -.0006 -. 0006 -. 0008 -. 0006 - . 0005
-375 -1,500 -750 - 813 -1, 000 -750 -625
-0.65 -2. 25 -1.875 -L625 -1.60 -1.30 - LOO
-. 0006 -. 0021 -. 0016 -. 0013 -. 0016 - . 1100 - . 0010
- 813 -2, 625 -2, 000 -1,625 -2, 000 -1, 438 -1, 250
-0.875 -2. 45 - 3. 00 -2. 05 -2.00 -L 50 -L 10
- . 0008 -. 0022 - . 0024 - . 0016 -.0020 - . 0013 -. 0011
-1, 093 -2, 815 -3,000 -2, 000 -2, 500 -1, 625 -1,375
-1.05 -2. 75 -6.00 -2. 75 -2. 325 -1.675 -1.15
- . 0010 - . 0025 -.0047 - . 0022 -. 0023 - . 0014 -. 0011
-l,313 -3, 125 -5, 875 -2, 750 -2, 940 -1,813 -1, 438
_I_ -
- -
0.175 o. 50 LOO 0. 50 0. 45 o. 375 o. 25
. 0001 . 0005 .0009 .0004 . 0005 . 0003 . 0002
219 625 1, 157 563 625 438 344
o. 325 o. 85 !. 50 !. 00 o. 90 o. 75 0. 50
. 0003 .0008 . 0014 . 0009 . 0009 . 0007 .0005
407 1, 063 1, 750 1, 157 l, 219 75 657
0. 475 L 125 2.00 1. 50 !. 50 1.15 0.90
.0005 . 0011 .0018 . 0014 . 0016 . 0011 . 0009
625 1, 375 2, 313 1, 750 2,000 1, 375 1, 188
0. 60 1. 45 2. 40 2. 00 2.00 1. 50 1. 20
.0006 .0014 . 0022 .0019 . 0021 . 0014 .0012
782 1, 750 2, 750 2, 375 2, 657 1, 813 1, 594
o. 75 !. 75 2. 80 2. 45 2. 50 1. 95 1. 55
.0007 . 0016 . 0025 . 0023 . 0027 . 0019 . 0016
875 2,063 3, 188 2, 907 3, 375 2, 31s I 2,000
27
FIGURE 26
FIGURE 27
FIGURE 28
COMPUTATION OF STRESSES DUE TO CHORD
LOADS
The action of the wing in resisting chord loads will
most likely be that of a cantilever beam, with the cor­rugated
flanges acting as web members. The leading
and trailing edges will probably be effective in resisting
tension loads, but their resistance to compressive loads
is doubtful.
COMPUTATION OF SECTION PROPERTIES AT
STATION 4}1,
The diagonal \\"eb members will be neglected in com­puting
the moment of inertia of the section.
Since the 0.014-inch covering over the box portion of
the wing is situated close to the neutral axis, it will
not be highly stressed and will be considered as being
effective in resisting chord loads.
FIGURE 2Q.-Chord-load neutral axes
28
TABLE 20.- Determination of neutral-axis location and section properties-Station 47'2
Distance Moment
ofC. G. of area
Item Area from about h I Ah' Io Total
leading leading
I edge edge
-------------·~------------- 1---~1 ---------
Leading edge (periphery=56 inches, t=0.020 inches) ___ ___ ___ _______ _
Front flange angles __ __ __ --- -- ------- __________ ___ ________________ __ _
Sq.in. I l. 120
. 574
.260
10.0 11.2
22. 7 13. 0
36. 3 1,470 ---------- 1,470
23. 5 318 318
Front web ______ --- -- _______ ------ ___ ------ _______ _______ __________ _
LUopwpeerr ccoorrrruuggaatteedd ffllaannggee aanndd ccoovveer'_- -__-_-_--_-_-__-_-_-_-_- -__- -_-_-__-_-_-_-_-_-_-_-_-_-_-_-_-_-_- -_
Rear flange angles-- ----- --- ----- ------- ------ - - ---- -----------------
TRreaairl iwnge-be .d•g -e- c-o--v-e-r- .-.-.- _____ .- -__-_--_-_-__-_-_-_- -_-____--__--__-_-_-_--__-_- -------_-__- -. -_-__- -. -_-. _____. _-_- __
Trailing-edge strip._. _______ _______ ___ . __________ . __ . ______ . ___ . ___ _
Total, box only _________ ___ --- - - __ --- ---- --- ---- - -- --------------- __
Total, entire section __ ____________ ----------------------------------
Total, entire section less 712 trailing edge ____________________________ _
Total, entire section less leading edge ____ ___________________________ _
Total, entire section less trailing edge ______ _________________ _____ ___ _
3.210
l. 790
.574
.184
1.494
. 056
6. 590
8.820
8. 050
7. 700
7. 270
23.0
45.5
45. 5
68.2
68.0
94.0
119. 6
46. 3
51.8
48.0
57.8
43. 5
6.0 23. 3
146. 0 .8
81.5 .8
39.2 22.9
12. 2 22. 7
140. 5 48. 7
6.8 74. 3
304. 7 ----------
456.4 ----------
386. 2 ----------
445. 2 --- -------
316. 0 ----------
----------
141 ------538- 141
538
341 341
300 300
100 ---------- 100
3, 540 ---------- 3,540
309 ---------- 309
---------- ---------- 1, 738
-- -------- ---- --- --- 7,057
---------- ---------- 5, 132
---------- ----- ----- 5,587
----- ----- ----- -- --- 3, 208
t h=distance of elements from line through centroid box section only.
Moment at station 4~ due to 1,258-pound chord load=282,000 inch-pounds.
Front flange angle stress=282,000X~3~-l,350 pounds per square inch.
Rear flange angle stress=282,ooox~3;-1,320 pounds per square inch.
STRESSES DUE TO TORQUE LOADS
The stress-distribution curves obtained at stations
4~ and 14% yield little quantitative information due to
their erratic nature.
The records of station 14% indicate the presence of a
bending moment in the plane of the corrugated flanges
due to shear caused by the applied torque load. The
shear forces act in the upper flange to cause compres­sive
stresses in t he vicinity of the front web and ten­sile
stresses at the rear web. The action in the lower
flange is reversed.
These records shed no light on the existing shear
stresses in the web members. It is regretted that a
study of the distribution of shear between the web
members was not made, for this information would
have been valuable in determining the correctness of
using the center-of-pressure location in distributing
shears between the webs, rather than the elastic-center
location.
