Created: 1/1/1968

OCR scan of the original document, errors are possible


in .ni

air coitoitiohino (continued)

Prcaaurlaatlon {Continued)

ebla Rate of Climbhe cabin altitude changes at constant rates aet by the pilot to carry the cabin from Initial to

final altitude in the exact time spans of tbe airplane's climb and descent.


a cabin differentialal ia reached at initial crulae altitude, aad remains constant during cruise, resulting again5 ft. cabin at altitude. However, this system bas the inherent peculiarity of requiring oabln differential toighel on the way to or from cruise altitude. It was selected for study as requiring the mrima pressurlsatlon nitrogen of any practical schedule,ounds for the mission flight noted shore. Pilot comfort during desoent, on the other hand, is by far the best of any poaslble system, aa Indicated by the cabin pressure rate changespmnot descent and onlypm at JjOO knots. The latter rat* compares to the above0 fpm maximum for the lsobarlc system, and0 fpm reached with the military-type system below.

Military-Typenpressuriaed ram operation from sea-levelt. oabln altitude is reached,o baric pressure is heldt. until cabin differential has built opel, withsl differential at all higher airplane altitudes5ince the first two ayatoras above spanned the mission flight nitrogenfrom minimum to maximum, this more normal system's nitrogen usage

im/ aircraft corporation


A.IH SQKDmOHffiO (Continued)

Presmirlaatton (Continued)

Military-Type Schedule (Continued) was Investigated for descent only. Herenot descent nitrogen amounted toounds, compared toouoda and UO pounda on'the lso-barlo and constant rate systems, respectively, for the same descent. LCO knot descent nitrogen was also calculated for thie system,toounds. The mala use made of this particular schedule, however, waa In investigating the relative pilot comfort between lt and the Isobarlc schedule, during the JaOO knotinutes). Oa the military-type schedule, after leaving mart mum orulss altitude at time aero, tbo pilot epeode theeconds subjected to cabin descent rates varying55 inohoo of mercury/minute, whereas, with the lsobarlo schedule no cabin change whatsoever occurs during this time span. For the remainingeconds both of these schedules would follow the earns rate curve (approachingaches of mercury/minute atxoept that ateconds the pilot with the military-type systemalf minute of reprieve at aero rate int. isobarlc cabin, then returning to the high rate for tho finaleconds. This comparison shows both systems to be relatively severe on the pilot, such that ho should not attemptapid 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 isobarlc system


ircraft corporation AIR CONDITION PJQ (Continued)

Prossurlzatlon (Continued)

Hllltary-Typ* Scheduleligbt edgetha military-type, on th* basis of sustained times at high rata of chaagai thia could probablyoint of argument between eny two given pilots, however.

For the present, no attempt is being made to finalize th* preseuri-zation schedule, but merely to have at hand tbe information on which the discussions above were based. This is necessarily the case because of the lnter-relatlonship between the nitrogen required for prossurlcation, and that required for cooling. It is felt that until final decision on the cooling aysten 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 pressurlzatlon can be stated at present; however,howsa simplified general concept covering theontrolling that must be accomplished. The master controller, whatever Its form, must accept signals from both the temperature and the pressure sensors, and than 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


AIRContinued) Pressurigatlon (Continued)

Mllltajy-Type Schedule (Continued) does not point up the necessity for such an inter-relating Master control, ainoe it la obvloualyormal functionreaeure sensor tocontrol Its outflow valve towards open forase. However, reversing the example, assuming sufficient decrease in cooling-nitrogen inflow toirect-controlled outflow valve fully closed, wouldin depreseuritation. (The outflow valve by ltaalf cannot "pump up* the cabin, being capable only ofigher preaeure generated at or beyond tlie point of cabin inflow. Note the diasimllarity between the more normal case of an "infinite" bleed air source availableable, and the present "release lt only as you need it" source). Now the need for the master becomes more evident, since it must recognise that even though temperature-wise the nitrogen flow oan be reduced, lt must etlll signal for nitrogen' aa an inflowing pressure source. (Thecontroller at thie time will function only to position thebypass valve or valves).

