PROPOSAL - A-11

• Document Type: CREST
• Collection: General CIA Records
• Document Number (FOIA) /ESDN (CREST): CIA-RDP74B00752R000100170001-7
• Release Decision: RIPPUB
• Original Classification: K
• Document Page Count: 110
• Document Creation Date: December 22, 2016
• Document Release Date: December 9, 2010
• Sequence Number: 1
• Publication Date: March 18, 1959
• Content Type: REPORT

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When I ;saw you in Washington last week, you asked for a proposed second
phase engineering program and cost analysis on the airplane we currently
designate as the A-11. I am attaching two copies of a complete report on
the A-11, which pretty well summarizes all the basic features of the air-
craft. A separate report will be furnished shortly on the radar aspects of
the type.
We currently have authorization to do a certain amount of engineering on
the A series airplanes, through 31 March 1959. For the period 1 April
1959 to 1 October 1959, I would propose the following work be done:
1. Design engineering. This includes the basic engineer-
ing required to carry on wind tunnel testing, major component layouts,
and provide basic information for structural testing.
2. Structural tests. We are to the point where it is necessary
to do a sub stantial amount of testing on titanium structures. W e already
have $10, 000 worth of titanium material, some of which has been used and
tested, but our investigation of this material would have to be greatly
accelerated in the next few .months.
3. HEF testing. We discussed the outlines of this program
briefly with you and Mr. Kiefer during my last visit to Washington. It
would envision setting up a basic part of the aircraft fuel system, shroud-
ing the pertinent parts in ovens capable of simulating temperatures up
through Mach #4. 0, and a considerable amount of work of a chemical nature
on such things as tank sealing material, seals, metals, etc. The basic
problem of how to handle safely the appropriate HEF to be used would be
studied.
4. Wind tunnel tests. It is vital that wind tunnel tests on
both subsonic and supersonic models be run as a next step in the program.
These tests would investigate lift, drag, stability, control problems, and
obtain basic load data for design. A good deal of work would be done on
the nacelle design but, in all likelihood, this phase could not be completed
within the six months period referred to. The low speed tests would be
run in the Lockheed wind tunnel, while the supersonic tests are proposed
for the Ames Laboratory of the NASA.
5. Wing temperature tests. A scale model of as large a
section of the wing as practical would be constructed and method provided
for applying at least 1 g flying loads. It is proposed to put this model into
an NASA tunnel at Langley Field or one of the Tullahoma tunnels of the
Air Force, to get some indication of wing deflection and smoothness under
flight load conditions at Mach # 3. 2.
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Fage 2
6. Cockpit and equipment bay temperature model. The
largest feasible model of this part of the airplane would be constructed and
instrumented completely, to determine heat transfer data to the critical
areas of the model. Data would be obtained for windshield design, cockpit
insulation and equipment bay environment.
7. Mockup. A full scale .mockup of the nose section of the
fuselage, including the equipment bay and the nose landing gear, would be
constructed. A separate mockup would also be made for the power plant
and main landing gear section. This work would not be completed within
six months, but should be about 80 to 85 percent complete in this period.
8. Flutter analysis. The basic aircraft flutter modes would
be investigated theoretically and computed data obtained to indicate the
structural safety for the design flight conditions.
9. Antenna model and tests. Due to the expense of flying the
A-11 type aircraft, and the importance of good communications and navi-
gation, a model would be constructed so that the antennas could be developed
to give optimum performance.
10. Shop layout and tool planning. A small amount of work of
a planning nature would be undertaken to determine the optimum way to
tool the airplane, determine the heat treat facilities required for the titanium,
and investigate the availability of critical items from a time standpoint.
The over-all price for the above six months study is $1, 722, 000. As you
can see, there is no proposal to construct any part of the airplane except
models and test structures. Whereas I think we could schedule a period of
18 months time to an initial flight test with a full go-ahead, if we apply the
following phase approach a somewhat longer period is necessary -- probably about
20 to 22 months. Of course, all the items proposed would necessarily have
to be done under any program to construct the A-11.
You will note that there .is no proposal to build a radar model as such. We
are, of course, willing to make such a model, should you find it desirable,
but my own current thinking is that our scale model approach, considering
all the factors involved, would give us sufficient information to cover the
desirable aspects. of the problem.
The HEF test part of the program above is based on a price of $226, 500, which
does not include the cost of the material to be tested, This number is in fairly
good agreement with the rough estimate I mentioned to you last, of between
$200, 000 and $300, 000. I would be hopeful that the producers of the HEF
would furnish the material to be used (600 to 1, 000 gallons in return for the
results of the testing on the airplane system. It would be contemplated to
build the HEF test rig in such a form that it could be transferred bodily to
our engine fr.iends~ test location, as we did in the previous program, so that
the total compatibility of the system, except for altitude effects, could be
studied with an actual running engine.
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Page 3
I do not have information at this time regarding the over-all cost of 12
aircraft. We are having difficulty in evaluating the exact effect of use of
the titanium in terms of our shop hours. My best horseback guess on the
cost of a 12 airplane program would be between $78, 000, 000 and $85, 000, 000,
in addition to the above engineering study. Of this cost, it appears that
$9, 000, 000 to $10, 000, 000 is the cost of the raw titanium itself.
