X-15 Design Proposal

8/8/2019 X-15 Design Proposal

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NOTICE: This document conta ins informationaf fec ti ng the na tional defense o f the Un it ed S ta teswithin the meaning of the Espionage Laws, Title18, U.S.C., Sections 793 and 794. Its t ransmiss ionor the revelation of its contents in any manner toan unau thor ized per son i s p rohibi te d by law.Thi s documen t contains prop ri et ary informat ion o fNorth Amer ic an Aviat ion, Inc ., a nd i s s ubmi tt edinconfidence . Information conta ined herein shouldnot be disclosed to other than military organiza-t ions of the U. S.Government .

A-.ORTH AMERICAN AVIATION, INC.

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REPORT NO. NA-55-2219 MAY,1955SERIAL NO.

DVANCEDCH AIRPLANE

DESIGN SUMMARYNORTH AMERICAN AVIATION,I NT ERN A T ION A l A I R PO R T . lO 5 A N GEL E 5 4 5, CALI FOR N I A

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Introduction

In response to a request for proposals for a New Research Aircraft"Project 1226, " North American Aviation presents in this brochurea summary of a proposed design. Additional detailed reports cover-ing all phases of this proposal are listed under Supporting Data.

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DESIGN CONCEPT

PROORAMOBJECTIVE: The program objective isinterpreted as exploration of the flight spectrumwhich will be traversed by a future generation ofmilitary aircraft. The regime of interest en-compasses an altitude of 250,000 feet and speed of6600 feet per second. The program, to be suc-cessful, must provide information which is usefulin the design of aircraft in the near future. Theelement of time is therefore important, and theattainment of flight within 2-1/2 years, as spec-ified, is vital.

DESIGNOBJECTIVE: It follows that the design ob-jective must be to provide a minimum practicaland reliable vehicle capable of thoroughly ex-ploring this regime of flight. Limiting factors aretime, safety, state of the art, and cost.DESUNSOLUTION: It has been determIned that thespecification performance can be obtained withvery moderate structural temperatures; however,the airplane has been designed to tolerate muchmore severe heating in order to provide a practical

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temperature band within which exploration can beconducted, and to provide a safety margin for theunpredictable.Available aerodynamic data indicates that theconfiguration presented is reasonable when thecomplete speed range is considered. The all-movable surfaces for pitch, roll, and directionalcontrol are known to be satisfactorily effectiveat the higher Mach nwnbers. Negative dihedral isincorporated on the horizontal tail to lessen abrupttrim changes due to shock impingement or wakeimmersion. Small variable thrust rocket motorsarranged to provide moments about the three con-trol axes comprise a separate space control sys-tem during periods of ineffectiveness of the aero-dynamc cmtrols due to reduced dynamic pressures.The problem of providing a landing system com-patible with a simple fixed lower vertical tail hasbeen solved by simply allowing the airplane totouch down and "rotate in" about the tail bumperand providing adequate energy absorption in themain and nose gears.The airplane presented is based upon a specificpropulsion system; however, it appears feasibleto use any engine or engines in the same perform-ance category.A secondary, but important, factor considered inpreliminary design is the desirability of meshingwith the present operational pattern for researchaircraft. By following the established patternof operations, a considerable saving in learningtime should be achieved.

s.... NA-55-221The basic philosophy regarding the thermal prob-lems associated with bearings, lubricants, elec-trical insulation, seals, actuating fluids, etc, isto provide a controlled environment to suit presentthermal capabilities rather than attempt to extendthe thermal limits o~these components by furtherdevelopment.The basic concept of a specialized research air-craft, of which only several will be produced andwhich do not require being rapidly serviced as ina combat aircraft permits considerable compro-mise in favor of extreme simplicity in order toassure a high degree of ruggedness and reliability.NAA, through detailed study of current researchaircraft and as a result of conferences with theoperating personnel is acutely aware of and hasincorporated this concept in the preliminary de-sign of the proposed aircraft.Detailed definition and solution of all problemswhich will be encountered in this program arebelieved impossible for a proposal of this scope;indeed, if this were possible, there would be little.need for a research airplane. However, it hasproved possible to design an airplane capable ofsuccessfully coping with these problems byallow-ing for easy modification of critical areas if theneed arises.CONCLUSION: North American Aviation has thetechnical ability, the facilities, and the desire toproduce a successful research vehicle having thedesired performance and operational flexibilityrequired for a long-range research program.

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DESIGN BRIEFPERFORMANCEMAX. VELOCITY AT BURNOUT (DESIGN MISSION). 6800 FT. PER. SEC.MAX.ALTITUDEDURINGCOAST(DESIGNMISSION).. . . .. 250,000 FT.TOTAL FLIGHT TIME (DESIGNMISSION) . . . . . . . . . . . . . . .. 20 MIN.MAX.ATTAINABLELTITUDE. . . . . . . . . . . . . . . . . .. 800,000 FT.MAX.TIMEOF "WEIGHTLESS"FLIGHT. . . . . . . . . . . . . . .. 6.5 MIN.WEIGHTEMPTY. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 9, 959 LB.USE FUL LOAD . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17, 763 LB.(PROPELLANTONLY).. . . . . . . . . . . . . . . . . . . . . . . .. (16,410 LB. )GROSS.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 27,722LB.POWER PLANTREACTION MOTORS INC. XLR-30RM2MAX.THRUST,0,000FT . . . . . . . . . . . . . . . . . . . . . . .. 57,000 LBS.WINGAREA . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 200 SQ. FTSWEEP . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 25 DEGREESTHICKNESS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 5 PERCENTASPECTRATIO.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 2. 5

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GENERAL ARRANGEM

FLIGHT CONTROLSRolling tail; horizontal tail surfaces operatein unison for longitudinal control and differentiallyfor lateral control.All-movable vertical tail provides directional control.SPACE CONTROLSHydrogen peroxide rocket motors for three-axis control

SPEED BRAKESIndependently operable split trailing edge surfaceson upper and lower vertical tail.