For tha 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 tbe contained weight of cabin atmosphere


US COMDrnOWlfC (Continued) Preasurltatlon (Continued)

HUltary-Typo Schedule (Continued) during climb and descent). alues for nitrogen,

Quoted under the various schedules sieve, vers calculated on this basis,


Note inhat the series arrangement of outflow and ssfsty valves gives double protection to the pilot against cockpit depreesurUa-tlon. For example. If the equipment bay were to depressuriss for any reason the cockpit remains fully pressurised. As an alternate example, if the cockpit's outflow valve became stuck ln the open position, tbe cockpit would again remain fully pressurised by riding on the equipment bay's valve. The probability of simultaneous-open failures is very low.

With the ram operation proposed for the unpreesurlsed portions of the above described schedules, and the variable pressure source available during pressuriaatlon, lt" Is considered that no negative pressureproblem can normally exist, even during tbe IjOO knot descant. In thia regard calculations were made to determine the required variation between nitrogen flow rates for0 knotnot descents, to maintain full preaaurliation on the military-type schedule. This had beena possible problem on the fast descent fron the standpoint of that nitrogen portion required Just to increase the contained weight of cabin atmosphere. The results show, however, that even though the hOO knot descent time was faeteractor of almostta nitrogen



AIR CONDITION fJQ (Continued) Prossurttatlcn (Continued)

Military-Type Schedule (Oontiouod) discharge rate (including leakage make-up) ma rely doubled. Further In regard to negative differentials, for the caseitrogen 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.


In the early atages ofook waa 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 thisossible solution.

Engine bleed air was peremptorily eliminated for cabin use due to the airplane performance losses associated with bleed at our altitude. Mote, however, that it le planned to use limited amounts of bleed air for windshield defogging and for ram air heating, if required, during the unpreseurlzod portion of the pressurltation schedules discussed above. The latter usage would become especially Important were in-flight refueling to be considered, aa here the normal time of ran operation would be far exceeded.

Tbe most recent investigations have been aimed at accomplishing certain assumed design direct irea as

f AIRCRAFT lifll


Cooling (Continued)

space temperature ln the oookpit and equipment bay,

maximum touch temperature of the trim ln areas not directly overstructure, with minimum possible touch temperatures In all other sraas.

Cooling to be accomplished entirely by liquid media stored aboard the alrplans (nitrogen, and possibly water).

Cooling medium to double as cabLn pressure source, as discussed above under prossurliatlon.

The above temperatures take Into account the fact that the pilot's comfort will be at tbe much more suitable level associated with direot suit ventilation by nitrogen gas, as oni>. This vill include pilot-selected temperature controls.

One of the most attractive features of having aboard liquid nitrogen ia the ease with which spot cooling of critical areas or equipment components can be accompli shed. Thus such local areas are considered to be no problem.

A recirculation system was lnvestlgsted on the basis of the above temperatures, wherein cabin atmosphsre was cooled ln two stagesi first by passage through the air sideater boiler, and then by injecting into lt liquid nitrogen which topped-off the required cooling. This system was considered unattractive weight-wise at the time.




AIR CQirorriOMIllO (Continued)

Cooling (Continued)

The most recent work, Just completed,etailed study madeater-panel system, for handling tne major cooling load in those portions

of the cabin wall between structural rings. Until the final stegea of this


investigation were reached, and the results could be integrated, thia system looked extremely attractive. Regrettably, the final system weight has turned out to far exceed that of much less elaborate systems, even though asthe amount of water expended wae 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 toractical system within the weight allowance set forthin this report. This would be so even for aalone" ay stem, and note in this regard that nitrogen's heat of vaporisation amouota only toenth that of water, at the pressures Involved.

In ensuing lnvcstigatiene lt is intended to exploit still further the advantages of using water's high heat of vaporisation, in combination with top-off cooling by the "double-duty" preSBuriaation nitrogen.




ill be provided with both iow altitude and high altitude eeoape capability. Zero velocity escape systems will be carefully evaluated and used if feaalble.

