Heavy Lift Vehicles for 1995 and Beyond

8/6/2019 Heavy Lift Vehicles for 1995 and Beyond

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a n d


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NASATechnicalMemorandum

N A S A T M - 8 6 5 2 0

A 0 7 / M F A O i

NASA

H E A V Y LIFT L A U N C H V E H I C L E S F O R 1995 A N D B E Y O N D

Compiled by Ronald ToelleProgram Development

September 1985

S t i V Y L I F T L A U S C F i V ^ . - i l C i . H j^oCl 221,

; . 3 u - I 1 2 I b

G 3 / 1 5 21^32

National Aeronaut ics andSpace A dminist rat ionGeorge C. Marshall Space Flight Center

M S F C - Form 3190 (Rev. M ay 1983)

R E P R O D U C E D B YN A T I O N A L T E C H N I C A LI N F O R M A T I O N S E R V I C EU .S . D E P A R T M E N T O F C O M M E R C ES P R I N G F I E L D . V A . 22161

F L 2827W S M C / P M E T T E C H N I C A L L IB RA R Y.V A N D E N B E R G A F B C A 93437-6021


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TECHNICAL. REPORT STANDARD TITLE PAGE1. R E P O R T NO.N A S A T M - 8 6 5 2 0 2. GOVERNMENT ACCESSION NO. 3. RECIPIENT'S CATALOG NO.4. TITLE AND SUBTITLE

H e a v y L if t L a u n c h Vehicles for 1995 a n d Beyond5. R E P O R T DATE

ueotember 19856. PERFORMING O R G A N I Z A T I O N C O D E7. AUTHOR(S)

C o m p i l e d by Ro n a ld Toelle8. PERFORMING ORGANIZATION REPORT

9. PERFORMING ORGANIZATION NAME AND ADDRESS

George C. M a r s h a l l Space Flight CenterMarshall Space Flight Center, A l a b a m a 3581210 . WORK UNIT NO.

1 I. CONTRACT OR G R A N T NO.

12 . SPONSORING AGENCY NAME AND ADDRESS

N a t i o n a l Aeronautics and Space AdministrationW a s h i n g t o n , B.C. 2 0 5 4 6

13. TYPE OF REPORT & PERIOD COVERED

Technical M e m o r a n d u m14. SPONSORING AGENCY CODE

15. SUPPLEMENTARY NOTES

Prepared by P r o g r a m D e v e l o p m e n t .16. ABSTRACT

A Heavy Lift Launch Vehicle ( H L L V ) designed to deliver 3 0 0 , 0 0 0 Ib t o a 5 40n . m i . circular polar orbit may be required to m e e t national needs for 1995 andbeyond. The vehicle described herein can a c c o m m o d a t e payload envelopes up to50 ft diameter by 200 ft in length. Design requirements include reusability forthe mo re expensive components such as avionics and propulsion systems, rapidlaunch turnaround t i m e , minimum hardware inventory, stage and c o m p o n e n t flexi-bility and c o m m o n a l i t y , and low operational costs. All ascent propulsion systemsutilize liquid propellants, and overall launch vehicle stack height is mi ni mi z ed whilemaintaining a reasonable vehicle diameter. The ascent propulsion systems are basedon the development of a new liquid oxygen/hydrocarbon booster engine and liquidoxygen/liquid hydrogen upper stage engine derived from today's SSME technology.W h e r e v e r possible, propulsion and avionics systems are contained in reusablepropulsion/avionics modules that are recovered after each launch.