These costs likewise do not include an astro-inertial guidance system, which
.may be desirable for the type. If these units get into production, they would
have costs varying between $165, 000 and $ 300, 000. If they were hand built,
the~.approach $1, 000, 000 apiece.
I will provide you with the best estimate we can make on an over-all program
cost prior to 26 March 1959.
STAT Sincerely,
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~1~oc~/~ced t ~ax~oxalio~a
SP
Mar. 18 1959
~~~~
TITL! Pi~pOSAL
STAT .
r~~r~eto ~
~ ~
~'P ~~ _ _ STAT
Clarence L. Bohn
Vice President
Advanced Development Pro~ecte
REVIf1ON:
o ? ~ t
.aw .o:. ~
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~~,~fr'~~~'~( AIRCRAFT CORPORATION
TAffi.E OF CONTENTS
Summarg
General Description
PerYormance
Structural Description .
Cockpit Emriromnent
Fuel System
Zhermodynamice
Miscellaneous Systems
Alternate Fuel
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--b~ ~-~
,~r~~~~~'~'f( AIRCRAFT CORPORATION
The airplane herein proposed is designed around two {2) Pratt and
~Jhitney J-58 afterburning engines using HEF type fuel in the after-
burners and~JP-150 in the engines. The fuel load is approximately 65~
HEF and 35~ JP-150. Below 10,000 feet no HEF fuel is burned in order
to avoid undesirable sma~e and contamination.
The airplane has a; 2000 n. mi. mission radius at Mach 3.2 acid
crosses the target at 9L,300 feet as shc~m in Figure 1 in the
"Performance" section of this report.
Provisions are made for a crew of one and a nominal design payload
of 500 lbs. The design strength is consistent with transport criteria.
Modern titanium alloys are used extensively in the interest of
simplicity and weight saving. The strength-temperature characteristics
of these titanium alloys provide for a stretch in airplane speed to
T'~ch 3.5. This is compatible with the J-58 engine stretch potentials
The configuration is as shown in Figure 1 in the "General Descrip-
tion" section of this report. It consists basically of a low aspect
ratio triangular planform wing carrying a long slender fuselage and
the two (2) engine nacelles underneath the wings This arrangement
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rag@ 1-Z
5~~~~~1~~t~/ AIRCRAFT CORPORATION
SUP~'1ARY (Cont. )
is consistent with the maximum in structural simplicity and aerodynamic
performance. In this manner the size and wBight of the airplane is held
to the minimum consistent with mission requirement.
In the section entitled "Alternate Feel" it is sho~,m that the same
airplane can use JP-150 entire],y and accomplish the same 2,000 n. mi.
mission radius at appraxiJaately 1,500 feet less altitude.
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Page II-1
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Sanitized Copy Approved for Release 2010/12/09: CIA-RDP74B00752R000100170001-7
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?11?0?11111 Sanitized Copy Approved for Release 2010/12/09: CIA-RDP74B00752R000100170001-7
Page IV-1
J(,4'&(r/ AIRCRAFT CORPORATION
SECTION IV - STRUCTURAL DESCRIPTION
Item
Weight and Balance
Design Loads
Material Selection
Structural Design
Wing
Fuselage
Landing Gear
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Page IV-2
AIRCRAFT CORPORATION
CALIFORNIA DIVISION
WEIGHT AID BALANCE
This section contains a brief discussion of the weight estimate and
the airplane balance.
The configuration achieves by structural simplicity
the lightest airplane to perform the mission. The weight estimate is based
on the use of present day production techniques and good weight control
activity in design. Sufficient analyses have been made of the structure
and major aircraft systems to determine the validity of the component
weights; these analyses are the basis for the weight estimate,
The airplane balance is shown on Figure 1, The center of gravity
envelope is tailored to give minimum trim penalty during the supersonic
position of the mission, while retaining reasonable c.g.'a for take-off and
landing. The most forward e.g. is at take-off, as fuel is used the e.g.
moves aft to give the most aft e.g. at the mid-point of the mission and then
forward for landing,
Page 3 contains the weight summary followed by a brief discussion of
the component weights on pages iv-5 to IV-8.
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Page IV-3
AIRCRAFT CORPORATION
WEIGHT StTh ARY
Wing
9,1,30
Fin
Fuselage
1,5 0
14
Landing Gear
1,900
Surface Controls
1,120
Nacelles
1,900
Propulsion Group
13,110
Instruments
110
Hydraulics
550
Electrics
300
Electronics
1,25
Furnishings
150
Air Conditioning
750
Tail Parachute
70
Weight Empty
35015
Oxygen
40
Oil
60
Unusable Fuel
100
Pilot
285
Payload
Soo
Zero Fuel Weight
36,800
Fuselage Fuel
30,925
Wing Fuel
17,100
Take-off Weight
814s825
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 I Sanitized Copy Approved for Release 2010/12/09: CIA-RDP74B00752R000100170001-7
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 Sanitized Copy Approved for Release 2010/12/09: CIA-RDP74B00752R000100170001-7
rage IV-5
AIRCRAFT CORPORATION
WEIGHT AND BALANCE
Component Weight
The wing and fuselage weights are derived fY the structural analyses
briefly presented in this section of the report. The fin structure will be
the same type as the wing, reduced in weight due to the lower load intensities.