The pilot's seat will eject upward and will be providedocket oatapult.

This catapult assures adequate ejection clearance to the tail of the- aircraft during all operating regimes. The long nose of the aircraft, which places the pilot far forward of the tall, and the low aspect ratio of the tall allow the man and the seat in an emergency ejection to clear the airplane tallonsiderable margin of safety.

The seatuitable survival kit and bail-out oxygen system, andpecial shoulder harness and lap belt automatic release system.

Emergency ejection at the high altitudes and high speed at which the mission of the aircraft will be performed introduces new aspects to be considered In the problem of escape. The opening shock of the parachute, chute oscillation and the rotation of the pilot's body during free-fall muet be retained within tolerable limits. The heat generated during ejection at cruise Mach number and deceleration of the man have been considered to Insure that they are within body tolerances.


EXEfflBiCI ESCAPE (Continued) Wind blastonsideration during emergency ejection. The srulsa Mach number, although high, correaponda tonots EAS, Maximum EAS to be used will be0 knots which occurs leas thanof the flight time.

Wind blast effects and deceleration at theae Telooitlea arelow and eafe ejections can be made within airspeeds with which we have had experience starting with the first Jet fighters.

Heating of the nan during ejection iserious problem. eet altitude,ith the man in tha seat, deceleration time Is approximatelyeconds. During this period the man and seat

decelerate fromo terminal falling Telocity.


The initial ten seconds of this deceleration is the critical period so far as aerodynamic heating la oonoemed. Afteren eeoonda, tbe ejected pilot has slowed to approximatelyorresponding reduction In temperature. Some Improvement of tbe temperature resistance of face pleoea and pressure suit details la probably required to acoomnodate this transient oondition.

It is proposed to use automatic seat belt and shoulder harness release of the type used ino permit the pilot to remain in the seat until the initial deceleration Is complete. This will reduce tbe basard of pilot injury due to epln and tumbling, and prevent premature deployment of the primary chute.







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PKB33NAL EQPffKEHT In order to be compatible with tbe airplane and emargenoleeescape, it io necessary toersonal equipment development program for this particular airplane. The general nature of this program is to adapt and modify presently proven equipment and to develop aof new equipment as required. artial -pre sau re suit, such as used so successfully inrogram is proposed as the basis for the pilot's wear.

The cockpit temperatures ofrequire that the pilot be provided With some sortooling and ventilation garment.

Lockheed understands that pressure suit development Is being done at Edwards Air Force Base aod proposes to make use of this program to provide2 with the latest advance in full pressure or partial-pressure suits that the state of development permits.

A liquid oxygen system will be installed to provide oxygen forand emergency suit pressure for the pilot.

To reduce the chances of system failure and Interruption of oxygen flow to tbeual system will be installed. Oxygen pressure failure, malfunction of system components, and loss of oxygen due to leaks will not abort the mission or endanger the pilot because the operative system can be isolated from the malfunctioning system.



The dual system design oonoept Is Included In the survival kit which contains the dual oJtygen regulators. The dual oxygen regulators are Joined togetheranifold valve which automatically accepts oxygen from the operative regulatoregulator malfunctions, and blocks loss of pressure through the inoperative regulator.

The pilot alsoontrol valve on the survival kit to isolate the malfunctioning regulator. ressure gauge le included to cheok the maintenance of oxygen preesure in the emergency bail-out bottle within the survival kit.