17. KEY W O R D SH e a v y L i f t Launch V e h i c l eLarge Payload V o lum eReusable Propulsion and AvionicsLarge Oxygen/Hydrocarbon Booster Engine

18. DISTRIBUTION STATEMENT

Unclassified U n l i m i t e d

19. SECURITY CLAS SIF. (of this report)U n c l a s s i f i e d

20. SECURITY CLASSIF. (of thl pag)Unclassified

21. NO. OF PAGES

14822. PRICE

N T I SM S F C - form 3 2 92 ( M y 1969) I For sale by National Technical Information Service, Springfield. Virginia 22151


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T A B L E O F C O N T E N T SPage

I. I N T R O D U C T I O N 1II. . C O N F I G U R A T I O N S U M M A R Y 2

I II . S E L E C T E D C O N F I G U R A T I O N R E S U L T S . 3I V . L A U N C H F A C I L I T I E S A N D G R O U N D O P E R A T I O N S 8

V . TEST P R O G R A M 8V I. D E V E L O P M E N T S C H E D U L E S 8

V I I. E V O L U T I O N A R Y C O N C E P T S 9VIII. D E S I G N AND M A N U F A C T U R I N G . .. 9

IX . C O N C L U S I O N S 1 0R E F E R E N C E S 12A P P E N D I X A . A E R O D Y N A M I C S 3 5A P P E N D I X B . A V I O N I C S A N D S O F T W A R E 4 5A P P E N D I X C . D E V E L O P M E N T S C H E D U L E S 6 5A P P E N D I X D . L A U N C H F A C I L I T I E S A N D G R O U N D O P E R A T I O N S 7 3A P P E N D I X E . P E R F O R M A N C E 7 9A P P E N D I X F . P R O P U L S I O N S Y S T E M S 8 9A P P E N D I X G . R E E N T R Y D A T A 9 7A P P E N D I X H . S T R U C T U R E S A N A L Y S I S 1 0 7A P P E N D I X I. TESTS 123A P P E N D I X J. T H E R M A L A N A L Y S E S 129A P P E N D I X K . C O N T R I B U T I N G P E R S O N N E L 1 3 9

P r e c e d i n g p a g e b l a n k


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L I S T O F I L L U S T R A T I O N SFigure

1.2 .3.4.5.6.7.8.9.

10.11.12.13.14. '15.16.1.7.18.19.20.21.2 2 .23.2 4 .25.

TitleH L L V options comparisonSpace Transportation Booster Engine ( S T B E )Space Transportation M a i n Engine ( S T M E 481)Heavy L i f t Launch VehicleAscent flight profileDesign reference mission profileP /A M o d u l e recoveryBooster recovery scenarioCenter of pressure variationCrosswind profile versus altitude -.

oPitch and yaw angle of attack history with four 6 7 5 - f t fins2Y aw gimbal angle c o m m a n d history with four 6 7 5 - f t fins

2Pitch and yaw angle of attack history with four 3 0 0 - f t fins2Y aw gimbal angle c o m m a n d history with four 300-ft fins

H L L V propellant feed line layoutCenter core stage descriptionSingle engine boosterT w o engine boosterP/A M o d u l e structure and propellant linesP / A M o d u l e component weightsPayload fairingPayload fairing longitudinal separation systemTest program elementsDirect approach development scheduleDerivative applications

Page131414151617171819192 02 021212 22 32 42 5

. ' 26. -, 26

2 7282 93031

IV


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L I S T O F T A B L E STable Title ' Page

1 . H L L V W e i g h t S u m m a r y 3 22 . H L L V Detailed W e i g h t S t a t e m e n t 3 23. Payload Fairing Structural W e i g h t s 33


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D E F I N IT I O N O F S Y M B O L S

a Alpha, angle of attack in pitchARef Reference area, TrD2Ref/43 Beta, angle of attack in yawB, 171 in. diameter boosterB2 246 in. diameter boosterC Forebody axial force coefficientAFC . Total axial force coefficientA TC,-. Drag coefficientC P Center of pressureCM Normal force coefficient, N/q AD -I N K e iCN N o r m a l force coefficient slope, (a


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D E F I N I T I O N O F S Y M B O L S ( C o n c l u d e d )g Accelera t ion of gravityM ax q M a x i m u m d y n a m i c pressureq D y n a m i c pressureq F r e e s t re a m d y n a m i c pressure