ELI
Banc Beam
Skin Panels 30000
Board Caps 1,390
Beam Webs 780
Ribs 1,150
Joints etc. 380
6,700
Leading Edge 1,020.
Trailing Edge 1,1180
Fillets-Wing to Fus. 230
Total 9,130
Fin 1,1150
Fuselage
Skin 1,225
Longerons 670
Frames 705
Wing & Fin attachments 350
Landing gear support structure 250
Bulkheads 190
Joints etc. in Shell 3110
Windshield & Canopy 250
Doors - Equip. Bay, Gear, etc. L70
L,550
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Page iv-6
AIRCRAFT CORPORATION
WEIGHT AND BALANCE
Component Weight (Cont.)
Landing Gear
Wheels and Tires
380
Brakes
320
Struts, Retraction, etc.
850
1,550
Wheel and Tire
110
Strut
180
Steering and Retraction
60
350
Surface Controls
The surface control weight is based on full powered irreversible
systems. An allowance of 50 lbs. is included in the autopilot weight to
provide arm stability augmentation that may be required.
Cockpit Controls
45
Autopilot
150
Elevon System
675
Rudder System
250
1:120
Nacelles
The total weight of this group is 1,900 lb. and includes the air intake
system and engine cowl. The engine cowl, that is the portion aft of the
front face of the engine, is estimated to weigh 900 1b. The air intake
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 Sanitized Copy Approved for Release 2010/12/09: CIA-RDP74B00752R000100170001-7
Page IV-7
,16(,&'d AIRCRAFT CORPORATION
WEIGHT AND BALANCE
Component Weight (Cont.)
Nacelles (Cont.)
system as drawn is somewhat tentative since the inlet configuration will
probably require some development, ho,erver, the weight of 1,000 lb. allowed
seems adequate for anything that can be envisaged at this time.
Propulsion Group
The J-58 'engine weight of 5,950 lb. each includes starting provisions
and self contained oil system. The fuel is contained in integral wing and
fuselage tanks, the simultaneous use of JP-150 and HEF will require some
ingenuity in the design of the fuel system plumbing to minimize the weight
penalty for this feature. The additional weight of 200 lb, carried for the
HEF system is based on some duplication of pumps, distribution and transfer
systems,
Engines
11,900
Engine Controls
50
Fuel System
1,160
Tank Sealing
350
Basic System
610
HEF Increment
200
13,110
Instruments
Flight Instruments
25
Engine Instruments
L0
Misc. & Installation
!,5
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 Sanitized Copy Approved for Release 2010/12/09: CIA-RDP74B00752R000100170001-7
Page IV-8
AIRCRAFT CORPORATION
W ST AMD 994=
Component Weight (Cont.)
HY 550
Slecttrics 300
Electronics
This group includes the navigational and communication equipment
described in Miscellaneous Systems section together with the wiring
and supports required to install these systems in the airplane.
ARC 62 Command set
75
ARM 44 Radio Compass 85
Inertial Navigation System 200
Drif'tsight 35
MAl Compass 30
425
Furnishings
Ejection Seat 100
Oxygen System (fixed items) 15
Misc. Consoles & Trim 35
.150
Air Conditioning
The air conditioning problem is discussed in Cockpit Environment section.
The weight allowance of 750 lb. for this system is a reasonable estimate
at this stage.
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 Sanitized Copy Approved for Release 2010/12/09: CIA-RDP74B00752R000100170001-7
Page IV-9
_/4d'K('(! AIRCRAFT CORPORATION
DESIGN LOADS
Loads used for the structural design of this airplane are based on
the requirements of Military Specification MIL-3-5700 with modified gust
criteria. The gust criteria modification refers to the variation of gust
velocities with altitude as shown . bar Figure !i.
Figure 3 shows the variation of, design speeds with altitude. Above
72,000 feet, maximum speed is limited to M = 3.2. From 72,000 feet to
sea-level the maximum design speed is 1425 knots, EAS. The design level
flight speed of 370 knots, EAS shown on this chart was selected for combina-
tion with a t 50 fps. gust.. Calculated aileron reversal speeds are also
shown on Figure 3. Adequate wing stiffness within the design speed range
is indicated by these reversal speeds.
V-n diagrams for gust and maneuver are shown by Figure 2. For the
maneuver envelope maximum accelerations of +2.5 g and -1.0 g are used.
The gust envelope shown is conservatively based on zero-fuel weight of
36,800 lbs. and therefore, results in the maximum design gust load factors.
Ultimate design loads for the various airplane components are included
in the pertinent sections of this report. Except for the forward part of
the fuselage, a 2.5 g sub-sonic maneuver 0 T.O. weight of 85,000 lbs. pro-
duces critical loads on both the wing and fuselage. The +50 fps. gust
condition 0 36,800 lbs. produces slightly higher loads on the forward
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 Sanitized Copy Approved for Release 2010/12/09: CIA-RDP74B00752R000100170001-7
~r,~1-%~~1 AIRCRAFT CORPORATION
Page IV-10
DESIGN LOADS (Continued)
part of the fuselage. A 2.5 g maneuver @ M - 3.2 is not critical because
fuel used to climb reduces the gross weight to 75,000 lbs. Wing loads for
this condition are approximately 86% of the "cold" condition loads. Fuse-,,
lage loads for this condition are not critical because the fuel used is
removed from the forward fuselage tanks.