ELECTRONICSUHF Command set AN/ARC-34.Radio homing AN/ARA-22.Voice recorder ANI ANH-3.ELECTRICAL

Two 28-volt, 100-ampere generators.750-volt-ampere, 115-volt , three-phase inverFLIGHT TEST INSTRUMENTATION40 cubic feet environment free.

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'ter.

ROCKET ENGINERMI XLR -30RM2 rated at 57, 000 pOWlds ofthrust at 40.000 feet altitude.

PROPELLANTForward integral tank contains 907, 5 gallonsof liquid oxygen.Aft integral tank and center fuselage tank contains1240 gallons of anhydrous ammonia.Suppression pressure supplied by stored helium.

AUXILIARY POWER UNITTwo RMI X50API: modified for hydrogen peroxide fuel,

HYDRA ULICTwo 8 gpm 3000 psi pumps supply independentsystems powering flightcontrols, speed brakes, andlanding flaps.

ALIGHTING GEARManually released free-fall extension assistedby air drag and b\lngee springs.Skid-type main gear; dual nose wheels

COCKPIT ENVIRONMENTLiquid nitrogen provides refrigeration and pressure,2-1/2 liter liquid oxygen breathing system.

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WEIGHT SUMMARY

WEIGHT EMPTY ... 0 0 . 0 0 0 0 0 0 0 0 . 0 . . 0 . 0 . 0 0 . 0 0 0 . 0 0 . . . . . . . 9,959 LBUSEFUL LOAD 0 . . 0 0 0 00 0 0 0 0 0 0 0 . 0 . 0 0 0 0 0 0 0 0 . 0 0 0 . . . . . . . .. 17, 763 LBPROPELLANT . . . . . . . . . . . 0 0 0 . 0 0 0 0 . 0 0 0 0 0 0 . 0 . 0 . 0 0 0 0 0 0 . . . 16,410 LB

FUEL (AMMONIA) . . . . 0 0 0 0 0 0 0 . . 0 . 0 . . 0 0 0 . 0 0 . . . . . . 7,240 LBOXIDIZER (LIQUID OXYGEN) . . . 0 0 0 0 . 0 . 0 0 0 0 . 0 . 0 0 0 . . . 9, 170 LB

HYDROGEN PEROXIDE . . 00 . . 00 00 . 0 . . 0 . 0 000 . . . . . . . 0 . . . . . . .. 835 LBPILOT. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . It. . . . 0 . . It. 290 LBMISCELLANEOUS 0 . . 0 0 0 0 0 0 0 0 0 0 0 0 . . 0 0 0 0 0 0 0 0 0 0 0 0 0 0 0 . 0 0 0 0 0 o. 228 LB

LAUNCHING WEIGHT 0. 00. . .. o. . . .0. . o. . . . . . .-. . 27,722 LBBURN 83 0 UT WEIGHT . . 0 0 . 0 0 0 0 0 . 0 . 0 0 0 0 0 0 0 0 0 . 0 0 0 . . 0 DO' 10, 443 LBSTRUCTURAL DESIGN WEIGHT (7.33 G) 0. . .00. .0.0. . 10,443 LB

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RESEARCH SUITABiliTY

LJ~-~The proposed airplane is considered to be a re-search tool, and as such has many outstandingfeatures. The effectiveness of this tool can bemeasured by its performance, reliability, safety,and ver satility.The proposed airplane equals or exceeds thespecified performance in all respects. Thisperformance is attained without recourse to un-tested or complicated solutions to design problems.This should allow the major effort to be expendedin obtaining the desired research information. Anyunconventional solution would impose the risk ofdiverting a major portion of research effort toitself and thus seriously penalize the programperformance.Reliability is achieved by simplicity and con-venient elimination of points where failure mightoccur. The proposed airplane is simple in con-cept and, with the benefit of the contractors ex-

perience and effort in detail design, wIll be areliable airplane. North American Aviation hasmade an effort to acquire and use the experiencealready gained by the NACA on research aircraftand will continue to do so in the future.Versatility is achieved by providing an inherentcapability of flying a large variety of flight pro-files, and by designing to allow convenient testingof a variety of shapes and structural configura-tions at the fuselage nose and wing leading edges.Removable wing tips can be easily provided toallow panel structures and aerodynamic shapesto be tested economically.Stability and control problems, at the extremehigh altitudes of the desired performance spectrum,may be studied through the control effectivenessafforded by special nonaerodynamic reaction-type attitude controls.