Due to the nature of the stress-distribution curves
obtained during the beam tests, which indicate that
the box section is acting as a whole in resisting bending
moments, and the rigid nature of the wing which tends
to equalize the loads carried by the spars, the assump­tion
of a division of shear determined by the center-of­pressure
location appears questionable.
From the low nature of the stresses shown at sta­tion
4}~ due to torque loads, it would seem that direct
stresses due to torque are not great. This can not
always be said of shear stresses, however, of which
data are lacking.
The shear stresses due to torque at station 4~
a.re calculated for H. A. A. in the following para-graphs,
using the general expression for any closed
section.
S=shear stress, pounds per square inch.
A=inclosed area, square inches.
Q=acting torque at the section considered.
t=thickncss of shell, inch.
At section 4).'2 :
Area of box portion, 45 X 24Yz
Area of leading edge, 18X 23
Square inches
= 1, 104 approx.
414 approx.
1, 518 approx.
Area of trailing edge, 52X 23/2= 600
2, 118 total.
Total shear outboard of station 4).'2 (reference H. A.
A. load schedule and computation of loading) =
10X 940+697= 10,097 pounds.
The torque at station 4Yz will be assumed due to all the
shear acting along the line of action of the sand load, with
a lever arm equal to the distance to the elastic center.
Lever arm, station 4Yz, high incidence= 12
inches. (See fig. 22.)
Torque= 10,097X 12= 121,100 inch-pounds.
Q 121,100
S=2At=2x2,118x co.045 + 0.014)
485 pounds per square inch shear.
0.059
S=485X0.034=
840 pounds per square inch lower chord.
S = 1,430 pounds per square inch in leading edge.
Thus it is seen that the stresses in the corrugated
flange is low, but that the leading edge is fairly highly
stressed in comparison.
Gage location __ 5 I 13
T ABLE 21.-Stress distribution study-Torsion test
STATION 14)6
14 15 16 17 18 8 10 11 12 19 20 21 22 23 24
g; 1------·I--------------------------------------------------------------
00
0 I I
Gage No ______ _ 93 90 88 87 86 85 93 90 88 87 86 85 93 90 88 87 86 85 93 90 88 87 86 85
l----.---·1--------------------------------------------- --- --- ---------------
Load
75
150
.MY
PIA
.MV
PIA
- 0.05 - 60 --03.2470 1
- 0. 15 - 0. 55
-190 - 690
- 0.10 0.05 0. 07
- 125 60 90
- 0.20 0. 10 0. 20
-250 125 250
0. 12 - 0. 15 -0.20 0. 05 0.02 o. 02
150 - 190 -250 -60 -------- ------
o. 25 -0.38 0. 15 -0. 05 0. 05 0.02
310 ---- ---- ---- --- - ---- ---- -------- ------
0. 12 0. 40 o. 05 0. 05 0 - 0. 10 -0.20 0. 27 0.15 0. 05 - 0. 03 - 0.10 - 0.35
150 500 60 60 0 - 125 -250 -340 190 60 -------- - 125 -440
0.20 0. 80 0. 10 0. 10 0 - 0. 25 - 0.48 0. 55 o. 27 o. 05 -0. 10 - 0. 40 - 0.80
-------- -------- ---- -- ------- ------- -- -- -- -- -------- --- ---- -- ---- - -- ----- -------- --------- -------
225 M V - 0. 30 - 0. 95 - 0. 40 O. 12 0. 30 0. 40 - 0. 60 - 0. 25 -0. 10 0. 10 O. 02 0. 22 1. 20 0. 15 0. 05 -0. 05 - 0. 37 - 0. 85 0. 90 0. 45 0. 10 - 0. 23 - 0. 60 -1. 30
PIA -------- -------- -------- -- - - - -- - --- - - --- -- - - ------ -- --- - - - -- -------- -------- - - - - -- -- - -- - -- ----- --- ------ - - -- - -- - -- -- - - -------- ---- ---- ------- ------- --- - --- -------- ----- ---- -- -----
300 MV -1. 00 -2. 62 -0. 65
PIA -l, 250 -3, 280 -810
375 { .MV - 1. 50 - 4. 35 - 1. 05
PIA - 1, 900 -5, 450 - 1, 300
Gage location __ 1 2 3
--· ---
Gage No _______ 93 90 88
--- ---
(Torque
In.-lbs.)
68, 775 { MV - 0. 45 - 0.45 0
PIA - 282 - 282 0
85, 970 { MV -5. 00 - 5.00 0
PIA -3, 125 -3, 125 0
All signs plus unless marked otherwise.
0.12 o. 45 o. 55
150 560 690
0. 10 o. 55 o. 70
125 690 870
4 5 6
------
87 86 85
------
- 0. 02 0 o. 25
-125 0 156. 5
- 0.15 0. 05 0. 30
-93. 8 31. 3 188
-1. 00 -0. 20 -0. 20 0.18 0. 10 o. 40
- 1, 250 -250 - 250 220 125 500
-1.20 - 0. 75 - 0. 15 - 0. 35 0.20 o. 53
- 1, 500 -940 -190 -400 250 650
S'rATION 4 ~
i 8 9 10 11 12
-----------------
93 90 88 87 86 85
--- --------------
- 2. 20 - 0. 20 -4. 45 -0.30 0.10 -3. 75
-l,37i -125 -2, 780 -188 62. 5 -2, 340
- 2.60 - 0.65 - 4. 20 - 0. 40 0 -4. 40
-1,823 -407 -2, 620 - 250 0 1-2, 760
1. 67 o. 20 0. 05 - 0.05 -. 48 - 1. 35 1. 20 o. 53 0. 05 - 0.35 - 0.90 -1.80
2, 100 250 60 -60 -600 -1, 700 1, 500 650 60 - 440 - 1, 120 -2, 250
2.10 o. 20 0 - 0. 20 - 0. 65 - 1.95 1. 55 o. 48 o. 05 - 0.50 - 1. 4.0 -2.30
2,600 250 0 -250 -810 -2, 440 1, 940 600 - 625 - 1, 750 - 2,880
13 14 15 16 17 18 19 20 21 22 23 24
--------------------------- --
93 90 88 87 86 85 93 90 88 87 86 85
---- - ---------------------- ---
0 0. 05 -0. 10 -0.10 -0. 15 o. 80 2.9 o. 97 0 - 0. 15 0 0. 20
0 313 -62.5 - 62.5 -93.8 500 1, 815 605 0 -93.8 0 125
- 0. 15 0 -0.15 J-o. 20 -0. 25 1. 20 2. 30 2. 50 - 0. 20 - 2.50 0 o. 25
-93. 8 0 -93.8 -125 -158. 2 750 1, 440 1, 565 -125 - l, 565 0 156. 5
30
FIGURE 30
SAND-LOAD TESTS
The stress-distribution curves for the sand-load tests
differ noticeably from those obtained in the beam-load
tests. Comparing the lower-flange stresses for the two
methods of testing reveals that the lower-flange stresses
in the sand-load tests have a decided maximum value
between the flanges, as may be seen from Figures 32,
33, and 34.