To obtain tho maxima performance ofl airplane, one of the available boron fuels will be used in tha engines' afterburners. Tbo auparior heating value of boron fuels overtu/lb. vs S.UOOillet result ofverall lncreaso in airplane range. uch greater increase of performance can be realised when the boron fuels can be burned in the main burners of the turbojet) however, since there haa beenimited amount of testing of boron fuel* In turbojets, this will not be considered at this date. On tbo other band, there have boon highly successful runs made with boron fuelsnd Hin both afterburners and ramjets on test stands and with ramjets in flight on theehicle.

l airplane will carry0 pounds of boron - fuel0 pounds of hydrocarbon fuel in separate tanks. This does not appreciably complicate the tankage problem aa multiple tanks are required so an optimum. can be maintained during the mission by proper fuel scheduling from the various tanka. (Seef "Structural Description" Section. location vs. time curve*1

It should bo noted that the tendency of HBF fuels to cause high-altitude vapor trails ia unknown at this time. Ground testing of this material results in large volumes of white vapor, but tbe dilution occuring Id flight may reduce this to invisibility. igh speed, high altitude test inould answer this question.





l airplane will0 pounds of fuel in the wing tanks0 pound* of fuel in the fuselage. The bulk of the wing foal will behe email amount ofo the wing will be burned before aerodynamlo heating of the fuelroblem. t

The mainanks in the fuselage will be preesurlsed to8 pslg. Ofel will be used for fuel transfer andpreseure will be uaed to prevent the fuel from boiling duebeating. Thie preaaura will be maintained by dryto Insure an inert blanket ovornitrogen used will be

oarrled aboard In liquid form to save tho weight of high pressnre bottle* and also to serveeat sink for functional components.

Canter of gravity control will be accomplished by scheduling the fuel from the various tanks a* shown inf "Structural Description" section.

All tanks, fuselage and wing, will be integral with the atruoture. This oaves considerable weight and gives maximum utllitation of volume for fuel. The tanka will not be in*ulated| however. If aerodynamic heating of the fuelroblem, it may be necessary to refrigerate the fuel prior to takeoff to insure adequate heat abeorbing capacity.


acacs fuels

There are several boron fuels currently being produced ln limited quantities. By the endUn-Mathieaon'a and Gallery's large plants will be "onhich will permit sccelersted engine development, and flight testing which has bean hampered to date by the shortage of material. Olln-Kathleson Chemioal Corporation will produoe KEF which la ethyldeoaborane. Callery Chemical will produoe Hi Cal-III which is similar to ethyldeoaborane (Cj Both of these fuels are relatively easy to handle and store aa compared to the earlier boron fuelspentaborene) andpropylpentaborane) which had many "nasty" characteristics and were difficult and dangerous to handle. The new fuelsnd Hi Cal-III are less toxic, have lower vapor pressure, are not pyrophorio, and are ooopatlble with Their thermal stability has been improved and their heating values are slightly higher. There is as increaee in viscosity which will require somewhst more power to run pumps, but this Is no major problem. Listed below are some of the properties ofnd HEF-3.



be investigated. Less than IX decomposition

Decomposition and solid formation

by weight as evidenced by gas evolution and aero solid formation st the conditions listed below.



1 min. IS eeo.

ircraft corporation

eui'omiu divis

boron fuels (cokt.)

b. increase la vi sens it;

heating valuebefined aa the reaotion yielding amorphous bj oj and hj vapor

specific gravity

vlecoelty at

vapor pressure

freeileg point

spontaneous ignition temperature


after being held forin. at too ft the vlscoaity ahall not have increased moresf. i

0 minimum

7 cas aaxisum

tuab.0 btu/lb0 btu/lb max.

5 minimum

aa naxlirua

to be lnveatigated. s maximum

sis nominal

maximum to be investigated.






flash point boiling point

zero solids formedonths storags in an inert atmosphere atin gba rangeo



zero sollda formedonths storags in an inert atmosphere at temp* eratures ln the range ofto



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* i



The products of complete combustion ofr Hi Csl-III sre water and boric acid, whichery mild acid (eye wash). Little, If any, trouble is expected from this sourcs with.the possible exception of some damage to vegetation around ground run-up areas.

The unburned fuel la highly toxic when inhaled or absorbed through the skin or swallowed. Since these fuelstrong distinct odor, there le no excuse for trained personnel to Inhale sufficient Quantities toealth hazard. Absorption through the skin comes in much the same category as inhalation aa it can be easily detected and can be washed off before any damage is done. Swallowing. ofa entirely possible, but extremely improbable by trained personnel.