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~THERMAL PROBLEM

Major problems to be approached and solved bythe use of the X-15 will be those material andstructural problems caused by the high tempera-tures encountered in the very high speed portionsof experimental flights. At a Mach number of 7,the boundary layer recovery temperature will beon the order of 4399F and the skin equilibriumtemperature, where heat input is balanced byradiation output, will exceed 1200F even at alti-tudes above 100,000 feet. 1200F approaches theupper limit for usage of Inconel X as a structuralmaterial.The time rate of approach of equilibrium tempera-tures to boundary layer temperatures allow the useof the airplane at"these high Mach numbers if flightduration is low and the skins are thick enough toform a heat sink of sufficient capacity.High temperatures will affect not only the staticproperties of structural materials, but also theallowable stresses under repeated loads. Lossin stiffness, or Young's modulus, is detrimentalfrom the flutter standpoint.The unequal deflections, caused by dissimilarmetals heated to the same temperatures or by un-equal heating of a homogeneous structure, resultin high order induced stresses built up betweenexpanding and contracting portions of the structure.This effect is very important in considering thedistortions set up by unequal heating between a hotskin, and cool inner structure, or between upper

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and lower skins on the wing during high-angle-of-attack flight. Widespread buckling of skins whichare restrained by stringers or longerons, warpingand curling of wing chord planes, large wing tipdeflections, changes in wing camber, gross wingflutter, or localized panel flutter may be the symp-toms of these thermal stresses.The cootractor's proposed design approach to thesethermal stress problems includes:1. Use of materials having different tensionmodulii where steep thermal gradients ex-ist between adjacent attached members.2. Use of corrugations of beads to provide a loweffective tension modulus.3. Use of thick skin structure with minimum

internal framing where feasible.Wing and tail leading edges and the fuselage nosemay experience temperatures approaching adiabaticstagnation maximums on the order of 4800F dur-ing extreme test conditions. It is proposed to pro-vide easily replaceable leading edges of laminatedglass cloth which will be permitted to actually meltor burn locally during these extreme cases. Theleading edges will be segmented for easy replace-ment. Tests indicate good dimensional stability,low thermal conductivity and stable decomposition,for this type of material.Thus it can be seeI;l that the problems of high tem-

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perature will form a major field of research forthe X-15 test program. The following pages de-scribe a few design features based on presentknowledge of these problems.The high temperatures experienced during flightalso pose the problem of furnishing a proper en-vironment for the pilot and sensitive equipmentwhose temperature and pressure limits are rela-tively low. The range of ambient temperaturesand near-vacuum pressures coupled with the de-sign for a simple reliable conditioning systemrequiring little development work make the optimumdesign of the environmental control a difficultone. The lack of any convenient source of largequantities of either compressed air or ram coolingair such as is associated with the conventional jetaircraft, requires that a new and different approachbe taken to the solution of pressurization and cool-ing. The uniquely simple conditioning system forthe X-15 is comprised of a liquid nitrogen storage

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tank plus the required distribution plumbing,regulating valves etc. The cold nitrogen liquidplus the available heat absorption inherent invaporization of the fluid, form the necessary heatsink for refrigeration. In addition the resultinggaseous nitr~en serves as atmosphere and a pres-surizing agent for the cabin, equipment compart-ments, and other localized heat sensitive areas.The pilot's pressure altitude suit is also pres-surized with nitrogen, his oxygen requIrementsbeing supplied 100 percent from a liquid oxygensource through an inner breathing mask. Thenitrogen atmosphere in the equipment compart-ments also tends to lessen the possibility of fireor explosion. Another desirable asset resultingfrom proper environmental control is the abilityto use presently available equipment instead ofhaving to laboriously develop new high temperatureand low pressure resistant instruments and equip-ment.

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THERMAL SUITABILITY.. . . wIngThe basic wing is fabricated as a complete full...span assembly for best rigidity and lightness ofstructure. Wing skin loads are transferred intofuselage ring frames for transfer across the fuse-lage. The ring frames are of titanium alloy withnumerous webbeads to minimize thermal stresses.The wing structural box extends from the 25 per-cent to 75 percent chord lines. A spanwise seriesofsrear beams, compJsed of corrugated 24S-T websand titanium-manganese alloy attach angles, providesupport for the taper milled Inconel X skins. Thesplr corrugations resist normal crushing loads andserve to relieve thermal stresses. The relativelylower modulus of elasticity of the titanium-man-ganese attach angles will reduce thermal stressesinduced from the hot Inconel X skins. Thicknessof the skins varies from O.060 inch at the tip toO.125 inch at the fuselage fairing intersection.Front and rear main spars are built up fromtitanium-manganese alloy with numerous beadsfor relief of thermal stresses. Panel stabilitywill be investigated in tests conducted in environ-mental circumstances.The airfoil shape has a rounded leading edge anda conventional cross section for low drag, reducedleading edge temperatures, and good subsoniccharacteristics. The leading edge section with achord of 3-1/2 inches throughout, is made of lami-

nated Fiberglas. The most severe flight profilewill result in about 5/8 inch of material be-ing burned away, but since the leading edge isseparated into 24-inch lengths, the damaged sec-tions may be cheaply and easily replaced. TheInconel X skin increases in thickness chord-wise from the front spar to the leading edge toform a heat sink.The square wing tip allows the tip rib to be at thesame temperature as the skins. The rib at thefuselage juncture is titanium-manganese alloy,while the mid-span and nose ribs are corrugated24S-T with titanium-manganese attach angles.Tail surfaces, in general, are constructed in thesame manner as described for the wing.