In the inverted-flight test the stresses seem to increase
directly with the applied load. In the low-incidence
test the stresses at the center of the beam increase
faster than the flange-angle stresses and do not seem to
go up as fast as the load. In the high-incidence test
the stresses at the center of the box increase in a direct
proportion to the load, but the stresses at the webs do
not, as may be seen by the crowding of the plotted
points at the webs in the upper flange in Figure 34.
The agreement of the computed and measured stresses
at low loads may be seen from the above-mentioned
figures. The plotted computed stresses are those
obtained by using the section properties of the box
alone. The computed upper-flange stresses are con­sistently
greater than the measured stresses, indicating
that to neglect entirely the upper covering of the box
was perhaps too severe at low stress values.
The computed stresses for the upper and lower flanges
at 11 factors, in high incisJence, are slightly more than
twice the values indicated by the strain gages for this
load factor, as may be seen by reference to page 23
and figure 34, page 32.
FIGURE 31
While the above-mentioned comp uted stresses do
not include those due to chord loads and torque, the
stresses due to these two loads are insufficient to account
for the enormous difference indicated . For instance,
the bending moment at station 4H, in high incidence at
11 factors, is 1,080,000 inch-pounds,
23
0.16X l,080,000X41913=810 pounds per square inch
approximately
due to chord loads by which the stresses at the front
upper and rear lower flange angles would be increased
and at the remaining corners decreased.
Reference to Figure 30 will show that direct stresses
due to torque, indicated by the strain gages, are small.
This increasing variance between the computed and
measured values of stress is probably due to the action
of the strain gages. Their limit would be but 10,000
pounds per square inch, but why they gave a stress
indication that was not proportional to the load is not
known. It is regretted that but six gages were avail­able,
for the use of twice that number would have fur­nished
a better check on the variation in stress across
the corrugated flanges .
The preceding study of the distribution of the stresses
in the wing, clue to varying types of loading, indicates
that the final check on the design of the structural
elements of the wing should be based on stresses indi­cated
by the common expression
F= Mc
I
31
where
F = stress-pounds per square inch at a point
c inches from the neutral axis of a section
having a moment of inertia of I inches 4
when acted upon by a bending moment of
M inch-pounds.
The study indicates that the value of moment of
inertia that gives a good agreement between measured
and calculated stresses due to beam loads, and is at the
same time conservative, is obtained by considering the
box portion only of the wing and by neglecting the
0.014 flat covering when it is in compr ession.
In calculating stresses due to chord loads the pro­cedure
is the same as above, except that the moment of
inertia is calculated about an axis perpendicular to the
chord, or to the chord loads, and that the leading edge
and one-haU of the trailing edge should be considered
to have obtained stresses in agreement with measured
stresses.
The stresses due to chord loads are small in compari­son
to those due to beam loads and need not be con­sidered
unless some section appears critical, due to the
beam loads.
The consideration of the low-incidence condition in
the design of the flanges does not.seem to be warranted,
for a rearward shift in the center-of-pressure location,
together with the lowered design factor of this condi­tion,
does not give stresses in excess of those in the
high-incidence condition at the high-incidence design
load in the vicinity of the rear web. High incidence and
inverted flight are the conditions that design the wing.
The limited number of strain gages available pre­vented
a study of the distribution of the shear between
the webs. This is regrettable, for it would probably
have shed light on both the web design and the stiff­ener
design, as well as determining whether or not the
low-inddence condition could be eliminated entirely
in the design of the wing.
T ABLE 23.-Strain-gage voltmeter readings- Inverted­jlight
test
Gage
loca-ti
on_ 13 16 18 1 3 6
------------ Gage _ 85 86 87 88 90 93
------ ------
Load
2.0 -3.00 4.40 3. 40 2.50 2.95 1.90
3.0 -5.20 -7.30 -5. 05 4.00 4. 70 2. 80
3.5 -6.30 -8. 90 -6.20 4.80 5. 50 3.40
ELONGATIONS
I
2. 0 i -0. 00150 I o. 0019
1
-0. 0017 I o. 00120 I o. 00140 I o. 00095
I
3.0 -. 00260 . 0031 -. 0025 . 00185 . 00220 . 00140
3. 5 -. 00315 . 0038 - . 0031 . 00220 . 00255 . 00170
STRESS
I
2.0
I
-1,875 1 -2. 380 I -2. 120 I 1,500 I 1 • .,50 I 1, 090
I
3.0 -3. 250 -3, 880 -3, 120 2,310 2, 750 l, 750
3. 5 -3, 940 -4, 750 -3, 880 2, 750 3, 190 2. 120
TABLE 24.-Low-incidence stress measurements-­Section
4%
Gatigoen _lo__c_a_-_ 1 4 6 18 16 13
--- ------------
Gage ___ ____ 85 86 87 88 90 93