Vll


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LIST O F A C R O N Y M S A N D A B B R E V I A T I O N S

APS Auxiliary Propulsion SystemB SM Booster Separation MotorC A D Computer Aided DesignC A E Computer Aided EngineeringC A M Computer Aided MachiningC G Center of GravityCP Center of PressureDRM Design Reference MissionE A F B Edwards Air Force BaseET External Tank *G L O W Gross Lift Off WeightG S E Government Supplied Equipment

.*G S E Ground Support EquipmentH C HydrocarbonH L L V Heavy Lift Launch VehicleIO C Initial Operational CapabilityI _ Specific ImpulsebpK S C Kennedy Space CenterL C C Launch Control CenterL/D Lift-to-Drag RatioL H g Liquid HydrogenL O Lift O f fL O X Liquid OxygenL R B Liquid Rocket BoosterM E CO Main Engine Cut OffM L P Mobile Launcher Platform

Vlll


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L I S T O F A C R O N Y M S A N D A B B R E V I A T I O N S (Concluded)

M M H M o n o m e t h y l HydrazineM S F C Marshall Space Flight CenterN ^ O . Nitrogen Tetroxiden . m i . Nautical M i l eO M S Orbit Maneuvering SystemO T V Orbit Transfer VehicleP / A ' Propulsion/AvionicsP L F . Payload FairingPSF Pounds (force) per Square FootRCS ' Reaction Control SystemRF Radio FrequencyS D / H L V Shuttle Derived Heavy L i f t VehicleSD V Shuttle Derived VehicleSRB Solid Rocket BoosterS S M E Space Shuttle M a i n EngineS T A R Shuttle Turnaround Analysis ReportSTS Space Transportation SystemSTBE Space Transportation Booster EngineS T M E Space Transportation M a i n EngineTPS Thermal Protection SystemV A F B Vandenberg Air Force BaseW T R Western Test Range

IX


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T E C H N I C A L M E M O R A N D U M

H E A V Y LIFT L A U N C H V E H I C L E S F O R 1995 A N D B E Y O N D

I. I N T R O D U C T I O N

A Heavy L i f t Launch Vehicle ( H L L V ) designed to deliver 3 0 0 , 0 0 0 Ib payloads toa 540 n. mi . circular polar orbit may be required to meet national needs for 1995 andbeyond. The vehicle described herein can accommodate payload envelopes up to 50 ftdiameter by 200 ft in length. Payloads utilizing this capability may be Space Stationelements, commercial space facilities, or advanced military systems.Design requirements include reusability of the more expensive components suchas avionics and propulsion systems, rapid launch turnaround time, m i n i m u m hardwareinventory, stage and component flexibility and commonality, and low operational costs.

All ascent propulsion systems utilize liquid propellants and overall launch vehiclestack height is m i n i m i z e d while maintaining a reasonable vehicle diameter.The ascent propulsion systems are based on the development of a new liquidoxygen/hydrocarbon booster engine and a liquid oxygen/liquid hydrogen upper stageengine derived f r o m today's S S M E technology. The upper stage engine will havemore thrust than the S S M E , be more reliable with less maintenance, and have a two-position nozzle. The requirements placed on the avionics system are more stringentthan on present launch' vehicles because of the rapid turnaround of reusable com-ponents and the necessity to maintain continuous launch readiness after stackup.Wherever possible, propulsion and avionics systems are contained in reusablePropulsion/Avionics (P/A) Modules that are recovered after each launch. The P/AModule has an ablative non-reusable Thermal Protection System (TPS) and a crushablehoneycomb nose cone to absorb landing loads. P/A Module recovery is baselined as aterrene landing to avoid the complexities associated with water landing and recovery.The storable propellant Reaction Control System (RCS) and Orbit Maneuvering System( O M S ) are also located in the P/A Module.The H L L V structu ral design is based on current Space Transportation System( S T S ) technology to meet the Initial Operational Capability (IOC) date of 1995. Tech-nology advancements in- structural design and materials may increase the payloaddelivery capability, but at the cost of a longer development schedule.T w o development approaches are considered. The first approach is the direct