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4fII!(- . NA-55-221,THERMAL SUITABILITY

. . . fuselageThe necessity for use of moderately heavy skinsfor heat sink, use of a shape compatible with thepressures required within the integral tanks, andconsideration of thermal stress prqblems has ledto adoption of semi-monocoque fuselage with in-ternal structure consisting only of rings or bulk-heads except in local areas around the cockpit.andempennage. This method of construction allowsthe thermal deflections of the monocoque shell totake place without the crippling of conventionallongerons and stringers which would result fromtemperature gradients. The shell is stabilizedby bulkheads spaced approximately 25 inchesapart. The bulkheads are stiffened with a seriesof radial beads to reduce thermal stresses. Themonocoque shell is fabricated from Inconel X sheetmetal for high-temperature strength. Internalpressures are considered to assist in stabilizingthe monocoque shell.Side fuselage fairings housing controls, hydrauliclines, etc, are made from Inconel X and seg-mented every 20 inches to reduce thermal deflec-tions and stresses. Primary structure behind thefairings is beaded for relief of thermal stresses.Panel stability will be investigated through en-vironmental tests.

--The Inconel X nose cone forward of the cock-pit is tapered in thickness from O.060 to O.100 inchat a point 8 inches from the tip. A replaceablelaminated glass cloth fairing forms the foremostnose tip, which at the temperatures attained dur-ing the most severe flight profile, will partiallyburn away at the surface, thus protecting the re-mainder of the nose.All bulkheads and frames, with the exception oftank bulkheads, are constructed from titanium-manganese alloy. The lower modulus of elasticityof this material reduces the magnitude of the in-duced thermal stresses, yet working allowablestresses are high. The use of service tunnelsin the outer fairings minimizes the need for cut-outs in the bulkheads for wire and plumbing routing.

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THERMAL SUITABILITY. . . cabin & equipment bay

The cabin and forward equipment bay utilize low-temperature, inner shells to retain compartmentpressure. Conventional sealing techniques areused. Insulation between the inner shell and thestructural skin comprises one inch of glass fiberplus aluminum foil radiation shielding.The canopy seal is isolated, by distance, from thehot skins, permitting use of a conventional "blow-up" seal operated by stored nitrogen.Liquid nitrogen storage systems provide a pres-surized and refrigerated atmosphere.The windshield consists of heavy fused silica or

Pyrex outer panes and stretched acrylic innerpanes. The inner low-temperature panes pro-vide the normal pressure seal. All panes aregeometrically flat to simplify fabrication.A stabilized ejection seat and the pilot's full pres-sure suit combine to provide an escape systemhaving minimum weight and complexity, environ-ment control, and a probability of surviv~l whichis believed to exceed that of any capsule of accep-table weight which could be developed within theallowable time period. Since the design dynamicpressures are inthe same range as those of present-day airplanes the seat should clear the airplanesatisfactorily under all conditions.

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THERMAL SUITABILITY. .. . remote equipment & routing

All fluid lines, controls, and wiring extendinf.along the fuselage are routed outside of themonocoque fuselage structure and within the re-movable side twmels. In the region. of the integraltanks, insulation will be provided between the tankwall and the tunnel to afford protection from theextremely low temperature of the propellants.Insulation is provided adjacent to the inner sur-face of the tunnel walls subjected to aerodynamicheating. All propellant lines are provided withinsulation to alleviate the low-temperature problem.All wing routing, excepting terminal portions ofthe research instrumentation, is within a thermallyinsulated portion of the wing structural box, aft ofthe front beam. Access is obtained by removinglower surface access doors. Terminal portionsof the research instrumentation will be servicedby suitably located access panels on the lower wingskin.Remote equipment which is not capable of with-standing ambient conditions will be suitably pro-tected. Items requiring pressurization and cool-ing will be installed in thermally insulated pressure

cans and, if necessary, pressurized and cooledwith the liquid nitrogen cabin pressurizing and re-frigeration system. All bearings subjected to hightemperatures will be provided with sufficient heatsink to allow use of dry lubricants such as molyb-denum-sulfide, or be designed to operate withoutlubrication.

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STABILITY& CONTROL PROBLEMS800

7001. SPACEREGIONAerodynamiccontrol power is notavailable.

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SUPERSONIC REGIONLoss of lifting effectiveness of thin airfoilsurf aces r esults in reduced stability and control.

3 . TRANSONIC AND SUBSONIC REGION.Excell en t g lide, approach and landing condit ionsmus t be prov ided.

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DYNAMICPRESSURE, q - LB/FT21200 1400 1600

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The stability and control of a vehicle moving atvery highvelocities in an atmosphere of lowdensitywill be one of the principal fields of investigationof the X-15. At the present time sufficient data isavailable from wind tunnel and model tests to indi-cate that the problems to be encountered will notbe- insurmountable. One major effect will be thereduction in stability at very high Mach numbers,and the total loss of stability and control effective-ness at altitudes where dynamic pressures fall tovery low value's. The proposed solution to the for-mer problem will be provision of symmetricallysplit speed brakes which can transform the basicdouble-wedge vertical stabilizer airfoil trailingedge into a relatively obtuse blunt wedge. Thisconfiguration greatly increases the tail lift curveslope at high Mach numbers, and thus providessufficient directional stability without actual in-crease of tail area. The relative location of air-plane CG and supersonic aerodynamic center ofthe wing overcom es the trend to longitudinalinstability so that a conventional horizontal tailairfoil, deflected through large angles, is suf-ficientlyeffective. The problem of control dur-