--- ------------
Factor riv -0. 55 -0. 95 -0.60 + 1.60 +2.10 +1.25
2.0 . 00055 . 0008 . 00065 . 00125 . 00195 . 0011
P/A -687 -1, 000 -810 +1,560 +2,440 +1,375 r -1.1 -1.90 -l.25 +2.10 +2.95 +J. 75
3.0 .0011 .00165 . 0013 .0016 . 00270 .0015
MP/VA -1, 375 -2, 060 -1, 630 +2,000 +3,380 +1,875 -1.8 -2. 75 - 1.90 +2. 10 +a. 75 +2.00
4. 0 A .0018 .0023.s .0019 .0021 . 0034 . 0017
P/A -2, 250 -2, 940 -2, 380 +2, 620 +4, 250 +z, 120 rv -2.0 -3.10 -2.10 +3. 00 +4.10 +2.18
4. 5
ft/A
.0020 . 00265 . 0021 .0024 . 00365 . 00185
-2, 500 -3. 310 -2, 630 +a,ooo +4, 560 +2, 310 riv -2.25 -3.50 -2. 45 +3.10 +4.50 +2.45
5. 0 .00225 .0030 .0025 . 00245 .0040 .0020
P/A -2, 810 -3, 750 -3, 120 +3, 060 +5,000 +2, 500 {MV - 2.50 -4.05 -2.33 +3.40 +4.90 +2. 75
5. 5
AA
. 0025 . 0034.5 .0023 . 00265 .0044 . 00225
-3, 130 -4, 310 -2, 880 +3, 410 +5,500 +2, 810
TABLE 25.- Stress distribution-High-incidence test
Gatigoen _lo__c_a_-_ 1 3 6 13 16 18
---------------
Gage _______ 85 86 87 88 90 93
---------------
Factor {MV -1.40 - 1.2 - 0.80 +J.45 +2.00 +1.45
2. 0
AA
. 0014 . 0010 .ooo .0013 .0020 .0014
{;yv - 1, 750 -1, 250 - 1, 000 +1,e25 +2,.500 +1,750 -2. 75 -3.00 -l.80 +3.oo +3.50 +3.10
4.0 . 00275 .0026 .0013 .0028 .0033 . 0029
P/A -3, 430 -3, 250 -2, 250 +3,500 +4, 125 +3, 620 rv -3. 75 -4. 50 -2. 90 +4.00 +5.00 + 4.50
6.0
AA
. 00375 .0030 .0029 .0038 .0046 .0041
-4,690 -3, 750 -3, 620 +4, 750 +5, 750 +5, 125 rv -4.00 -5. 20 -3.40 +4.80 +6.00 + 5.20
7.0
AA
.0040 .0045 .0034 .0046 . 00545 .0048
-5,000 -5, 625 -4, 250 +5, 750 +6,800 +6, 000 {MV -4. 40 -6.50 -3.90 +5.30 +6.55 +6.00
8.0
AA
. 00445 . 0057 .0039 .00505 .00600 . 0055
-5, 550 -7, 125 -4, 875 +6. 400 + 7, 500 +6,875 rv -4.50 -6. 90 ;-4. 05 +5.50 +7. 00 +6.10
8. 5
AA
.00455 . 00605 . 00105 . 00525 . 00630 . 0056
-5, 700 -7,550 -5,050 +6, 550 +1, 860 +1,000 r1v -4. 70 -7.20 -4.30 +5.90 +7.50 + 6.50
9.0 . 00•175 . 0063.s . 00425 .0056 . 00675 .0059
P/A -5, 940 -7, 940 -5, 310 +1,000 +8,440 +1, 370 riv -4.85 -7.60 -4. 45 +6.10 +7.90 +6.90
9.5 .0049 .0068 . 0044 . 0059 . 0071 .0062
P/A -6, 120 -8, 500 -5, 500 +1, 375 +8, 875 +1, 750 rv -5.00 -8.00 -4. 60 +6.50 +8.25 +1. 20
10. 0
AA
. 0051 . 0072 . 0045 .0063 . 0074 . 0065
-6, 375 -9, 000 -5, 625 +1. 875 +9, 250 +8,125 {"'lv -5.10 -8.30 -4. 70 +6. 75 +8. 75 +7. 55
10. 5 . 0052 .00745 . 0046 . 00655 . 0077 .0068
P/A - 6, 500 -9, 300 -5, 750 +8, 175 +9,625 +8. 500 {MV -5. 20 -8.65 -4.90 +1.10 +9.10 +8.oo
ll. 0
AA
.0053 . 00785 .0048 .00695 .0082 .0072
-6, 625 -9,800 -6, 000 +8,690 +10, 2-50 +9, 000
32
r
FIGURE 32
FIGURE 34
,,'
FIGURE 33
33
COMPUTATION OF STRESSES AT STATION 1
[B igh incidence-120 pounds per square foot= 15 factors]
.014
FIGURE 35.-Section station 1, 55-loot wing
Areas: Sq. in.
2 upper flange angles ________________________ __ _____ ________ __ O. 574
2 lower flange angles __ ____ __ ___ ______ ___ ___ ________ ___ _____ __ . 674
Front upper web, 0.051X8 __ __ _______ ___ ___ ___ ___ _____ _____ __ . 408
Front lower web, 0.051X8____ __ ____ _____ ____ ____ _____ ____ ____ . 408
Upper corrugated flange, 45.5X0.064Xl.22+2X0.046 _______ ____ 3. 683
Lower corrugated flange, 45.0X0.020Xl.22+ 2xo.020 _________ __ 1. 140
Lower skin=47X0.014 __ __ __________ __ ___ ______ ____ ______ _____ . 658
Rear upper web, 0.036X8 __ ___ ___ ___________________ _____ ___ __ • 288
Rear lower web, 0.036X8. ------------------ --- ------ --- --- - - - . 288
Total area · -- - - - - - -- - ----- --- -- - - - ----- - ----------- - - - -- ---- 8. 021
DETERMINATION OF NEUTRAL AXIS AT
STATION 1
Moments tn.ken about lower chord.
Only box portion considereci.
Upper 0.014 cover neglected.
Considering 8 inches of web as effertivc.
Moment abou t lower chord
Front upper flange angle __ 0.287 (27.75 - 0.72)= 7.75
R!',ar upper fl ange angle ___ 0.287 (24.37- 0.72) = 6.80
Front and rear lower flange angles_ 0.574 (0.72) = .41
Upper corrugated flange
3. 6832S . 87 + 'n~75 + 24.37 -0.38= 97.90
"
Lower corrugated flange __ ___ __ l.14X 0.387=
Front upper web ______ _ 0.408 (27.75- 1.25) =
Rear upper web ___ __ ___ 0.288 (24.38 - 4.25) =
Front lower web ____ _______ ___ _ 0.408 (4.25)=
Rear lower web ____ _______ ___ 0.288 (4.25) ~·=
.44
] 1.68
5.84
1.72
1.05
Total_ ____ -- - -- ---------- - - - ---- - - - - 133.59
Location of neutral n.xis :
133.59/8.021 = 16.50 inches above lower chord.
DETERMINATION OF MOMENT OF INERTIA AT
STATION 1
Front upper flange angle _ _ _ 0.287 (10.53)2 __
Rear upper f1~'1 nge a ngle ___ __ 0.287 (7.16)2 __
Lower flange angles ___ ___ __ 0.574. (i 5.78)2 _ _
Upper corrugated flange ___ _ 3.683 (10.07) __
F t b 0.05l X S
3
ron upper we - - - + O 408 (7' 2 12 . ·- i --
R ear upper web __ 0.036X 8 a +o '>18 (3 63)2
12 .~ . __
Front lower web_ 0·05{2'< 8 3
+0.408 (12.25)2 __
Rear lower web_ 0.03~t 8 3
+ 0.248 (12.25)2 __
Inch •
33. 0
14. 7
143. 0
396. 8
22. 2
5. 3
63. 4
44.. 5
Lower corrugat ed fla nge _____ 1.14 (16.12)2 __
Inch•
296. 0
Lower skin __ ____ ___ __ _____ _ .658 (16.50)2 _ _ 179. 0
TotaL __ ____ __ ___ _____ _____ _____ __ 1, 197. 9
Bending moment at station 1, high incidence, 15
factors=2,342,800 inch-pounds.