The Gallery Company, producers of Hi Cal, hare had lost-time accldenta dne to boron exposure, but in all cases these were not caused by the final product (Hibut by one or more of the more active agents used In procesalng such as diborane, boronhydrides, etc

Material Compatibility

There haa been enough work done on the material compatibility program to prove that the most of the materials proposed forl airplane are unaffected byr HI Cal-III. Tbe few queatisnable ones areso there will be no weight or performance compromises to be made on the airplane if substitutes are made.



Material Compatibility (Oont.)

Aa1 airplane structure is predominately titanium, which is inert to boron fuels, this will reduce the amount of teatingew metals.

Thereild reaction between boron fuels and aluminum whichor may not be tolerable for airplane oomponentej future test will verify thie. Tha amountand hum which could be used on1 airplane fuel systea is limited to tank bafflea and valve bodies for which there are several substitutes.

Steel, copper, stainless and aost other structural aetalst ia factory for use in boron fuele. Rubber and aost plastic materials must be avoided as they deteriorate rspidly in HEX. In aost casss, these materials must be replaced with Teflon,itron, or one of the other fluorineted materials.

There are eeveral currently actlvs programs for development of sealant materials being carried out for WADG. The results are promialngj however, lt ia anticipated that further testing and development will be required.

Boron fuela may be used with hydrocarbons such as the JP fuels and lubricating oils, provided the letter are free of water which hydrolit as



Material (Cont.)

the fuel andoric acid deposit which could clog fuel lines, etc.

Chlorinated material such as carbon tetrachlorld must not be used


around or with KEF as tho reaotion compounds are explosive and shock sensitive. Some of the other halogen compounds must likewise be avoided for the same reason.

Fuel Handling

The new boron fuelsnd Hi Cal-III can be handled safelyew safety precautions are taken. The fuel should be kept under an Inert atmosphere such as nitrogen gas. Spills should be avoided even though the fuel is not pyrophoric. If fuel is spilled, it should be washed away with water or burned to avoid the vapors from being inhaled. If lt is Impractical to burn the fuel, it may be hydrolitedater-methanol mixture.

Personnel handling quantltiea of boron fuels ahould wear protective clothing such as rubber gloves and sleeves, gas masks and safety goggles. The most important thing, however, is adequate training and good conson sense.

ontainer la fueled, it should be thoroughly cleaned and dried and purgedry inert gas such as nitrogen. The purging can



PUel Handling (Cent.)

beat be acooaplished by pressure purging!filling tbe container with

dryeveral tinea to its permissible operating pressure and bleeding tO_-,

ambient. If venting is required when fueling the container, the vapor should be disposed of by burning or bubblingater-methanol trap.

" IIVItldH


A. Power Plant System

I. Engine Selection

Early ln the Archensel series ofomparison of the Pratt nnd8 engine and the Oeneral3 engine was made. The performance comparisonpeolflo weight and specific thrust basis wars almost identical up toltitude. Above that height8 engine was better than

Since maximum altitudeajor criterion, the enginethe beit thrust/weight ratio could achieve the highestreliminary analysis showed thateet greater altitude could be achieved with8 engine. In addition the hardware development of8 engine isear shesd of3 engine. Consequently,8 englno waa soleated asl airplane power plant,

Engine Performance

The8 engine thrust and fuel flows at maximum power are presented in Figure 1. The data are based on the use ofn the primary burner and HEf in the afterburner. Theare based on the inlet recoveries shown in Tha data arelimb speed, up toeet andbove 7hpOO0 feet. hows the variation.



c *ii'aia division

ESKMcnr.wfica (cont.)

A. Power ^Unt System (Cont.)

II, Engine Performance (Cont.)

with afterburner power. The effect of ualng HEF in the -afterburner only resultsecrease in overall SPGver an allystem. This factor is substantially lesshooretical gain and ia due primarily to the present uncertainty of the condensation of boron-oxide in the nossle and secondarily to the difference of the molecular weights of the combustion products. It is believed that withreater factor will be achieved.