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ing space flight will be solved by the use of reac-tion jets arranged to provide control momentsabout the three aircraft principal axes. The pri-mary use of this space control will be to orientthe airplane attitude prior to re-entry into theatmosphere. Optimum control forces have beendetermined from an analog computer analysis.The configuration of the X-15 will allow a wideflexibility in the types of re-entry trajectories thatmay be followed. The aircraft may re-enter on azero lift curve, using the speed brakes for decelera-tion, and with subsequent pull up at reduced Machnumbers. Temperature differentials across thewing thickness will be at a minimum in this maneu-ver. At the opposite extreme a high lift coefficientand resultant high drag may be maintained for de-celeration without the speed brakes, and yet exces-sively high load factors or bottom skin temperatureswill not be encountered. This flexibility of flightpath programming will permit a much wider rangeof conditions to be investigated during the testprogram, and provide a wide safety margin tocover unplanned deviations.

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AIRFRAME SERVICE SYSTEMS. . . cabin & equipment conditioning

COMPARTMENT PRESSURE REGULATOR

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INSULA TIONPRESSURE SENSING ELEMENTTEMPERA TURE SENSING ELEMENT

SAFETY VALVE TANK PRESSURE REGULATOR

NITROGEN MODULATING VALVECOMP ARTMENT PRESSURE AND TEMPERA TURE CONTROL

EQUIPMENT CONDITIONING SCHEMATIC

A liquid nitrogen storage system is used for bothpressurization and refrigeration of the cockpit,forward compartment, midship equipment, andleft and right auxiliary power unit compartments.Four insulated tanks containing a total of 75 poundsof liquid nitrogen are self-pressurized. The nitro-gen will be injected into the conditioned areas; cool-ing results from the rapid evaporation of the spray,and pressurization is automatically kept at theproper level by relief valves.

Tanks will be filled with liquid nitrogen on theground prior to take-off.The pilot is protected by a nitrogen pressurizedaltitude suit. A 100 percent oxygen supply system,with regulator attached to the suit, is supplied froma 2-1/2 liter liquid oxygen supply converter. In theevent of failure of the cabin nitrogen pressuriza-tion, the oxygen system will serve to pressurizethe pilot's altitude suit to ensure survival. Agaseous oxygen bail-out bottle is provided.

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AIRFRAME SERVICE SYSTEMS. . . auxiliary power units.....

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P ower for operation of the airplane electrical andhydraulic systems is furnished by two RMIX50AP-1gas turbine units modified to utilize hydrogenperoxide as a monopropellant. Each unit drivesone 100-ampere 30-volt d-c generator through a 2:1gear boxon the 12,600 rpm pad and one 8 gpm, 3000psi variable-displacement hydraulic pump mountedon the 3850 rpm pad. Each unit is completelyautomatic except for mechanically controlled pro-pellant valves. The dual installation ensures sys-tem reliability in case of a power unit failure. Thepower units are located in the aft fuselage and areaccessible through the engine cowlings. Hydraulicline runs are thus kept at a minimum length.Propellant supply is from two pressurized bladdertype tanks located in the aft fuselage. The spacecontrol system is also supplied from these tanks

with 68. 5 percent of the total capacity of 10 gallonsbeing allocated to the auxiliary power units. Pres-surization is at 400 psi through a regulator con-nected to a separate 3000 psi helium bottle.Propellant capacity is sufficient for a 5-minuteprelaunch warm-up, plus continuous operationduring a maximum duration flight.A H202 propellant fill and return line and valve,with connections located in a common access portwith the main engine turbo-pump peroxide filler,are provided for closed-circuit propellant servicing.Propellant tanks, lines, valves, and catalystchambers are insulated sufficiently to preventfreezing in -60F ambient temperatures as well asto prevent boiling or decomposition during periodsof high-speed, high-temperature flight.

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AIRFRAME SERVICE SYSTEMS. . . electrical

The electrical system is of the regulated 24-voltto 28-'volt d-c type powered by two 100-ampereE -1601-1 (or equivalent) generators, each drivenfrom one of the gas turbine auxiliary power units.The generators are installed in pressurized andrefrigerated containers which provide the requiredenvironmental isolation. Each generator isequipped with a carbon-pile voltage regulator,reverse-current cutout, cockpit warning light, andon-off switch. Normal operation is in parallel,feeding both a primary and an essential bus sys-tem. In case of malfunction, the remaining unitwill automatically continue to service the essentialbus and will disconnect the primary bus. Allequipment considered necessary for flight and,landing is on the essential bus. Energization ofboth busses during captive flight is through anexternal receptacle connected to the mother air-plane. Pilot's controls consist of generator warn-

ing lights, generator on-off switches, and an"internal- external" switch.A secondary 115-volt, three-phase a-c systemis powered by a 750-volt-ampere main inverter.Malfunction of this unit will activate a warninglight, and the a-c load may be transferred to theresearch instrumentation inverter, with consequentdropping of the instrumentation load.The use of two generators with environmentalprotection gives dual unit protection that is notduration-sensitive since the turbine fuel supplyis based on the full system capacity demands.The normal need for only a portion of the usablegenerator capacity permits flexibility and growthof equipment requirements, without electrical re-design, as the flight program progresses.