M 2 342 800 . 7 ' 1 , 1~ 8 = 1,958 pounds per square mch.
Stress:
Front upper B.ange a ngle_
Rear upper flange angle __
Lbs./sq. in.
1,958X 11.25= - 22,000
l ,958X 7.88= -15,400
l ,958X 12.37= - 24,200
1,958X 16.50= + 32,300
Maximum urdinate _____ _
Lower chord __ __ ______ T ABLE 22.-S tresses at station 1 due to chord loads
Mo-men
t
Member Area arm Mo- " 1. Ah' To-about
me11t ta!
leading
edge
---------:---------- --- -
Front flange angle ______ __ Rear flange angle ___ ___ __ _ _
Upper corrugated flange __ _
Lower corrugated flange __ Upper skin ____ ____ ____ ___ _
Lower skin _____ ______ ____ _
Upper leading-edge skin __ _
Lower leading-edge skin __ _
Upper trailing-edge skin __ _
Lower trailing-edge skin _ _
Front web __ ___ __ _____ ____ Rear web _________________ _
0. 574
• 574
3. 683
1.140
• 665
• 658
.690
.600
.895
. 840
. 816
.496
26. 14
71.64
48.90
48.90
48.90
48. 90
12. 00
12. 50
101. 40
101. 40
26. 40
71. 40
15. 0 22. 7 - --- 295 295
41.1 22. 7 -- - - 295 295
180. 0 0 590 -- ---- 590
55. 6 0 184 - - ---- 184
32. 5 0 122 --- - - - 122
32. 2 0 122 - - ---- 122
~: ~ }36. 5 -- - - I, 715 1, 715
~~:~}52.5 ____ 12,3802,380
21. 7 22. 5 -- -- 413 413
35. 4 22. 5 --- - 252 252
I
Moment of inertia, entire section, about Y-Y·-- -- ------- -- - --- - -- 6, 368
Total less leading edge _ ___ iO. 341 - - ----- - 589. 3 -----1----1------ ----- Total less trailing edge_ __ _ 9. 890 - --- - -- - 429. 3 -- -- - ---- ----- - - - ---
~g~~ g~ii~e,J~~t~~~:::::::: 1~: fci :::::::: ~~~: ~ ::::r::I:::::: :::::
i One-hall of trailing edge.
Loeation neutral axis:
Less leading edge, 57 inches from leading edge.
Less trailing edge, 43.5 inches. ·
Entire section, 52 inches.
Box only, 48.05 inches.
34
LOWER FLANGE ANGLE
Reduction in area due to 3 rivets = 3X 0.101X
0.156 ____ _____ ____ _____ ______ ___ ____ ___ ~ 0474
Reduction in area due to strap = O.lOlX0.156_ . 0157
TotaL _ ___ _ ___ _ __ _ _ _ ___ _ . 0631
Total area of section = 0.287 square inches.
Area less cut-outs = 0.224.
Ratio of .&reas o:;~~ =l.28.
Stresses at rear lower fitting due to chord loads:
23
±0.16X 2,342,000X
6 368
=+1,755 pounds per
square inch. '
Total due to beam and chord loads:
1,755+32,300=34,055 pounds per square inch.
This stress may be considered as acting just outboard
of the rivets at station 1.
At the center line of the fitting, if the stresses vary
inversely with the net area, the stresses will be
l.28 X 34,055= 43,600 pounds per square inch
at 15 factors.
Comparison of computed and allowable upper-flange
stresses, station 1, high incidence: Inches
Pitch of corrugations __ ________________ 2. 50
Depth__________ _________ _____ ___ ___ . 75
Pitch to depth ratio ____ _________ ___ __ 3. 33
Radius of gyration, P------------ -- --- . 268
Radius of curvature, R _ _ _ . 71
Thickness, t_ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ . 064
Bulkhead spacing (center section) ____ __ 12. 5
!:'.=48.5
p
-Ry =L'.0
Lbs./sq. in.
Buckling' P/A, R/ l= lL _______ ___________ -42, 400
Column 1 P/A, C=2 L/p= 98.5 __ ____ __ . ____ -34, 500
Column' P/A, C=3 L/p=,4fl..5 ____________ - 37. 000
Computed P/A, maximum, upper flange= -24,200.
Which indicate.s that the upper flange is heavier than
need be by a considerable margin.
TESTS TO DETERMINE THE CONTRIBUTION
OF THE LEADING AND TRAILING EDGES
After the failure of the wing in the high-incidence
te~ t the break was patched by placi11g a flat sheet
across the break aud securing the two halves to it b:v
1 ee A. D. l!.L 1110, serial No. 3227, Column Properties Corrugated
Aluminum Alloy Sheet, by C. J;' . Greene, captain, Air Corps, and C. G.
Drown.
the use of Kalen screws. The flange angles were
united by riveting a % square cold-rolled steel bar to
the inner part of the augles.
Following this repair the damaged wing was used for
a col!nterbalance and beam and torque loads applied
to the good end, loading and tra ili ng edge deflections
being obtained in ench case.
The beam loads were applied as nearly as possible at
the elastic axis for both the up and down load~.
These tests were repeated :1Jter the remov:tl of the
entire leading-edge coYering from the good wing.
Deflection!' were obtained at the same points as before,
for each deflection point was at a rib, and the ribs were
left intact. A comparison of the deflections of this
and the preceding test should show the contribution of
the leading edge to the total rigidity of the wing.
Finally, the entire trailing-edge covering was removed
and the tests repeated as before. In this cnse a com­parison
of the deflections shows the variation of rigidity
due to the trailing edge with respect to the wing less
the lead ing edge.
Figures 37, 38, and 3!) show the relative twist of the
\ving intact, less leading edge, and less leading and
trailing edges, for three torque loads. These figures
show that the leading edge contributes an appreciable
amount to the torsion:tl rigidity of the wing, and thnt
the trailing edge contributes a lesser amount. In
viewing these curves it must be borne in mind that
curve C represents the increase in deflection of the wing
less b0th leading edge and trailing edge over the wing
less leading edge, and not OYer the wing intact. Hence
the effect of the trailing edge on the wing :ts a whole
will be less than the difference between curves B and
C of Figures 37, 38, and 39.