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 Sanitized Copy Approved for Release 2010/12/09: CIA-RDP74B00752R000100170001-7
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 Sanitized Copy Approved for Release 2010/12/09: CIA-RDP74B00752R000100170001-7
Page IV-1,4
AIRCRAFT CORPORATION
MATERIAL SELECTION
Investigation was made into new experimental materials available and
those still being developed in the laboratory. All of the common and
exotic metals and modifications thereof were considered. These were com-
pared to each other on strength/density basis, for ultimate, yield and
modulus of elasticity, for all temperatures up to 1200?F. For temperatures
up to 800?F titanium alloys indicated as good as or better strength/density
capabilities. Of the titanium alloys considered MST.185 and B-120VCA were
shown to be most promising..
From feasibility and producibility aspects B-120VCA is the most practi-
cal and the most efficient in strength at all temperatures up to 800?F. The
material selected is manufactured by Crucible Steel Corporation, Pittsburgh,
Pennsylvania, and is basically an all Beta titanium alloy. Its elements are
13% vanadium, 3.1% chromium and i% aluminum. It can be purchased in the an-
nealed, aged, or cold worked and aged conditions. Aging is a simple heating
procedure (800?F - 1000?F) for extended periods of time ranging from 8 to
100 hours, followed by air cooling.
This material indicates the following characteristics*
1. Good bendab it ity and formability.
2. Good weldability.
3. Non-directional characteristics.
Ability to be brazed.
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 Sanitized Copy Approved for Release 2010/12/09: CIA-RDP74B00752R000100170001-7
Page IV-15
AIRCRAFT CORPORATION
MATERIAL SELECTION (Continued)
5. Cold headability.
6. Readily machined.
7. Exceptionally low creep rates at elevated temperatures.
The physical properties of solution treated or annealed material are
as follows:
1. Density: 11.82 GMS./c.o. (0.175 lbs./cu.in.).,
2. Specific Heat: .131 BTU/lb./?F.
3. Thermal Expansion: 5.2 x 10-6 in./in./?F (68 - 200?F)
4. Thermal Conductivity: 3.90 BTU/hr./Ft.2/?F/Ft.
The mechanical properties furnished by material vendor are as follows:
Annealed
Room Temp.
600?F
Ftu - psi
152,000
109,000
Ftv - psi
151,000
103,000
% Elong.
12
21
Elastic Modulus - psi
111.3 x 106
13.2 x 106
Aged
Boom Temp.
600?F
Ftu - psi
200,000
175,000
Fty - psi
190,000
1115, ooo
% Elong.
5
9
Elastic Modulus - psi
15.3 x 106
13.8 x 106
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Page IV-15a
J('4'&64/ AIRCRAFT CORPORATION
MATERIAL SELECT IDN (Continued)
The above values have been verified by a number of coupon tests in
the Lockheed Research Laboratory.
General temperatures expected throughout the airplane structure are
expected to be 500?F with peak temperatures-on leading edges equal to
780?F. The above allowables indicate this material has good mechanical
properties in this range.
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Page IV-16
AIRCRAFT CORPORATION
Description
The construction of the wing is as shown in Figure 5. The structural
box extends from 15 percent to 80 percent of the wing chord. Forward of
15 percent, the leading edge consists of a solid leading edge arrowhead
and skins supported by multiple ribs and stiffeners perpendicular to the
swept leading edge, The structural box itself consists of multiple beams
sapoed at 16 inches along the chord.. Beams are built up. of beam caps, webs
and stiffeners.. Caps are located under contour in order to allow for the
passage of surface corrugations in a chordwise direction. Shear attachment
of beams to outside skin is accomplished by tabs between corrugations. The
beams are designed to carry the wing beam bending load and vertical shear,
The surfaces of the box consist of an outer skin and an inner corru-
gated skin with corrugations running in a ohordwise direction. This surface
structure is designed to carry normal pressures to the beams and to resist
wing torsional moment. This type of surface design, acting together with
intercostal ribs spaced approximately 40 inches along the span, provides
good chordwise form stiffness.
Aerodynamic heating of the structure results in a temperature gradient
from outside skin to inside structure. This gradient can be accommodated by
this type of structure easily since expansion of the outside surface results
only in buckling or waving between corrugations in the streamwise direction,
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64.k'ld AIRCRAFT CORPORATION
Page IV-17
WING (Continued)
Description (Continued)
Hence, the stresses due to temperature gradient are held to a minimum and
aerodynamic smoothness is maintained.
For produoibility and transportability, a joint in the wing is pro-
vided just outboard of the engine nacelle as shown in Figure 5. The trail-
ing edge structure from 80% to 100% of chord consists mainly of control
surfaces.