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AIRFRAME SERVICE SYSTEMS. . . flight controls

--------.The aerodynamic control of airplane attitude is bymeans of all-movable horizontal tail surfacesoperated symmetrically for pitch, and differentiallyfor roll, plus an all-movable upper vertical sur-face for directional control. Dual tandem hydraulicactuators powered by twin independent hydraulicsystems provide irreversible control actuation.The conventional cockpit control stick is connectedto the servo valves by cable runs down each sideof the fuselage. A nonlinear pitching mechanicaladvantage increases stick travel per degree ofstabilizer travel in the leading edge up range andreduces it in the leading edge down range. Thisreduces the variation of stick force per G producedby the feel bungee and any tendency for pilot-induced pitching oscillation. A trim button onthe stick controls an electrically actuated bungee

trimmer, and a small bobweight contributes tostick feel. Roll and directional feel is providedby manually trimmed bungee springs.Split speed brakes on upper and lower verticalsurfaces are individually controlled by separatecockpit levers which are cable-connected to theservo valves. A pre-select type control allowsthe brakes to follow upat rates proportional to the -air loads.Wing flaps of the plain type are operated hydrau-lically and are controlled by an electric switch inthe cockpit and an electric servo actuator-synchro-nizer connected by flex shafting to the individualflap servo valves. A solenoid valve cuts off pumppressure to the servo valves when the flaps areup and locked.

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AIRFRAME SERVICE SYSTEMS. . . space controls

A separate space control system is provided forattitude control during periods of ineffectivenessof the normal aerodynamic controls due to reduceddynamic pressures. The system comprises a setof small variable-thrust hydrogen peroxide rocketmotors arranged to provide variable momentsabout the three control axes. Four motors arelocated in a cruciform arrangement at the nosefor pitch and yaw, and one motor at each wing tipfor roll control.The rocket motors are similar to the RMI XLR-32-RM-2 type. Hydrogen peroxide is suppliedto each pair of control motors by a two-way

throttling valve connected by mechanical linkagesto a three-mode maneuvering lever in the cockpitoperated by the pilot. The lever is spring-loadedand normally neutral so that break-out forces andspring rates give a realistic feel, and resultingfuel flows and control moments are proportionalto lever deflection.Control operation is of the simple on-off no-feedback type, with the initial lever movementstarting a certain acceleration about any axis,followed by a constant-velocity rotation with thelever in neutral, and ending with an oppositecontrol application to damp out the rotational

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velocity to zero as the desired heading is reached. freezing during flight at -60F and to preventboiling or decomposition temperatures beingreached during high-speed flight. A maximumrotational acceleration of 2. 5 degrees per secondper second about any axis may be attained. Nor-mal duty cycles will ensure five gross attitudechanges about each axis at approximately 6 de-grees per second.

The hydrogen peroxide monopropellant is suppliedfrom two pressurized tanks, which also furnishfuel for the auxiliary power unit. The tanks, lo-cated in the aft fuselage, are of the positive-expulsion bladder type pressurized to 400 psi bya 3000 psi helium bottle through a pressure regu-lator. Total capacity is 10 gallons, with 3.15gallons allocated to the control system. Auto-matic 50 psi check valves are located at each motor.A propellant quality gage is provided also. Suf-ficient thermal insulation is provided for the tanks,valves, lines, and catalyst chambers to prevent

The flexibility inherent in using a common fuelsupply for both space control and auxiliary poweris consonant with the exploratory nature of theintended aircraft research program.

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PO.WER PLANT. . . rocket installation

The research airplane is powered by a single-chamber liquid propellant variable-thrust rocketmotor. The unit is a Reaction Motors, Inc.XLR30-RM-2 motor burning liquid oxygen andanhydruus ammonia. Maximum sea-level thrustis 50,000 pounds, and the thrust may be variedfrom 30% to full thrust by controlling power out-put of the propellant pump turbine. This turbineoperates on superheated steam furnished by de-composition of 90 percent concentration hydrogenperoxide in a gas generator.

Throttling of the steam turbine input flow variesthe propellant flow, and consequently the rocketthrust.Starting is accomplished by utilizing the hypergolicreaction of hydrogen peroxide and ammonia in therocket chamber prior to admittance of the mainoxidizer. No electric ignition is required. Theairplane helium pressurizing system is of sufficientcapacity to allow three starts and shutdowns withfull engine purging.

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POWER. PLANT. . . propellant system

The main propellants are liquid oxygen and anhy-drous ammonia, which are supplied under pressureto the suction side of the propellant pump. Thepropellant pump turbine is supplied with super-heated steam formed from 90 percent hydrogenperoxide in.a catalyst chamber.Pump inlet pressure for the main propellant sys-tem is supplied by helium regulated at approxi-mately 45 psi and stored at 3000 psi and approxi-mately -300F in a cylindrical tank on the center-line of the liquid oxygen tank. A similar tanksupplies pressure for the propellant pump peroxidesystem, the engine, and system controls. The

two pressurization systems are separated to reducesize and weight, since the main propellant systemmay operate to a much lower exhausted pressurethan the engine and control system.The liquid oxygen tank comprises a section of theforward fuselage structure and is constructed ofseam-welded lnconel X. The aft tank, of similarconstruction, contains ammonia. A center sectionnonstructural tank provides the remaining am-monia capacity. All tanks comprise intercon-nected compartments arranged to pr9vide auto-matic sequential propellant flow for CG controlduring normal or jettison operation.