.development of a mature H L L V without intermediate steps. The second approach hasan extended schedule where the booster systems are initially developed for applicationto the STS and Shuttle Derived Vehicle (SDV) programs, or to a new intermediatevehicle prior to H L L V application. This approach will reduce front end developmentcosts but extend program development time and delay IOC. The first approach willcompress overall development time leading to an earlier IOC but will increase initialcosts accordingly.


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I I . C O N F I G U R A T I O N S U M M A R Y

Three configuration concepts were investigated during this study, all satisfyingthe previously described requirements. Figure 1 shows the m o l d line and base viewof each. The aerodynamic fairing for the 50 ft by 200 ft payload is jettisoned at350,000-ft altitude for all concepts. The booster stages of all three configurationsuse new L O X / J P 4 gas generator cycle booster engines, designated Space Transporta-tion Booster Engines (STBEs), having a sea level thrust of 1.5 to 2.0 million Ibfdepending oh specific booster application. All upper stages use staged combustioncycle L O X / L H 2 engines derived f r o m the S S M E and designated Space Transportation

M a i n Engines ( S T M E s ) . The S T M E has a two-position nozzle for altitude compensationwith thrust varying f r o m 3 9 7 , 0 0 0 Ibf at sea level to 4 8 1 , 0 0 0 Ibf in vacuum. Theoperating characteristics and performance parameters for the S T B E and S T M E areshown in Figures 2 and 3, respectively. Very brief descriptions of the three con-figuration concepts f o l l o w ; however, Configuration II shows the greatest potential andis discussed in more detail in the remainder of this report.

A. Configuration IConfiguration I is a series/parallel burn three-stage vehicle designed for com-monality of propellant tanks. All tanks have the same diameter to reduce the costsof development, design, tooling, production, and qualification. The L O X / J P 4 firststage, or booster, consists of four tank sets of sub-stages, each with two 1.75million Ibf sea level thrust STBEs. The second and third stages use L O X / L H 2 pro-

pellants and have four and two STMEs, respectively. The second stage consists oftwo tank sets. The third stage consists of a single tank set centered within thefirst and second stages. This configuration m i n i m i z e d the vehicle stackup height.The aft ends of the booster sub-stages are connected by box beams to distributeatmospheric flight bending moments and m i n i m i z e structural weight of the upperstages. Longitudinal thrust loads during booster flight are distributed to the vehicleat a forward payload attach ring. The upper stages are carried in tension duringbooster burn with the S T M E nozzles in the retracted or stowed position for moreefficient packaging and better thermal control.

After first stage separation, the second and third stages are ignited simul-taneously and burn in parallel. The second stage, which consists of two tank setsor sub-stages, crossfeeds propellants to the core third stage. At second stage pro-pellant depletion, the two sub-stages are separated and expended. The high stagingvelocity (Mach 16) results in a down range distance too great for practical recovery,even if the hardware could survive reentry.After second stage separation, the third stage continues to burn into theperigee of a 100 x 540 n.mi. elliptical orbit. Following a coast to apogee, the thirdstage reignites, placing the payload into the operational 540 n.mi . circular orbit.After payload deployment, correct separation distance and orbit phasing, the thirdstage reignites to deboost the stage for disposal.First stage hardware is recovered after water landing; however, designs forsecond and third stage recovery systems were not pursued for this configuration.Additional design details for Configuration I are available in Reference 1.