Material throughout the wing is B-120VCA titanium in?various forms..
r~+G?!7
\Fl:
Design Loads
Ultimate wing shear, bending moment and torsion is shown in Figure 6
for critical 2.5 g heavy weight condition. This is a room temperature
condition at M a 0.8. Supersonic "hot" conditions are 14% less and are
not critical on the box structure since the material reduction factor at
500?F is only 10%.
Section Properties
The airfoil section is presented graphically in Figure 7. Using this
section and the wing basic dimensions, the structural section properties
are calculated and presented graphically in Figure 8.
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 P- -.ems'..... ---- Tv-19
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CHFC-KFn RY CALIFORNIA DIVISION --m- -
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 .,..matrn o.. To~20
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lei
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 Sanitized Copy Approved for Release 2010/12/09: CIA-RDP74B00752R000100170001-7 r-2)._____
DATE LUt-F%rlttU AIKC,KAI- I (,UKPUKA I IUN MODEI
CHECKED BY CALIFORNIA DIVISION REPORT NO.
I '
T-77 . . .
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Page IV-22
AIRCRAFT CORPORATION
WING (Continued)
Internal Loads and Analysis
The internal loads are calculated from the critical external loads
given in Figure 6 for the subsonic room temperature condition by means
of the structural section properties given in Figure 8,
The beam cap design loads, stresses and cross section areas are
sunnarized in Figure 9. The axial load shown is for the highest loaded
beam. All beams are similar in cross sections and as noted in the figure
have a constant area for most of their span. This makes for ease of fabri-
cation and is efficient because tapering of material is accomplished by the
number of beams decreasing with span station. Beam caps are machined from
B-120VCA titanium rolled bar.
Beam web ,design shear flows, stresses and web gages are summarised in
Figure 10. Due to the effects,of beam taper, the vertical shear in'.the
beam webs is very low and a minimum gage of .016 sheet is sufficient.
Material is B-120VCA titanium cold rolled sheet. Stiffeners are sheet
metal angles of the same material spaced at approximately three inches
along the beams. Front and rear closing spare are of similar construction
but web gage is .0110 in order to maintain the torque box stiffness.
Wing upper and lower surfaces are designed by the torsion shear flows
given in Figure 11 plus the effects of bending due to air loads and, in the
case of the wing fuel tank region, fuel vapor pressure. The outer skin is
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Page IV-23
(-Klrw AIRCRAFT CORPORATION
WING (Continued)
Internal Loads and Analysis (Continued)
,020 B-120VCA titanium cold rolled sheet and the inner skin is .025
B-120VCA titanium sheet which is formed in the annealed state and then
heat treated. The depth of the corrugation varies according to the
shear stability and pressure load bending requirements along the span.
The wing torsional stiffness for aileron effectiveness is pre-
sented in Figure 12,
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 Sanitized Cony Approved for Release 2010/12/09: CIA-RDP74B00752R000100170001-7 TV-4,
CHECKED BY. MODEL -_
CALIFORNIA DIVISION
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 U
4.
6
Tff 25
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 Sanitized 7
Copy Approved for Release 2010/12/09: CIA-RDP74B00752R000100170001- "~
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 _ /oo 2Q0 i s 300 .: y
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rN~,r?
LOCKHEED AIRCRAFT CORPORATION MOOS /L_~
rAI ?rrn.n? '..?.. r.....
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Page IV-28
JAIRCRAFT CORPORATION
CALIFORNIA DIVISION
Description
The fuselage consists of three major assemblies; the forward, mid
and aft sections as noted on the Inboard Profile. The construction of
the fuselage in three sections will greatly facilitate fabrication of
the structure and installation of the functional equipment required in
each section. The provision of service joints on these fuselage sections
permits rapid disassembly of the aircraft for transporting purposes.
The forward fuselage section contains the Flight Station, Military
Equipment compartment, nose landing gear, air conditioning compartment
and suitable compartments for the installation of electronic, navigation
and communication equipment. The remainder of the forward section con-
tains the forward fuel tanks.
The mid fuselage section provides for attachment of the wing box
section and contains the main landing gear and mid section fuel tanks.
The aft fuselage section' provides for attachment of the aft portion
of the fin box section and contains the aft section fuel tanks and the
landing chute.
The fuselage fuel tanks are of the integral type providing maximum
fuel capacity for a minimum size structure.
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/
AIRCRAFT CORPORATION
Page IV-29
FUSEUGE (Continued)
Description (Continued)
The lVaelage structure is of semi-monocoque construction, consisting
of skin, rings and four longerons. Since most of the fuselage structure
adjacent to the skin is subjected to high temperatures for long periods
of time, the material used is a titanium alloy (B-120VCA). For internal
structure, where temperature is maximum at 300?F, 20214T6 or 2021jT81 aluminum
alloys will be used. The minimum skin gage is .016. at the nose, increasing
to a maximum of .0440 at the center section. Rings will be of gage comparable
to the akin except the main frames in the center section. Rings (2.0 in.
deep channel sections) will be spaced approximately 15.0 in. c.o., with two
(1.0 in* deep) $ section intermediate rings spaced between, giving a-panel
spacing of 5.0 c.c. Four longerons, B.L. 14.0, left and right, resist up
and down bending moments. Side bending is restated by tension in the side
skin and B.L.- 14.0 upper and lower longerons on the.-. opposite, aide. hongerons
will be formed sheet metal channels, with'inner and outer caps tying the
channels together. The outer '' cap. also acts as a splice plate for the skin
and rings, and the inner cap 'aa a splice plate for the inner flange of the
2.0 in, deep rings. Spot welding will be used extensively because of weight,
low cost, reliability and strength.