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ALIGHTING GEAR

The main gear consists of two strut-mountedskids that fit flush to the fuselage when retracted.Gear extension is initiated by manual release ofthe uplock. Gravity and aerodynamic loads extendthe gear and a bungee spring insures proper posi-tioning of the drag linkage and actuation of the downlock. The skid is properly positioned for groundcontact by a bungee system. The ground crewretracts the gear by manually unlocking the dragbrace and rotating the strut and skid up againstthe fuselage.The nose gear comprises dual wheels mountedon a retractable air-oil strut. Manual groundretraction of the gear is done by releasing thedown-lock, collapsing the strut by means of anexternal air supply, and rotating the gear into thewheel well by hand. Extension is started by manual

release of the uplock, followed by an initial rota-tion caused by the expanding air-oil strut operatingagainst a cam, and then by gravity and air streamloads which bring the gear to the down-locked posi-tion. A fairing attached to the strut fully enclosesthe gear when retracted.A manually retractable skid in the lower verticaltail is released by the pilot and is extended bygravity and a spring bungee.The preferred method of landing will probably beto touch down tail-first, with a subsequent pitch-over onto the main skids and nose gear for grounddeceleration and stopping. The wide tread of theskids and the castering ability of the nose wheelwill minimize the tendency for directional oscil-lation while decelerating.

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INSTRUMENT ATION

Selection of this equipment has been guided by thebasic principles of safety, ability to follow a pro-grammed trajectory, and minimum need for de-velopment. All instrumentation is standard GFAEexcept the Air Data System, which is defined inNAA Report NA-55-231 included in this proposal.A uhf ANI ARC-34 command transceiver will beused for voice communication. This unit providesmore functions than are necessary, so the feasi-bility of using a smaller, more compact set oflower power (such as the AN/ARC-45) will be in-vestigated.Lateral steering during the initial portions of the

flight will be principally aided by the use of a J-4gyrocompass, while the later part of the mission,including the approach to the landing field, will befacilitated by the use of an ANI ARA-22 radio hom-ing group. This unit time-shares a portion of theuhf command set circuits.At altitudes above the limits of the air data unit,airplane attitude and orientation will be determinedby a K-4B remote-indicating vertical gyro. Thisinstrument does not tumble and has a very low rateof drift. The pilot will need this instrument whileleveling the wings and adjusting pitch attitude priorto re-entry into the atmosphere. A panel-mounted

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accelerometer will aid in accurate determinationof the desired climb paths subsequent to launching.The pilot's immooiate reactions and comments uponincidents during the course of a flight will be re-corded on an AN/ ANH-3 tape recorder for laterplayback and analysis.A multiple-purpose air data system, defined inNAAReport NA-55-231, will be designed to fur-nish certain information. Development time forthis system will be minimized, since many of thecomponents have already been successfully de-veloped. Pitot-static pressures, differential dy-namic pressures due to angle of attack and angleof skid, and air stream temperatures will be fedinto a data computer, and outputs illustrating Machnumber, indicated airspeed, rate of climb, baro-metric altitude, angles of attack, and skid will bepresented to the pilot for his evaluation. Compen-sation for the total pressure errors at high Machnumbers \yill be applied to the observed pressurereadings. The present-day NASEC (Static ErrorCompensation) device may be easily extended intothe range of this test airplane. Conventional rate-of-climb instruments are too sluggish for a craftof this type, but a fast-acting sensitive instrumentcalled NARODI(Rate of Descent Indicator), designedfor missiles, may be adapted. A usable accuracyshould still be available at a 100, OOO-footaltitude.Accuracy is very good at low altitudes, and it will

aid in determining the rate of descent during alanding. The Machmeter will have limits of O.7and 7. 0 Mach number.Re-entry into the atmosphere at very high speedswill result in great importance being placed on theknowledge of airplane angle of attack and skid. Thesensitive differential pressure transducers willfortunately allow the detection of these angles about30 seconds before the dynamic air pressures riseto significant values. Optimum boom shapes tomaximize this sensitivity will be the subject offurther research. Both angles will be presentedon a single cross-pointer indicator for easy reading.The specified research instrumentation equipmentwill be installed in the airplane in a compartmentcooled and pressurized by means of a lightweightliquid nitrogen system. Components will includeautomatic recorders for airspeed, altitude, three-axis linear accelerations, angular velocities andaccelerations, airplane attitude, airflow directions,control positions and forces, trim settings, hingemoments, structural temperatures and stresses,and stagnation temperature. Automatic cameraswill record visual deformations of wing and tailsurfaces. A multichannel telemeter will relayengine operational data and cosmic radiation in-formation to the ground. An AN/APN-65 radarbeacon transpondor will allow accurate. groundtracking and ranging of the airplane during flight.

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~C"':::c1"'=c:,.-1"'

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PROGRAM SCHEDULE

MONTHS AFTER GO AHEADENGINEERING DESIGNMOCKUPMATERIALS TESTINGWIND TUNNEL TESTING

HYPERSONICSUPERSONICLOW SPEEDSPIN

MANUFACTURINGFUEL SYSTEM TEST STANDGFAE ENGINEFLIGHT TESTSTATIC TESTAIRP LANE DOLLYENGINE DOLLYMISC. GROUND EQUIPT.