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B . Configuration IIConfiguration II is a parallel burn two-stage vehicle designed without the hard-w a r e c o m m o n a l i t y constraints of the previous configuration. The flight profile usesdirect insertion into a 100 x 540 n.mi. orbit by the first and second stages w i t hcircularization at apogee achieved'by either a kick stage or pay load-supplied propul-sion system.The L O X / J P 4 first stage, or booster, consists of four tank-sets or sub-stages,tw o are 246 in. in diameter and two are 171 in. in diameter. The larger diametersub-stages have two 1.616 million Ibf thrust STBEs each and the smaller tank setshave one S T B E each, for a total of six booster engines. Each booster tank set con-tains three propellants: liquid hydrogen, liquid oxygen, and a hydrocarbon f u e l( J P 4 ) . The second stage is 396 in. in diameter and has five two-position nozzleS T M E s . All first and second stage engines are ground, ignited and f l o w n in parallelburn until booster staging. During booster burn liquid oxygen and liquid hydrogenare crossfed f r o m the first stage tanks to the S T M E s . This procedure shortens thesecond stage hydrogen tank by 30 ft and reduces weight.It is feasible to recover the booster sub-stages and the second stage propulsionand avionics hardware (housed in a P/A Module).

C. Configuration IIIConfiguration III is a two-stage inline series burn vehicle that has been definedonly to the depth necessary for comparison with the other options. The L O X / J P 4first stage is 50 ft in diameter and has eight STBEs. The L O X / L H 2 second stage,

also 50 ft in diameter, has five S T M E s . All LOX is carried in the second stage (i.e.,the first and second stages share a c o m m o n LOXtank). The booster stage isrecovered and the second stage propulsion and avionics hardware is assumed to berecoverable by the use of an appropriately designed P/A Module. This option isviewed as having very little growth potential, essentially no capability for stageelements to be used as intermediate class vehicles, and would pose difficulties inground transportation and handling.

III. S E L E C T E D C O N F I G U R A T I O N R E S U L T S

A. Two-Stage Parallel Burn Configuration (Configuration II)The configuration resulting from detailed analyses is shown in Figure 4. The

L O X / J P 4 recoverable first stage consists of four tank sets or sub-stages. Two of thesub-stages are 246 in. in diameter and two are 171 in. in diameter. The large diame-ter sub-stages have two STBEs and the small diameter sub-stages have a single S T B E ,Each sub-stage of the booster contains three propellants: liquid oxygen, liquid hydro-gen, and JP4. The L O X / L H 2 second stage is 396 in. in diameter, has five S T M E scontained in a recoverable P/A M o d u l e , and a forward structural adapter/payloadattach ring. A vehicle weight s u m m a r y is displayed on Table 1. The resulting grossl i f t o f f weight of this configuration is 8.6 million Ib. The sub-stage and second stageweight statements are detailed on Table 2. Structural details are contained inAppendix H. .


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B. Ascent Flight ProfileThe ascent flight profile is shown in Figure 5. All engines are ground ignitedon the launch pad and burn in parallel until booster c u t o f f and staging. The secondstage S T M E nozzles are retracted (e = 55:1) until booster staging, then extended(e = 150:1) and the second stage continues into the perigee of the 100 x 540 n.mi.

orbit. All propellants for the parallel burn portion of flight are carried in thebooster tanks with LOX and LH9 being crossfed into the second stage. This pro-ucedure allows the second stage tanks to be f u l l at booster burnout, reducing the tankvolume and weight required to be carried to orbit.C. Design Reference Mission and Deorbit Events

