y
DIVISION
PERFORMANCE
l configurstion is capable. ml. radios missionett altitudes0 feeteet. The mission is summarised onistance-veight profile is shown on Figureirplane performance is summarised on Figure 3.
Tho missionull powerclimb and cruise. Fuel allowance for take-off and acceleration to climb speed is one minute at full power.
The climb performance is shown on Figure h. The sea level rate of climb0 feet per minute and decreases with altitude toeet per minute ateet. This part of the climb is madeonstant EAS0 Knots and an increasing true speed. arge part of , tho excess thrust is required for acceleration. Aboveeet the climb is madeonstantnd all of the excess thrust is available for cllnb. Ateet the rate of climb0 feet per minute and thereafter decreases rapidly to zero0 feet, the start of cruise. The climbOO pounds of fuel,,ndinutes.
The climbing cruise is made at maximum power at. Ihe cruise
timeoursegree turn at the target. rai.
from take-off at an altitude ofeet. The end of cruise is
feet over the baae at. An actual mission would include an idle
Jt'cMmf AIRCRAFT CORPORATION
DIVISION
PERFOHMAHC5 (COHT.?
power descent. mi. from the base and would use less fuel than continuing the cruise to the base at altitude. Idle powerof the engines at altitude is not yet established making the descent characteristics difficult to define. eserve allowance is includedingleninute loiter at subsonic speeds0 feet altitude.
The take-off and the landing ground rolleet respectively. Speeds required for take-off and landing are based on an angle of attack ofegrees, which is the clearance angle with the Rutin gear struts compressed. This provides an adequate ground clearance margin over theegrees provided with tho gear struts extendod. Single engine safety during take-off is excellent since the total airplane drag la less0 pounds Including dead engine and trim drag and the operating engine provides0 pounds of thrust. Single engine performance during lande, of course, better due to the reduced weight.
In tho ovont of an engine failure at some pointission, two courses of action are open to the.pilot. He can descend to0 faet and subsonic speed and return to base from any point during the mission. Or, ne can maintain his speed stnd descend0 feet. At this flight condition, he can return to base if the engine failure occurs not. roi, from base on tho outbound leg or not over :' n. ml. on the return leg of Vie mission, Between these points the airplane cannot return to base. If the engine failure occurs at the target, the airplane
HpM III-3
rER?cRXA :jce (coht.)
will run out of. ml. short of the base. The single enginereturn capability is shown on Figure 5. If penetration is assumed to occur at the end of. ml. from base, then the airplane canupersonic single engine return to the penetration point from all pcdnte during the mission exceptistance of. mi. before reaching the target. mi. after passing the target. igh degree of multi-engine reliability is assured.
jssiom nma
t.l/ o .
climb
cruise out target crulaa backfw
r. ki. (
lbs. total
bs, hrf used in0 lbo. jplso used in primary)
craft corporation
ClilfORNIA DIVISION
/ 3
( air "
PSPJORKAKCE s'jmhart
n.
Waight
Spssd'
Take-off Oroand 0
Rate of Climb.0 Kts.
Cruise
Mach
Speed
Target
Altitude
Weight
Landing
Weight
Speed
Distance
TRUCTURAL DESCRIPTIOM
Weight and- 2
Design- 9
Material-Design
- 16
- 28
Landing- 35
rv-i
ircraft corporation
F DIVISION
WEIGHT AKD BALANCE
This sectionrief 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 oontrol activity in design. Sufficient analyses have been made of the structure and major aircraft systems to determine the validity of the component welghtej these analyses are the basis for the weight estimate.
The airplane balance is shown on Figure 1, The center of gravity envelopo is tailored to give aunirsum trim penalty during the supersonic position of the mission- while retainings for take-off and landing. The most. is at take-off, as fuel is ueed the eg. moves aft to give the most. at the odd-point of the mission and then forward for landing.
ontains the weight au*rarj followedrief discussion of the component weights on pageso IV-8.