An engine weightbs. was used to allow for the large ejector diameter, for the structure required to use the engine upeet, as well as the dual fuel compatibility of the afterburner.

nduction Systea

A two-dimensional external-internal compression inlet was selected for the induction system ofirplane. chematic diagram of this inlet type is shown in Figure 3. This inlet was chosen since its layout fits tho general layout of the airplane without sacrificing performance. hree dimensional inlet would

provide slightly greater recovery, however, lt would be harder to


control. The final selection of the inlet would be made after an intensive wind tunnel program.

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A. Power Plant System (Cont.)

III. Induction System Performance (Cent.)

NASA has re canhown that at Xach ran bars greater, the inlet recovery is almost Independent of inlethe pressure recoveryi la dependent on the aaoont of ccntpression surface boundary layer renoved. ums rises the effect of boundary layer removal,


Total Head Recovery,BleedCaptnred Plow

Although the high pressure recovery ia attractive, theand/or efficient disposal off of the Inlet captured flow at high Mach number is extremely difficult without incurring large quantities of drag. For example,f flow in engine secondary noxsle results in an increase of approximatelyin SFC.

Figure li shows the inlet pressure recovery assumed in the engine analysis. Also, shown are test data obtained in wind tunnel teats conducted both by this contractor in its supersonic transport studies and by NASA Also shown for coanjarieon la the new. standard recovery curve.




A. Power Plant System (Cont.)

III. Induction System Perform oca (Cont.)

Another factor ln Inlet selection Is inlet drag. .This drag is composed of the external drag (cowl pressure and friction drag) and the Inlet spillage drar. With the corn on external compression inlet, the cowl pressure drag that eeoompaniee the large compression surfsce sngles, ae wall as the spillage dreg at off-design conditions make the slrple inlet impractical at high Hach numbers. ixed variableow angle external oowl surface is possiblethe cowl pressure drag, and the spillage drag is supersonic which Is considerably lower than subsonic spillage drag associated with spillageormal shock.

The engine locstion was determined by the. An under-wing inlet was selected since this type of inlet would be insensitive to angle of attack. It was then necessary to determine inlet location, that is, ahead of or behind the wing shock.

Locating tha inlet ahead of the wing shock has the following advantages!

s) ttoqulres no boundary layer diverter.

Utilizes the top surfsce of the inlet for additional wing area.

Allowong subsonic diffuser thereby improving pressure distribution.

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A. .Power Plant System (Cent.)

III. Induction Systea Perform nee (CcnQ

On the other hand locating the Inlet behind the shock]

) Lowers the inlet Mach masher thereby saucing it possible to

igher recovery and lowering the Inlet capture area.

Reduces the inlet momentum or ram drag of the engine thereby Increasing net thrust, since the momentum decrement has been charged to the wing pressure drag. (No allowance was mads for this in the performance).

Reduces the inlot capture area thus reducing the inlet-engine matching problem and aasoclated aplllage drag.

The principal disadvantage cf this lstter location isoundary layer diverter and ite ettsndant dreg. This contractor is currentlyull acale diverter wind tunnel teat at the NASA Lewis Research Center to verify the drag dataobtained with scale model divertars. Boundary layer diverter drag has been included ln tho performance.

In addition to the aerodynamic footors the latter location is shorter and therefore lighter. Consequently, the behind-the-ving shock under-wing inlet was Initially selected subject to the wind


tunnel program.

JfA/im/ aircraft corporation

riiiiii.ii DIVISION

riEry.CCTMAKICS (Coot.)

A. Power Plant System Bxhaust System

The Pratt and Whitney rectangular ejector is proposed for the l exhaust system. The engine manufacturer's teat data shew that there is no difference in performanceectangular and circular ejector noitle. The useectangular ejectorhe base drag problem and allows greater ground clearance angle. The inlet compression surface boundary layer air. approximately t% at cruise, will be ducted aft to aot as the secondary airflow required by the ejector. At off-designortion of the by-pass air required to cirlmize spillage drag will be dumped in the secondary nozzle. Tha quantity will have to be determined in wind tunnel teste.