The fuselage shell will be made in four parts, spliced longitudinally
at the longeron points. This "quarter shell" breakdown permits spotwelding
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,,4 'krW AIRCRAFT CORPORATION
Page IV-30
FUSELAGE (Continued)
Description (Continued)
to be used extensively. The "quarter shells" will be spliced together by
a maximum of two longitudinal rows of.titanium (B-120VCA) rivets at each
of the four longerons.
Figure 13 is a shear and moment curve, for the forward fuselage,
critical for.room temperature condition. The shear and moments for ele-
vated temperature conditions are almost as critical. Figure 14 shows the
longeron loads, areas, stresses, akin shear flows and skin thicknesses re-
quired. A detailed sketch, Figure 15, of typical lower longeron is shown.
The upper longeron is similar but approximately half of the area of the
lower longeron at any given fuselage station. Also a sketch, Figure 16,
showing typical side shell construction and ring splice at longerons, is
included.
The cockpit section is similar to the basic shell except-that the
upper longerons support the canopy and cockpit pressure loads. Pressure
bulkheads in this area and other internal structure will be considered to
be made of 20214ST aluminum alloy if temperatures are below 300?F.
Fuselage skin is also considered to carry internal pressure of
15.0 psi ult. due to fuel pressure in the fuselage fuel tank region.
Surge bulkheads, where temperatures remain below 300?F will be made of
2021ST aluminum alloy.
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_------ ~. . ,- r, a.r~r~r 1 L.L'RrIKP1 I IUIV
MODEL ...~._~1_._...
REPORT NO
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 Sanitized Copy Approved for Release 2010/12/09: CIA-RDP74B00752R000100170001-7 -
Prepared
Checked
Approved
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 Sanitized Copy Approved for Release 2010/12/09: CIA-RDP74B00752R000100170001-7
LOCKHEED AIRCRAFT CORP.
77
L ``"Q ~'CU coo '77
~9,!Lovs~9.8.CE"S
F = %/c oc c ~f/
~'/4 TL
r
.200, 000.
7 el/1"-7
/AZ= _ ,203 Ooc
0 -1
TEMP. PERM.
Page I -33
Model A -//
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NAME I DATE
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I Prepared I
LOCKHEED AIRCRAFT CORP.
TEMP.
IV
P6fU1.
Model A -/1
Report No.
TYf'/C~ Z~ 5/UE 5// ZL
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AIRCRAFT CORPORATION
Page IV-35
LANDING GEAR
The landing gear is of the conventional tricycle configuration with
both main and nose gears retracting forward.
Static MG hub is at Sta. 915, WL 29.6. Static NG hub is at
.Sta. 477.5, WL 214.2.
The main gear stroke is 181 in. and NO stroke is 15 in. With the
ratio of 85,000 lbs. takeoff, weight to 140,000 lbs. landing weight, and
by virtue of the lengthy MG stroke, gear strength capabilities are deter-
mined mainly by ground handling conditions, (taxi, braking, and tow).
V
Y1 ~, (for landing) - L
C
b h_
`sTP'Ur
Assuming 7 ft./sec. sink speed and - .9,
strut
149
.2x.9 - .84
and for the ground conditions, static loads, with 00 at 25% MAC (Sta. 868) -
859000 390.5
PV r (~37 5) - 38,000 lbs.
QM
PVG 85,000 - 76,000 - 9,000 lbs.
N
The main gear tires are 40 x 12 of 26 ply rating. Nose gear tire is
26 x 6.6 EEP, Type VII.
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10
AIRCRAFT CORPORATION
Page V-1
COCKPIT ENV IIONMENT
The general arrangement of the cockpit, windshield and canopy is
as shown in Figure 1.
This is the simplest and lightest configuration which we believe
adequate to provide the required comfort and safety for the pilot in
the flight regime which this airplane will encounter.
In this section the problems of air conditioning, emergency escape
and personal equipment are given separate consideration.
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'//Cd AIRCRAFT CORPORATION
Page V-3
AIR CONDITIONING
Pressurization
The gaseous nitrogen released from liquid storage for cooling pur-
poses, as discussed below, serves also to pressurize the cabin. Various
cabin pressurization schedules were investigated, each selected for study
on the basis of its particular effect on the following "critical" criteria:
Nitrogen required for pressurization alone.
Cabin differential pressure.
Pilot comfort in descent.
Isobaric Schedule.- Unpressurized ram operation from. sea-level until
26,275 feet cabin altitude isreached, then cabin altitude remains isobaric_
On a 400 knot descent the cabin rate of descent goes as high as 33,500 fpm;
however, the more important rate of absolute pressure increase associated
with this at the altitude concerned is exactly the same as the maximum en-
countered with the militazy-type schedule below, and involves less sustained
time at high rate.