The design reference mission profile is shown in Figure 6. The flight eventsare depicted and times of occurrence covered from l i f t o f f to P/A M o d u l e landing. Thesecond stage propellant tanks are expendable and require controlled debcost f r o morbit. Following payload separation, the second stage coasts in orbit for approxi-mately one revolution. It is then oriented for the deboost burn attitude by the P/AModule and a slow roll started for attitude stabilization during the retro burn. TheP/A Module separates and performs an evasive maneuver a w a y from the tankage. Anonboard timer provides the deboost ignition signal to six solid rocket motors mountedin the forward conical adapter (see Appendix F). At the appropriate time, the P/AModule performs an orbital adjustment burn to allow correct phasing to align theorbital plane with the landing site (Edwards Air Force Base for this study).The P/A Module performs a deboost burn using the aft firing QMS storablepropellant engines. The reentry environment of the P/A M o d u l e is described inAppendix G and the resulting thermal protection system requirements in Appendix J.W h e n the P/A Module has slowed to approximately M a c h 1 after reentry, drogue para-chutes are deployed for added drag and stabilization. Parawing type steerable

devices were selected for the terminal landing event (Fig. 7). Terrene landing on acrushable nose cone has been baselined to reduce the refurbishment requirements andturnaround time. The honeycomb nose cone structure and the TPS are consideredsacrificial and are replaced after every launch.

D. Booster RecoveryAt booster burnout, the sub-stages are separated and fall into the ocean forpartial retrieval. Figure 8 shows a recovery flight profile of a booster stage fromdrogue release to water impact. The retrieval technique is called the hydropneumaticoption. Following separation the boosters coast through apogee and reenter theatmosphere. The induced environment during reentry is detailed in Appendix G.

More detailed studies are required to define when the aerodynamic fins are to beseparated f r o m the sub-stages. After the boosters have slowed to approximatelyM a c h 1, drogue parachutes are deployed for stablization. After the m a i n chutes aredeployed, a linear shaped charge severs the forward tank dome f r o m the barrelsection just aft of the "Y-ring." Shaped charges are used to provide "vent" holesin the tank barrel section just forward of the aft d o m e "Y-ring" to allow the trappedair to escape at water impact, providing the pneumatic cushion that yields a softlanding. Vent hole size will be traded between deceleration and rebound. The JP4tank will provide the flotation and a flotation ring will be deployed to provide stabilityuntil recovery. A protective spray bag will be deployed before impact to enclose the


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a f t end of the engine compartment and thus keep the booster engines dry. Thebooster propulsion and avionic subsystems are recovered for reuse and the JP4 tankhas a high probability of reusability. The boosters will float in the inverted positionuntil removed from the water. Cleaning for refurbishment can begin on the way backto port.E. Fin Size Selection

Fin size effects on ascent aerodynamics were traded early in the study. Thedetailed aerodynamics generated for this launch vehicle are contained in Appendix A.The static stability of a launch vehicle is determined by the distance between the.centerof gravity and the center of pressure. Figure 9 displays the estimated pitch plane9center of pressure (CP) variation versus the M a c h number for no fins, four 300 ft" .2fins, and four 675 ft fins. The estimated longitudinal center of gravity is alsoshown. The center of pressure is always forward of the center of gravity, resulting2in an unstable vehicle. Based on these data, four 675 ft fins were baselined untilmore detailed analyses could be performed. A cursory control study was performedafter the design was frozen.

F. First Stage ControlFlight simulations with six-degree-of-freedom rigid-body dynamics and three-axis control were performed for the booster phase of flight. The preliminary resultsshow that, with only, the booster engines controlling, sufficient control authority isavailable with a square 6 deg gimbal pattern to withstand the effects of one boosterengine out at 45 sec and the M S F C 95 percentile synthetic wind profile with theembedded gust [2]. Figure 10 displays the crosswind profile versus altitude.

2 2Simulations were made for the 675 ft fins, 300 ft fins, and for no fins with.the engine failure at 45 sec of flight and the wind speed of 270 ft/sec peaking at4 3 , 0 0 0 ft altitude (near m a x i m u m dynamic pressure). One engine of the right sidetwo-engine booster was assumed to fail. This causes an immediate yaw m o m e n t of5 0 million ft-lb. The available control torque is 90 million ft-lb. A left crosswindis assumed, thus creating positive yaw m o m e n t , additive with the engine failurem o m e n t .2Figures 11 and 12 show that with the four 675 ft fins, the required enginegimbal c o m m a n d does not reach the 6 deg limit due to the engine failure alone, andjust momentarily touches it when the crosswind peaks. The pitch angle of attack( A L P H A ) is smooth, but the yaw angle of attack ( B E T A ) has a m a x i m u m value of