The effect of the leading and trailing edges on the
torsional rigidity can be seen more clearly from Figure
36, where CTR has been plotted for the three test
conditions.
Figures 41 and 40 show the effect of the leading and
trailing edge on the beam rigidity of the wing for three
loads. It is apparent from these curves that the con­tribution
of the leading a!ld trailing edges to the resist­:
tnce of a heam lond is insignificant. This is reasonable
when it is considered that that part of the leading and
tr:tiling edges, especially the latter, when subject to
compressive loads is practically ineffective, due to
buckling of the flat sheet. This leaves only the ten­sion
sides of the leading and trailing edges to contribute
to the total rigidity. Their total area, however, is
small in comparison to the total area of the upper and
lower flanges.
36
\
' '
FIGURE 38
~ ~
" T
!
-l-\
I
1 I'
-:la-~ - ~ -
~
\
k i\
' ~
i\I.
~\
-- ~ , -,
~t
~ -M--+t-++-t-t-1-+i
~ ~ ~ · t-t-1-HH--<-H-I l - +-+-14-,H--1--1--1 I+
FIGURE 39
37
FIGURE 40
FIGURE 41
38
TABLE 26.-Determination of ej)"ect of leading and trailing edges on bending rigidity
Load
(pounds) 17 15 12
!Load at station 17 (on elastic axis)]
SECOND BENDING 'l' ES'r
D EFLECTIONS, UPLOAD
Station
9
--- F R F R F R ~ FW RW -==-I F --==---- F R F 1'' W RW R
~ LIB l.W QU Q~ QOO QOO QO QW QW Q44 QM QM Q~ Ql2 Q03 QOO Qm Qm
I, 000 2. 58 2. 75 1. 98 2. 12 I. 36 1. 43 . 88 . 88 . 89 . 93 . 51 . 55 . 26 . 29 . 06 . 04 . 09 . 13
l,~ 4. 25 4.48 3.29 3. 48 2.27 2. 39 1.51 1.50 1.54 1.57 1 . 91 .97 .47 .51 . 18 .16 . 19 .22
DEFLECTIONS, DOWNLOAD
500 I i. 26 1 L 351 o. 971 Loo I o. 63 1 o. 70 I o ~2 1 o. 42 1 o. 45 1 o. 43 1 o. 231 o. ;.s I 1,000 1.71 2.80 l.23 2. lO 1.38 1.42 .91 .92 .92 .90 . 53 .06 0..2193 1 o.. 2163 1 o.. ol_o~ I o.. J0O4 o. 03 I o. oo I .o<J . lJ
J, 500 2.99 4.17 3.03 3.19 2.05 2. 15 1. 35 1. 36 1.39 1. 38 . . 82 .83 . 46 .39 . 17 . 14 .JI .15
. -
'l'HIRD BENDING 'l'ES'l'
D EFLECTIONS WJTH LEADING EDGE REMOVED, UPLOAD
500 I I. 22 I l. 30 I 0. 95 I 1. 00 I 0. 62 0. 67 I 0. 39
1
0. 40 I 0. 40 I 0. 43 I 0. 21 I 0. 27 I 0. JI 1, ()()() 2. 70 2. 91 2. 10 2. 20 1. 41 1. 50 . 90 . 91 . 96 . 99 . 53 . 65 . 27 I o.. 1353
I , ~ 4. 45 4. 72 3. 49 3. 63 2. 39 I 2. 51 1. 56 I I. 57 1. 63 1. 66 . 94 1. 05 . 49 . 54
0.. 1005 10..0 049 , 0..0 173 1 0.. 0158 1
.20 .21 . 23 .26
DEFLECTIONS WlTll LEADING EDGE REMOVED, DOWNLOAD
I
~· i l. 25 I ~- 35 I o. 9 I l. 01 I 0. 64 I 0. 70 I o. 43 I 0. 4;! I o. 44 I o. 45 I o. 25 I o. 28 I o. JI I o. J~
1
o. 02 I 0. 05 I 0. 05 I o. 06 ~ooo 20 -OO l.M 200 Lh i.w .a .• .a .E .n .w .n .~ .05 .10 .10 · .14
.L,500 3.84 4.10 2.97 3.13 1.97 2. 11 1. 26 1.25 1.33 1.36 . 72 .84 . 32 .41 .09 .11 .12 .18
- -- ·- - - ---- - -- - ------~-----
FOURTH B ENDING TES'l'
DEFLECTIONS WITH L EADING AND TRAILING EDGES R EMOVED, UPLOAD
500 I I. 24 I l. 29 I 0. 96 I 1. 00 I 0. 63 I 0. 69 I o. 38 I o. 42 I o. 42 I o. 44 I 0. 22 I 0. 26 I 0. 12 I 0. l 3 I 0. 02 I o. 04 I o. 03 1, ()()() 2. 77 2. 92 2. 13 2. 26 1. 43 1. 64 . 90 . 93 . 97 I. 02 . 55 . 63 . 28 . 31 . 07 . 06 . 09 I o.. 1043
l, ~ 4. 44 4. 67 3. 47 3. 64 2. 37 2. 52 1. 53 1. 55 l. 62 1. 68 . 96 1. (}l. . 49 . 55 . 22 . 22 . 22 . 27
DEFLECTIONS WITH L EADJNG AND 'rRAJLING EDGES REMOVED, D OWNLOAD
~ I l. 31 I l. 41 I l. 01 I l. 09 I o. 67 I 0. 71 0. 42 I 0. 42 I 0. 45 I 0. 47 I 0. 22 I o. 30 I 0. 11 ~ooo 264 2M 205 2D 1.n i.~ .~ . ~ .m .oo .a .oo .u
l , ~ 3. 89 4. 25 3. 00 3. 30 1. 99 2. 20 1. 23 1. 27 I. 35 1. 43 . 69 . 89 . 29
~--~-~---'---~-~-~--~-
0.. 2172 1 0.. 0046 1 o.. 0046 10 .. 0038 1 0.. 1106
. 40 . 09 . 07 . 11 . 21
- -
39
TABLE 27.-Determination oj efject of leading and trailing edges on bending rigidity
Load
(pounds)
500
1,000
1, 500
500
l, 000
1,500
500
1,000
1,500
500
1,000
1, 500
500
1,000
1, 500
500
1,000
1, 500
[Load at station 17 (on elastic axis)]
SECOND BENDING TES'!'