SEPT 1955GO AHEAD

COMPLETION &SUBMITT AL OFREPORTS

:~ ~ ~ ~ ~ m m

::::::::::::::::::::::;:::t:::::::::{{:

.1" ... ...T'T.'~TT'~.~. ",cm... .FIT .T ..m.......................': : : : : ~ ~ ~~ ~ : : : : : : : : : : : : : : : ::::

5 0 0 1 0\0 1 0 0 1 0 1 0 1 0

:::;:::::::::::::::::::::1:::::::::::::::::::::::::::::::::::::::::::::::::::::: tlJAA I:::::;::::::)h::::::::::::::::jNGINE MFG. TESTINGFUEL TESTNGINEERING RELEASE:~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~ ~~ J~ ~ ~ ~ ~ ~ } ~ f

:::::k::::::~:~:~:~:~:::~n:::::::::::::::::;:::::::::::::::::::::::~:~J:I

;::::::::tm}~l}}}~:~:~:~:~:~:~:~:~:~:~:~:~:::~:~:~:

100100\1

1 0 0 1\0 0

1001\001

NOTE: 1. Contractor modified B-36 Carrier Airp lane to be available for telt program by Dec. 195'1.Carr ier to be f lown and maintained by USAF personnel during Contractor 'l f light tel t program.

2. GF' engine ""n "and and loadlug~able byDeoemb", 1957. 26 --


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NORTH AMERICAN WIND TUNNELS. . . type & availability1. NORTH AMERICAN AERODYNAMICS

LABORATORY 3. SOUTHERN CALIFORNIA COOPERATIVEWINDTUNNEL

(NORTH AMERICANHAS1/6 PART OWNERSHIP)MACH RANGE: .2WIND TUNNEL SIZE: 7.75 FT x 11 FTMODELSIZE: 7 FTAVAILABLE: NOW

MACH RANGE:WIND TUNNEL SIZE:MODEL SIZE:AVAILABLE:

.2 TO 1.808.5 FT x 12 FT5 FTSUBSONIC& TRANSONIC, NOV 55SUPERSONIC , FEB56

2. SUPERSONIC AEROPHYSICS LABORATORY 4. NORTH AMERICAN TRISONIC

MACH RANGE: .7,1.22,1.56,1.872.48, 2.87, 3.24WIND TUNNEL SIZE: 16 INCH x 16 INCHMODELSIZE: 10 INCHAVAILABLE: NOW

MACH RANGE: .2 - 3.50WIND TUNNELSIZE: 7 FT x 7 FTMODELSIZE: 5 FTAVAILABLE: MAY 1956

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DESIGN MISSIONSaltitude and speed vs. time. .- DESIGN PROFILE

I 1_-- HIGH TEMPERATURE PROFILE

PULLUP PULLUP CLIMB TO 75,000 FTPOWEREDCLIMB TO BURNOUT, CL=O - -2 G PUSHOVERPOWEROFF CLIMB TO POWEROFF LEVELFLIGHTMAX. ALTITUDE, CL MAX. i DESCENTTO 10,000 FT. OPT. GLIDERE-ENTRY& PULLOUT CL=.75DESCENT TO 10,000 FT OPT. GLIDE--- -

. BURNOUT100 200 300 400 500 600 700 800TIME FROM LAUNCH - SECONDS 900 1000 1100 1200

7000~ 6000~ 5000~~ 4000Q 3000~ 20000-4rn 1000

o o 100 200 300 400 500 600 700 800TIME FROM LAUNCH - SECONDS 900 1000 1100 1200

28 .

I I I .1. I I" - DESIGN PROFILE - - - HIG H TEMPERATURE PROFILE . BURNOUTJ

.........,- -'" '\I " '\.I \ -,- .....-I/ ............... -,

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1400

-200 o

DESIGN MISSIONS. .. temperature V5. t.me

o50 100 500 55050 200 250 300 350 400

TIME FROMLAUNCH- SECONDS450 600

29

12000

1000

800p:;600

Z1-4 400r:t.:I

200

I I I,- ......... - - - HIGH TEMPERATURE MISSION; DESIGN MISSION.' ", "I r FUSELAGETATION150, BOTTOMI ', .06 SKIN THICKNESSI ""-. '..... " - WING LOWER SURFACE 5'10CHORDI

/..........12 SKIN THICKNESSI '..

I .,--"---- --- .....I " -' ---.... .......I -.......' -- --- --- '...-- .....I/1!,... ",..- '\ --- ......,, ,-' ---....., / , r:USELAGE STATION100, BOTTOM .. .... / . 06 SKIN THICKNESS. '/ I1/ '-- WING LOWERSURFACE25'10CHORDI / - .08 SKIN THICKNESS

I \.

AY \ - WING LOWER SURFACE 5'10 CHORD'" . 12 SKIN THICKNESS

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PERFORMANCE CAPABILITIES900

800

600

500

400

200100

o4000 4800 5200 5600 6000 6400

VELOCITY AT BURNOUT - FEET/SECOND6800 7200400

30

E-t700

000..-4I

E-t.....E-tH

X-15 Design Proposal

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Transcript of X-15 Design Proposal