Peps VTJ-7



'ii. ^ercdy ramie Heat Transfer

I. Structural Temperatures

A steady eta to heat transfer analysis was madeypical ing chord. hows the tempera tore distribution'on tbe upper and lower surface of the wing. The data show that the

average upper surface will be approximatelyF and the average


lower surface temperature will be approximately, Thein temperature between the upper and lower surface is due to the compression resulting from the wing operating atangle of attack. The heat transfer analysis was based on the Van Driest method which has been checked exporimontslly by NASA in free flight tests.

Tho temperature on the fuselage will vary fromF toF. The minimum temperature will occur on the bottom surface with maximum occurring at the'clock positions due to effects of boundary layer crossflow at angle of attack. emperature ofF will occur at the wing-fuselage intersection as well as at the intersection of any protuberances such as antenna naste. The protuberanceitself can be decreased fcy sweep. These data are based on calculation methods checked experimentally on$ configuration;

The external windshield surfece temperatures will vary fromF at the stagnation point to an average ofF over most of the

AfflHffft aircraft corporation


mejwodiwkica (coot.)

p. aerodynamic heat transfer (cent.)




tho engine manufacturer requirements ss veil as the fuelnever exceedf for eitherr he? fuel.

the resultsreliminary invsstigation show that through judicious echeduling of fuel to regions which sre leas susceptible to rpaid transient temperature rise, an uninsulated fuel tank system can be used forl airplane.

integral fuel tanks, both for the fuselage and wing, hare been designed for the airplane. at thisetailed analysis has been made for tho wing integral tanks and is presented below*. no analysis, as yet, has been made for fuselage integralowever, preliminary spot checks as well ss conclusions drawnag type fuselage tank system doalgn, discussed below, show that the fuel tanperature limitation can be adequately met. the beg type fuselage tank design study madeeasible tank design configuration in the event the sssling material problem associated with integral tanks fails te be resolved.

a preliminary heat transfer was made on typical fuel systea

Je-rk/ifm aircraft corporation

th KfTOrNAMICS (cont.)

a. aarcdynamle heat transfer. fuel system (cont.)

configuration. tho etudy vas divided into wing fuel tanks andfuel tanks. the wing fuel tanks were assumed to be'integral and that the wing fuel would be used either during climb or the initial portion of the cruise. the fuselage fuel tanks were assumed to be in bags and to be ueed only during the cruise portion of flight. no insulation waa used in either the wing or fuselage tanks. It is assumed that the wing tanks will be keptel differential and fuselage tankssl differential.

hows the time-temperature history of the wing integral tanks. the curve for use of wing fuel for climb shows that at the beginning of cruise the fuel temperature has risenF. thereaches thef limit, assuming an initialf temperature, afterinutesf the total flight. the rapid temperature rise experienced in tho tanks are due principally to the tanks being empty. also, shown ina condition if wing fuel is not used until beginning of cruise. results show that tio minutes are available beforef is reached assuming initial fuel temperature isf. this time is more than adequate to use up the wing fuel.

figureresents the tirae-teraperature history of the fuselage tanks. it should he noted that maxirauiri fuel temperature rise ief



B. Aerodynamic Beet TransferI. Fuel System (Cont.)

Were the Taper at top of tank risesF.

It is emphasized that this preliminary study should be used to give en order of magnitude since results can vary If, fuel quantities snd achedules or tank configuration are different from those assumed.

The study slso indicates that fuel can be routed from theto tho ving tanks, daring the entire mission if necessary, without exceeding tbe limit tempore tore. It is Intendedater date to usei Thermal Analyser toore accurate analysis and thereby determine the optimum fuel routing compatible. requirements as well as the actual tank configuration.

Another fuel problem other than material compatibility is the temperature effects on the residual fuel. It is presently planned to purge the tanks completely upon emptying the tanks so ss to insure no residual fuel, since any residual fuel will result in tank coking.

Original document.

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