MEAN DEFLECTIONS, UPLOAD
-
Station
--17 1 15 12 9 6 3 1
------ ------
Web Web
1.220 I 0.950
0.630 0.425 0. 395 o. 240 0.125
I
0.025 0.010
2. 665 2.050 I. 395 . 905 .885 . 530 . 275 .095 .065
4.365 3. 385 2. 330 I. 540 I. 520 . 940 .490 .200 .175
MEAN DEFLECTIONS, DOWNLOAD
I. 305
I
0. 985
J
0.665
I
0.425
I
0. 435
I
0.255 0.130 0.060 0.035
2. 255 1. 665 I. 400 .905 .920 . 545 .275 .130 . 095
3.580 3. 110 2.100 I. 365 1. 375 .825 . 425 .160 . 130
'!'HIRD BENDING 'l'ES'l'
MEAN DE>"LECTIONS WITH LEADING EDGE R EMOVED, UPLOAD
1. 260
2.805
4. 585 I
g:i~ I ?:~g I o:~!~ I o:~~ I o:~~ I o:~~ I o:r4~ I o:?rg I 3. 560 2. 450 1. 610 I. 600 . 995 . 515 . 230 . 220
- --- - --- - -- --- -- ----
M~~AN DEFLECTIONS WITH LEADING EDGE REMOVED, DOWNLOAD
1. 300 I 0.995 I 2. 590 2. 000
3. 970 3. 050
o. 670 I o. 440 I o. 440 I l. 355 . 855 . 845
2. 040 I. 3!0 1. 290
0. 265
. 515
. 780
FOUR'l'H BENDING TES'l'
0.120 I . 245
.365
0.040
. 095
. 135
MEAN DEFLECTIONS W1TII LEADING AND '!'RAILING EDGES REMOVED, UPLOAD
I . I I I I I I
i.• Q~ Q~ Qm Q~ Q• Q~
~2. 8w45 I 2. 195 I. 535 . 960 . 950 . 590 . 295 aw 2w 1.~ LE 1.000 .~ 0.030 I . 100
. 245
MEAN DEFLECTIONS WITH LEADING AND 'l'RAJLING EDGES REl!OVED, DOWNLOAD
1
21.. 376400 I
4.070
1.050 I 0.690 I 2. 130 I. 675
3. 150 2. 095
0. 445 I 0. 4.35 I . 895 .890
I. 330 l. 310
o. 260 I . 525
. 790
0. 115 I 0. 070 I . 240 .110
. 345 .150
0.050
.100
.115
0.035
.075
. 220
0.035
. 070
• 090
Platform
40
TABLE 28.-Detennination of effect of leading and trailing edges on torsional rigidity
LEADING AND TRAILING EDGE DEFLECTIONS
SECOND 'l'ORSION TEST
Station
load 17 15 12 9 6 3
(pounds) i--~--1-- ~---1--~---1--~--~-~---1--~---. -----1
F R F - I R F R F FW RW I R F R F R F FW R\'i - -R-
~ 0.17 -;;;-~l~-;;:;-~~~~1~~ 0.02 0.08 0.02 0. 00 0.02 0.07
300 .28 .79 . 21 .55 .14 .36 .JO .03 . 13 . 26 .08 . 05 .17 . 04 . 01 .02 .12
375 .40 1.00 .30 .65 .18 .43 .14 . 06 .15 .33 .11 . 05 .20 .04 .03 .04 . 17
'l'BJRD TORSION TEST
150 I 0. 25 I 0. 38 I 0. 18 I 0. 22 I 0. 14 I 0. 16 I ~ .u .n .n .u .~ .u
375 . iO 1. 03 . 50 . 63 . 33 . 46
0. 11 I 0. 05 I 0. 01 I 0. 13 I .20 .11 .08 .25
.M .14 .11 .M
FOURTI! TORSIO~ TEST
0. 09 I 0. 09 I 0. 06 I o. 04 I 0. 02 I 0. 01 I 0. 01 I .• 2 .20 .12 .14 .04 .01 .00
. 17 . 25 . 13 . 18 . 05 . 01 . 00
0. 05
.13
.15
150 I o. 26 I o. 39 I 0. 19 I 0. 26 I 0. 12 I 300 .57 .84 .40 .54 .27 m .M 1.08 .U .68 .H
o.2o I .40
. 52
0.. 0198 1 0.. 0096 1 0.. 1036 1 0.. 3115 1
. 22 . 12 . 16 . 39
0.. 1016 .I 0.. M12 I
.14 . 31
0. 07 I 0. 06 I 0. 02 I 0. 01 I 0. 00 I . 10 . 14 . 03 . 00 . 01 0.. 1072
.13 . 18 . 01 . 01 • 01 .15
COMPLETE WING, SECOND TORSION TEST, d8
Station
17 I 15 I 12 I 9 I 6 I 3 I 1
Distance between stations, dL
Torque 4.2 I 60 I 54 I 48 I 46 I 30
Distance from tip
58 I 109 I 166 I 217 264 I 302
34, 388 0.145
I
0. 101
I
o. 054 I 0.030
I
0.020 I 0.010
68, 775 . 298 . 245 .110 . 070 .040 .022
85, 970 .442 .300 . 130 . 080 . 052 .029
M~; (millions pounds-inch')
34, 388 10. 000
I
20.400 I 34.400 I 55. 000 1-~~
I
104. 000
68, 775 9. 700 16. 800 33. 700 47. 200 79.000 93. 800
85. 970 8.170 17. 200 35. 700 51. 500 76. 000 88. 800
Mean 9.300 18. 100 34. 600 51. 200 78. 000 95. 500
VvlNG LESS LEADING EDGE, 'l' IllRD TORSION TEST, d8
34, 388 0.180 0.110
I
0.060
I
0. 044 0.030
I
0. 018
68, 775 .400 . 248 . 132 .088 .OU .030
85, 970 .560 . 330 .165 .112 .080 .048
lvf ~; (millions pounds-inch•)
34, 388 8. 500
I
18. 700
1 30. 900 137. 500
I
52. 700
I
57. 200
68, 775 7. 200 16. 600 28.100 37. 500 59. 700 68. 700
85, 970 6.450 15. 600 28. 100 36. 800 49.500 u. 700
Mean 7.400 16. 950 29.000 37. 300 54. 000 59.800
WING LESS LEADING AND TRAILING E DGES, FOURTil TORSION TEST, d8
34, 388
68, 775
85, 970
34.,388
68, 775
85, 970
Mean
0.197 I . 450
. 648
.269 .147 .096 .062
. 350 . 192 . 120 . 082
0. 021
. 033
.048
o. 125 I o. 064 I o. 048 I o. 033 I
-~--~--~~--~-----
CrR=ll1 ~;(millions pounds-inch')
7.340
6.420
5. 580
6.450
1
16. 500 I 29. 000 I 15. 300 25. 200
14. 700 M. 200
15. 500 26. 100
34. 400
34. 400
34. 400
34. 400
1
48.000 I 51.000
48. 200
49. 200
48. 200
62. 500
53. 700
54. 900
41
TABLE 28.-Determination of effect of leading and trailing edges on wrsional rigidity- Continued
J,EADING-EDGE PLUS TRAILING-EDGE DEFLECTIONS
SECOND TORSION TEST
150
300
375
17
0. 50
1. 07
1. 40
15
0.43
. 76
.95
12
o. 25
.50
.61
SLation
9
0.19
. 36
.47
6
0.10
.22
. 25
-1
0.09
. 16
. 21
'fHIRD TORSION TEST
150
300
375
0.63
1. 32
1. 73
0. 40
. 85
1.13
o. 30
.62
. 79
0.24
. 45
.60
0.18
. 32
. 42
0.10
. 26
. 31
0. 07
.17
.20
FOURTH 'fORSION 'l'EST
150
300
375
0.65
1. 41
1. 84
0.45
.94
1. 21
0. 32
. 67
. 86
0.24
. 49
. 61
0.18
. 35
. 45
0.13