~ti tine khq r ;: i eon. .. 1 i
During descent-at 200 knots the maximum cabin rate of descent is only 3930 .
2 `pounds for the same flight; using the .constant rate of climb system below).
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Page V-14
AIRCRAFT CORPORATION
AIR CONDITIONING (Continued)
Pressurizatio&(Continued)
Constant Cabin Rate of Climb Schedule - The cabin altitude changes
at constant rates set by the pilot to carry the cabin from initial to
final altitude in the exact time spans of the airplane's climb and descent.
A cabin differential. of 5 psi is reached at initial cruise altitude, and
remains constant during cruise, resulting again in a 26,275 ft. cabin at
maximum altitude. However, this system has the inherent peculiarity of
requiring cabin differential to reach a high of 6 psi on the way to or
from cruise altitude. It was selected for study as requiring the maximum
pressurization nitrogen of any practical schedule, using 202 pounds for
the mission flight. noted above. Pilot comfort during descent, on the
other hand, is by far the best of any possible system, as indicated by
the cabin pressure rate changes of 1160 fpm during 200 knot descent and
only 8070 fpm at 400 knots. The latter rate compares to the above noted
33,500 fpm maximum for the isobaric system. and to 21,200 fpm reached with
the military-type system below.
Military Type Schedule - Unpressurized ram operation from sea-level
until 5,000 ft. cabin altitude is reached, then isobaric pressure is held
at 5,000 ft. until cabin differential has built up to 5 psi, with constant
5 psi differential at all higher airplane altitudes (above 18,365 ft).
Since the first two systems above spanned the mission flight nitrogen re-
quirement from minimum to maximum, this more normal system's nitrogen usage
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Page V-5
i(I(*1~&6KI AIRCRAFT CORPORATION
AIR CONDITION3140 (Continued)
Pressurization (Continued)
Military-Type Schedule (Continued)
was investigated for descent only. Here the 200 knot descent nitrogen
amounted to 33 pounds, compared to 12 pounds and 40 pounds on the iso-
bario'andconstant rate systems, respectively, for the same descent.
400 knot descent Witrogen was also calculated for this system, amount-
ing to only 9 pounds. The main use made of this particular schedule,
however,,- was in investigating the relative pilot comfort between it and
the isobaric schedule, during the 400 knot descent (34 minutes). On the
military-type schedule, after leaving maximum cruise altitude at time
zero, the pilot spends the first 125 seconds subjected to cabin descent
rates varying from 0.85 up to 15.5 inches of mercury/minute, whereas,
with the isobaric schedule no cabin change whatsoever occurs during this
time span. For the remaining 70 seconds both of these schedules would
follow the same rate curve (approaching 21 inches of mercury/minute at
sea-level), except that at 20 seconds the pilot with the military-type
system starts a half minute of reprieve at zero rate in his 5,000 ft.
isobaric cabin, then returning to the high rate for the final 20 seconds.
This comparison shows both systems to be relatively severe on the pilot,
such that he should not attempt such a rapid descent unless blessed with
exceptional ear and nasal passages, or in an emergency. Study of the
exact rate curves vs. elapsed time would seem to give the isobaric system
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AIRCRAFT CORPORATION
Page V-6
AIR CONDITIONING (Continued)
Pressurization (Continued)
Military Type Schedule (Continued)
a slight edge over the military-type, on the basis of sustained times
at high rate of change; this could probably be a point of argument
between any two given pilots, however.
For the present, no attempt is being made to finalize the pressuri-
zation schedule, but merely to have at hand the information on which the
discussions above were based. This is necessarily the case because of
the inter-retionship between the nitrogen required for pressurization,
and that required for cooling. It is felt that until final decision on
the cooling system is reached, it will not be possible to give proper
consideration to all factors for both systems.
Thus, no physical concept of the actual control components for
pressurization can be stated at present; however, Figure lo-shows schemati-
cally a simplified general concept covering the inter-related controlling
that must be accomplished. The master controller, whatever its form, must
accept signals from both the temperature and the pressure sensors, and then
influence both the outflow and the nitrogen flow valves accordingly. For
example, with an increasing cooling requirement the master control must
simultaneously increase the flow of cooling nitrogen, while opening the
outflow valve to prevent over-pressure. Note that this example by itself
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Page V-7
AIRCRAFT CORPORATION
FIGURE 1c.
J*
Ti
~~ _ .off exa~a
c- I-
Z co
u]c~~'1 .zAU
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AIRCRAFT CORPORATION
Page V-8
AIR CONDITIONING (Continued)
Pressurization (Continued)
Military-Type Schedule (Continued)
does not point up the necessity for such an inter-relating master control,
since it is obviously quite a normal function of a pressure sensor to di-
rectly control its outflow valve towards open for such a case. However,
reversing the example, assuming sufficient decrease in cooling-nitrogen
inflow to drive a direct-controlled outflow valve fully closed, would re-
sult in depressurization. (The outflow valve by itself cannot "pump up"
the cabin, being capable only of controlling a higher pressure generated
at or beyond the point of cabin inflow. Note the dissimilarity between
the more normal case of an "infinite" bleed air source available to a
cabin, and the present "release it only as you need it" source). Now the
need for the master becomes more evident, since it must recognize that
even though temperature-wise the nitrogen flow can be reduced, it must
still signal for nitrogen as an inflowing pressure source. (The tempera-
tore controller at this time will function only to position the recircula-
tion bypass valve or valves).