Joe/dm/ AIRCRAFT
CALIFORNIA ISIOK
WEIGHT SDKMAnT
Wing Fin
Fuselage Lending Gear Surface Controls Nacelles
Propulsion Group
Instruments
Hydraulics
Electrics
Electronics
Furnishings
Air Conditioning
Tall Parachute
Weight
Unusable
Zero Fuel
Fuselage
Wing
Take-off
, -- Tin
JoChiimi aircraft corporation
ini BitiilOH
-TIGHT AKD BALANCB
Componont Weight
The wing and fuselage weights are derived froa the structural analyses briefly presented in this section of the report. The fin structure vill be the same type as the wing, reduced in weight due to the lower(load intensities.
Bca Beam
rsnela
Caps
Webs
etc.
Edge
Edge
to Fus.
50
solago
Skin
t* Fin attachments
gear support structure
etc. in Shell
& Canopy
- Equip. Bay, Gear, etc.
Page
WEIGHT AND BALANCE
Component Weight (Coot.)
CALIIDRMIl DIVISION
Wheels and
Strata, Retraction,
Poso
Wheel and
Steering and
Surface Controls
The surface control weight Is baaed on full powered Irreversible
systems. An allowance ofbe. is included ln the autopilot weight to
provide any stability augmentation that nay be required.
Cockpit
Elevon
Rudder
Hacolles
The total weight of this groupb. and includes the air intake system and engine vowl. The engine cowl, that is the portion aft of the front face of the engine, is estimated tob. The air Intake
CALIFORNIA DIVISION
WEIGHT AND BALANCE
eight (Cont,)
Kaceilos (Cont.)
systea as drawn Is ooraovfhat tentative since the inlet configuration vill probably require some developnent, hovever, the weight0 lb, allowed soena adequate for anything that can be envisaged at this tine.
Propulsion Group
ngine weightb. each includes starting provisions and self contained oil system. The fuel is contained in integral wing and fuselage tanks, theuse ofnd KEF will require seme ingenuity in the design of the fuel ays ten plumbing to minimise the weight penalty for this feature. The additional weightb, carried for the HEP system is based on some duplication of pumps, distribution and transfer systems.
Engine
Fuel System
Sealing Baslo System HEF Increment
Instruments Engine Instruments '"nstallation
ho
Joct'flfflf aircraft , . , division
p
Component Weight (Cont.)
lie-tropics
This groin; Includes the navigational and ccamminicatlon equipment described in Hlscellfu-eoua Syete-ia section together with the wiring and supports required, to Install these aysterns In the airplane.
ARCamMzA
ARB kk Radio
Inertia! Jtovigation
mai
425
ejection
oxygen systec. (fixed items
misc. console- &
air conditioning
The air conditioning problem la discussed in Cockpit Environment aeetion. Theove_ceb. for thie Bystemeasonable estimate at this stage.
Page lT-9
ivision
DSSIQN LOADS
Loads used Tor the structural design of this airplane are basedrequirements of Military Specificationith modified The gust criteria modification refers to the variation ofwith altitude as shown by Figure
hows the variation of design speeds with altitude. 0 feet, maximum speed is limited. 0 feet to sea-level the maximum design speed Is kZ$ knots, EAS. The design level flight speednots, EAS shown on this chart was selected0 fps. gust. Calculated aileron reversal speeds are also shown on Figure 3. Adequate wing stiffness within tbe design speed range Is Indicated by theee reversal speeds.
lagrama for guat and maneuver are shown by Figure 2, Foraneuver envelope maximum accelerationsre used. The gust envelope shown ia conservatively based on tero-fuel weight0 lbs. and thsrefore, results ln the maximum design gust load factors.
ultimate design loads for the various airplane componente are included ln the pertinent sections of this report. Except for the forward part of theub-sonic. weight0 lbs.critical loads on both the wing aad fuselage. ps. gust0 lbs. produces slightly higher loads on the forward
HA '
/ aircraft corporation
r ii iifiiriii
pes ion loads (coatlmied)
part of the fuselage. a not critical because fuel used to climb reduces the gross weight0 lba. wing loads for this oonditlon aref thecondition loads. loads for this condition are not critical because the fuel used is removed from the forward fuselage tanks.