6 deg during the period of m a x i m u m dynamic pressure.2Results from the analysis using the 300 ft fins are displayed in Figures 13and 14. The yaw gimbal c o m m a n d hits the limit for both the engine out and cross-wind disturbances. The resulting B E T A angle is larger and lasts longer than forthe larger fins.The no-fin case can only be controlled through the wind disturbance; if abooster engine fails, vehicle control is lost. Since performance effects of these con-2trol studies remain to be determined, the 675 ft fins were retained as the baseline.


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G . Engine Out OperationPropellant plumbing has been added to the launch vehicle to provide propellantcrossfeed in case of a non-planned engine out during ascent. The operationalphilosophy is to set the flight power level at a given percentage less than the nominal1 0 0 percent power level. The S T B E flight power level will be 83 percent, and the

S T M E 80 percent. If a S T B E or S T M E fails, the functioning engines will be advancedto 100 percent for the remaining burn time. Control gains and guidance presettingswil l be automatically switched by the onboard computers. Additional studies, his-torically called "Abort and Alternate Mission," will be required for premission planningbecause, once launched, the vehicle is c o m m i t t e d to space or destroyed.H. Avionics

The H L L V avionics is composed of several m a j o r subsystems: c o m m u n i c a t i o n sand tracking, data processing, guidance and navigation, flight control, propulsioncontrol, auxiliary flight control, electrical power, and range safety control. Designo f the avionic subsystems wil l utilize current and evolving technology to m e e t theobjective of improved performance with m i n i m u m risk, reduced turnaround t i m e , andreduced cost; Advanced technology in data processing will enable m u c h higher levelsof automation to support design, analysis and mission planning, reconfiguration,checkout, and launch. Advanced distributed fault tolerant processing architecturesand methodologies will be utilized to provide very reliable flight systems that can bepartitioned for vehicle modularity, contractual separability, and interface simplification.Advanced electronic technology will also make possible a high level of vehicle autonomy.Many functions that have previously been performed by ground support equipment orthe launch control center will be performed on-board the vehicle to m i n i m i z e checkoutand launch support personnel, vehicle turnaround time, and vehicle/GSE interfacecomplexity.

M u c h of the vehicle avionics is distributed and physically dispersed to achievemodularity and partitioning objectives. Other vehicle avionic functions such as RFcommunications tend to be centralized in nature and relatively independent of thevehicle configuration. Subsystems with these centralized functions are integratedinto a "central avionics package" located in the P/A Module. This central avionicspackage can be treated as a m a j o r vehicle m o d u l e , completely integrated within itselfand having "clean;t data bus and power bus interfaces.The elements of both the distributed and the centralized avionic systems on thevehicle are connected with a fault tolerant network. It is envisioned that the networkwill consists of a number of physically dispersed processing sites connected by afault tolerant communication network. The general purpose processors at each sitecan have varying levels of throughput, m e m o r y size, fault tolerance, and local

input/output.A more complete description of the requirements, objectives, and characteristicsfor the avionic systems is given in Appendix B.


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I. Vehicle DescriptionFigure 15 is a computer-generated drawing emphasizing the propellant feedlineso f the launch vehicle. The second stage propellant tanks, feedlines, and componentweights are detailed in Figure 16. The LOXfeedlines are routed externally to avoidpenetration of the LH9 tank. Two feedlines are used instead of one .large line toLtm a k e the stage more symmetric in roll for the deboost stabilization. The aft plumbingin the core stage is the LOXcrossfeed connection for the P/A M o d u l e and the JP4crossfeed to ensure booster propellant depletion in case of a booster engine out. Thebooster engine out LOXcrossfeed is available as a result of crossfeeding propellantsto . Stage 2, therefore, only additional JP4 lines are required. Various layout com-binations were investigated for the JP4 crossfeed and the one presented here is thesimplest. The LH2 feedline is a short sump feeding in to the P/A Module. The LH2