For the latter condition of pressure-nitrogen requirement exceeding
that for cooling, the outflow valve will close completely so that the
only nitrogen flow will be that required for leakage make-up (plus or
minus that involved in maintaining the contained weight of cabin atmosphere
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l~,rf~'~rrr~ AIRCRAFT CORPORATION
Page V-9
CALIFORNIA DIVISION
AIR CONDITIONING (Continued)
Pressurization (Continued)
Military-Type Schedule (Continued)
during climb and descent). The "pressurization alone" values for nitrogen,
quoted under the various schedules above, were calculated on this basis.
Note in Figure (]s that the series arrangement of outflow and safety
valves gives double protection to the pilot against cockpit depressuriza-
tion. For example, if the equipment bay were to depressurize for any
reason the cockpit remains fully pressurized, As an alternate example,
if the cockpit's outflow valve became stuck in the open position, the
cockpit would again remain fully pressurized by riding on the equipment
bay's valve, The probability of simultaneous-open failures is very low.
With the ram operation proposed for the unpressurized portions of
the above described schedules, and the variable pressure source available
during pressurization, it is considered that no negative pressure differ-
ential problem can normally exist, even during the 1100 knot descent. In
this regard calculations were made to determine the required variation between
nitrogen flow rates for the 1100 knot and 200 knot descents, to maintain
full pressurization on the military-type schedule. This had been con-
sidered a possible problem on the fast descent from the standpoint of
that nitrogen portion required just to increase the contained weight
of cabin atmosphere. The results show, however, that even though the
1100 knot descent time was faster by a factor of almost 6, its nitrogen
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Page V-10
AIRCRAFT CORPORATION
AIR CONDITIONSIO (Continued)
Pressurization (Continued)
Military-Type Schedule (Continued)
discharge rate (including leakage make-up) merely doubled. Further in
regard to negative differentials, for the case of a nitrogen system failure
during rapid descent, the outflow and safety valves will all be of the
vacuum relief type and so sized as to prevent excessive structural loading.
Cooling
In the early stages of investigation, a look was taken at air-cycle
ram cooling, with several variations of machinery and water boilers. As
might be expected, the size and weight of the required equipment, plus
material development problems due to the temperatures involved, eliminated
this as a possible solution.
Engine bleed air was peremptorily eliminated for cabin use due to
the airplane performance losses associated with bleed at our altitude.
Note, however, that it is planned to use limited amounts of bleed air
for windshield defogging and for ram air heating, if required, during
the unpressurized portion of the pressurization schedules discussed
above. The latter usage would become especially important were in-flight
refueling to be considered, as here the normal time of ram operation
would be far exceeded.
The most recent investigations have been aimed at accomplishing
certain assumed design directives as followat
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AIRCRAFT CORPORATION
Page V-u,
CALIFORNIA DIVISION
AIR CONDITIONING (Continued)
Cooling (Continued)
100?F space temperature in the cockpit
and equipment bay.
135?F maximum touch temperature of the
trim in areas not directly over con-
ductive structure, with minimum possible
touch temperatures in all other areas.
Cooling to be accomplished entirely by
liquid media stored aboard the airplane
(nitrogen, and possibly water).
Cooling medium to double as cabin
pressure source, as discussed above
under pressurization.
The above temperatures take into account the fact that the pilot's
comfort will be at the much more suitable level associated with direct
suit ventilation by nitrogen gas, as on the R-15. This will include
pilot-selected temperature controls.
One of the most attractive features of having aboard liquid nitrogen
is the ease with which spot cooling of critical areas or equipment components
can be accomplished. Thus such local areas are considered to be no problem.
A recirculation system was investigated on the basis of the above
temperatures, wHWeia cabin atmosphere was cooled in two stages: first
by passage through the air side of a water boiler, and then by injecting
into it liquid nitrogen which topped-off the required cooling. This system
was considered unattractive weight-wise at the time,
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CORPORATION ie V-12
CALIFORNIA DIVISION
AIR CONDITIONING (Continued)
Cooling (Continued)
The most *fecent work, Just completed, was a detailed study made of
a water-panel system, for handling the major cooling load in those portions
of the cabin wall between structural rings. Until the final stages of this
investigation were reached, and the results could be integrated, this system
looked extremely attractive. Regrettably, the final system weight has turned
out to far exceed that of much less elaborate systems, even though as ex-
pected the amount of water expended was very small.
The studies made to date serve to indicate that the cooling problem,
while severe, is not so extreme but that it is completely feasible to
accomplish a practical system within the weight allowance set forth else-
where in this report. This would be so even for a "nitrogen alone" system,
and note in this regard that nitrogen's heat of vaporization amounts only
to approximately a tenth that of water, at the pressures involved.
In ensuing investigations it is intended to exploit still further the
advantages of using water's high heat of vaporization, in combination with
top-off cooling by the "double-duty" pressurization nitrogen.
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