. 24
. 31
0.09
.15
. 19
WING 'fWlST
SECOND TORSION TEST
150
300
375
o. 403
. 864
!. 130
o. 312
. 552
. 689
0.156
. 312
.380
0.460
.102
.115
o. 039
.070
. 091
THIRD TORSION TEST
\
- 130050 -\-0. 509
!. 065
375 1. 395
o. 290
. 616
.820 I
o:~ I . 492
0.132
.248
. 330
0.090
. 161
. 211
0. 046
. 161
. 143
o. 031
. 074
. 087
FOURTH TORSION TEST
150 0. 525 0. 326 0. 200
300 1. 138 .682 • 418
375 1. 485 . 878 .536
Wing chord __
Distance Crom
71 79 92
tip ___ ______ 37. 5 79. 5 139. 5
WEIGHT DISTRIBUTION
The weights of the individual items entering into
the construction of the wing, and of the assemblies,
together with a per cent weight distribution, are given
in Table 29.
It is felt that the design could be altered in some
ways so as to effect a change in the wing weight. The
first of these possibilities occurs in connection with the
method of lightening both the rib and main webs.
The present boundaries of the lightening holes are
shown by the solid lines of Figure 42.
In section A- A, for instance, there is a considerable
width of web protruding on either side of the stiffening
angle. This flat material carries little compressive
load, as was noted during the high-incidence test.
Similarly, the web material in sections B-B carries
little compressive load and could be omitted. The
suggest ed outline of the lightening holes is shown by
the dotted lines in Figure 42.
Another means of saving weight would be to diminish
the sections of t he vertical stiffening members of the
main web in proportion to the loads that they are· to
0. 132 0.090 0.060 0.039
. 270 . 176 .111 . 065
.336 . 226 .143 . 083 I
104 114 124 131. 3
193. 5 I I 241. 5 287. 5 317. 5
carry rather than to use the same section throughout.
Flanging the outstanding legs of the stiffening angles
would effect an increase in their strength but at a
FmunE 42.-Suggested increase in rib lightening
slight increase in manufacturing cost. The rib flange
angles in the trailing-edge section could be tapered,
retaining their present section at the rear web, and
diminishing toward the trailing edge to a section just
42
sufficient to be riveted to the rib web. Had the out­standing
leg of the vertical stiffener of the nose ribs
been flanged, the failure of the leading edge would have
been appreciably delayed.
When considering the means of decreasing the wing
weight, it would be well to mention the items and
changes that would be necessary ·to place the wing in
service on an airplane. The first items to be con­sidered
would be gas tanks, tank fittings, changes
necessary to install the tanks, etc. Then would come
control rods or cables, aileron supporting fittings,
aileron spars and structure, wiring, Pitot static tubes,
possibly landing lights, etc. The wing as received was
not protected from exposure to weather. Such a
protective coating would be essential. When the
wing was opened for inspection after the test, the steel
parts were found to be rusted, and the gray powdery
film that indicates corrosion was found throughout the
wing. However, no attempt of any kind was made to
prevent this corrosion when the wing was made, for it
was purely a test article and was not expected to be
subjected to any exposure.
TABLE 29.-Weight report-Frame assembly
Name Num- Unit
-------------------1 ber re- weight
Item Station quired (actual)
---------- 1---------1------
Bulkhead •... ------------- -
Do ... -----------------­DO---
------------·- ---­DDoo
•_._ _-_-_-_-_-_-_-_-_-_-_-_-_-_-_-_-_-_- -_
DDOo .-•--• .-.-•--.-.- -. -_-__-_-_-_-_-_-_-_- -_
Do . . . --------- - --------
Do. _ --- -- - - - ___ - - - - - ---
Do. - - - - - - -- -- --- -- - -- - -
DDoo ._•_•__ --__-_--_-_-_-__-_--_-__-_- -_-__- -_
DDOo -__-_-_--__--_-_-_-_-_-_-__-_-_-_--_-_-_ -_
DO--------------------­Do
•.• ... . ..... . . .......
DDoO _-_-_-_--_-_-_-_·_-_-_-__-_·_-_-_-_-_-_- -_
DDoo .___-_-_-_-_-_-_-_-_-_-_-_-_-_-_- -_-_-_-_
DDoO -_-_- -_-__-_-_-_-_·_--_-_·_-_-_-_-_-_-_ -_
DDoo -__-_- -_-_-__-_-_--_-_-_-__-_-_-_-_-_-_- -_
DDoo .__._. _-_-_-_-_-_-_-_- -__--_-_-_-_- - -_
DDoo ._.__. -__-_-_--__--_-__--_-__--_-______ __
Do ___ ____ ________ _____ _
Do ___ _____ _______ _____ _
Do ____ _____ ____ ___ ___ _
Do ____________ _____ __ _
Do _____ ______ ____ ___ __ _
Do ___ ____ ______ ___ ____ Do ___ ________ ____ ___ __ _
o __________________ __ __ _
1-L '- ----- --------- - - - -
21--RL _ '_-_-_-_--_-_-_-__-_-_-_-_-_-. ,-_ -_-__- -_
32--LR _. _-_-_-_-_--_-_-__-_-_-_-_-_-_-_-_-_- -_
43--LH ._-_--_-__-_-_--_-__-_-_--_-_-__-_-_- -_
45--RL _•_._._._ -_-_-_-_-_- -__--_-__- -__--__- -_
5-R--------- - --- -- ----­&-
L. - - ------- -- -- - - - -- -
6-R