;
vtt:;;
3tk) ;
t*
Page rv-lii
IRCRAFT CORPORATION
DIVISION
MATERIAL SELECTIOM
Investigation was made Into new experimental materials available and those still being developed in the laboratory. All of tha common and exotic matale and nod If ic at ions thereof were cooaldered. These wereto each other oa strongth/dsnslty basis, for ultimate, yield and modulus of elasticity, for all temperatures up to For temperatures up totitanium alloys Indicated as good as or better strength/density capabilities. Of the titanium alloya consideredVCA were ahown to be most promising.
From feasibility and producibllityVGA la the moat practt-al and the most efficient in strength at all temperatures up to The aurterlal selected is manufactured by Crucible Steel Corporation, Pittsburgh, Pennsylvania, and is basically an all Seta titanium alloy. Its elementsanadium, Hi chromium and k% aluminum. It can be purchased la theaged, or cold worked and aged conditions. Agingimple heating procedure-for extended periods of time rangingours, followed by air cooling.
This material indloatea the following characteristics!
Good bendability and fcrmabillty.
Oood weldablllty.
Hon-dlrectlonal charaoteristics. li. Ability to be brazed.
ArMrrr/ aircraft
CORPORATION
ivision
MATERIAL SELECTION {Continued)
Cold headabillty.
Readily machined.
Exceptionally low creep rates at elevated temperaturea.
The physical properties of solution treated or annealed material are followai
Deoaityi 5 lbe./cu.in.).
Specific Heati 1
Thermal Expansion. li. Thermal Conductivity! 0
The mechanical properties furnished by material vendor are ae foUovai
*tu "
F^ -
t
Ft -
fty -
%
T
Elastic Modulus
Page IV-lfia
Jf-d'/tecd aircraft corporation
) n
MATERIAL SELECTION (Continued)
The abore values have been verifiedumber of coupon tests in the Lockheed Research Laboratory*
General temperatures expected throughout the airplane structure are expected to bewith peak temperatures on leading edges equal to The above allowables indicate this material bas good mechanical properties ln this range.
Page IV-16
IRCRAFT CORPORATION
CHIFORNIJ DIVISION
WINQ
Description
The construction of the wing is as shown in Figure 5* The structural box extends fronercent toercent of the wing chord. Forward ofercent, the leading edge conalstsolid leading edge^arrowhead and skins supported by Multiple rlbe and stiffensrs perpendicular to the swept leading edge. The structural box itself consists of multiple beansatnches 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 corrugationshordwlse direction. Shear attachment of beams to outside skin ie accomplished by tabs between corrugations. The beams are designed to carry the wing beam bending load and vertical shear.
The surfaces of ths box consist of an outer akin and an innerakin with corrugations runninghordwise 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 spsced approximatelyinches along the span, provides good ohordwise form stiffness.
Aerodynamic heating of the structure resultsemperature 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 oV waving between corrugations in the streamwise direction.
Jfiffifieeti AIRCRAFT CORPORATION
CALIFORNIA
WING (Continued)
Description (Continued)
Hence, the streeoea due to temperature gradient are heldinimum and aerodynamic sraoothnsae ia maintained.
For produclbllltT andoint in the wing laJust outboard of the engine nacelle as shown in Figure 5. Theedge structureJt of chord consists mainly of control surfaces.
Material throughout the wingVCA titanium In various forms. Design Loads
Ultimate wing shear, bending moment and torsion is shown inoreavy weight condition. Thisoom temperature oonditlon. Supersonio "hot" conditions areess and are not critical oa the box structure since the material reduction factor atia.
Section Properties
The airfoil section ia 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.
Original document.
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