crossfeed inlets are located at the top of the tank at a level which will allow adequateullage pressure v o l u m e at booster staging.The one- and two-engine boosters are detailed in Figures 17 and 18, respec-

tively, including a weight summary for each booster size. The moldline dimensionsare displayed and primary stage and crossfeed propellant lines detailed. The 675 ftfins are displayed. The upper feedlines are for LH2 crossfeed into the core stage.The LOX feedlines run d o w n the side of each stage and split, one part routed to thesecond stage and the other to the L O X / J P 4 engines of the respective stages. TheJP4 plumbing splits to the engine and to the crossfeed manifold in the second stage.The recovery parachute system is housed in the engine skirt region, and the postlanding flotation collar is housed within the aft JP4 tank area.

The P/A Module is the m o s t expensive component of the launch vehicle. Struc-tural and plumbing drawings and a weight summary are displayed in Figures 19 and2 0 . The four S T M E s are clocked 45 deg to the flight plane to reduce mechanical andheating interference with the boosters. The propellant distribution feedlines are ondifferent planes to eliminate plumbing interferences. The individual lines for eachengine straddle the thrust structure crossbeam. The gimbal points of the outboardengines (the inboard engine is fixed) will be 45 deg out of plane, i.e. , rock and tiltas on the Space Shuttle Solid Rocket Boosters. These commands will be transformedby software within the control system. The Orbital Maneuvering System (OMS)andon-orbit Reaction Control System (RCS) will be housed in four replaceable modules,plumbed to the storable tanks. Three axis RCS and longitudinal translation is pro-vided by this system. Thruster details are in Appendix F. The P/A Module housesmost of the avionics for prelaunch checkout, flight, and reentry.. Details andspecific requirements assigned to each launch vehicle component are presented inAppendix B. Structural design details are in Appendix H.

The Payload Fairing (PLF) is used to protect the payload through aerodynamicflight. A moldline drawing is presented in Figure 21, and structural weights inTable 3. The PLF has a 50 ft outside diameter and houses a 200 ft long payload,not including the length available within upper nose cone. A double angle nose conewas selected based on studies performed for the Saturn launch vehicle [ 3] . Thisgeometry provides an efficient trade between aerodynamic drag and internal volume.The PLF is designed in 24 ft long cylindrical sections for adaptability to variouspayload lengths. .


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The PLF will be separated into four 90 deg longitudinal sections which w i l l bejettisoned in flight and expended. The longitude separation system is a non-con-taminating expanding tube device as f l o w n on the Skylab Mission (Fig. 22). Separa-tion f r o m the PLF adapter ring can be by explosive bolts or some other suitabledevice.

I V . L A U N C H F A C I L I T I E S A N D G R O U N D O P E R A T I O N SThe baseline launch scenario for the H L L V is a due south polar orbit launch.Considerations in selecting the launch site include the large size of the vehicle, rapidbuildup and payload changeout requirements, non-interference with STS flights,launch azimuth and overflight restrictions, and other practical-factors. A new launchsite is recommended and possible locations include Hawaii, southern Alaska, theVandenberg Air Force Base area, and certain other regions of the continental UnitedStates. The facility size will be a function of launch rate, however, there should beat least two launch pads for parallel launch capability, served by a single LaunchControl Center. The vehicle is to be built up in an assembly building and transportedto the launch pad by a mobile launcher. This approach provides more efficient useof the launch pads, allows parallel vehicle processing, isolates the launch pads f r o mthe buildup area, and facilitates launch vehicle changeout. The overall operationalsequences given in Appendix D are similar to the STS processing f l ow at KSC, andsome timelines (such as rollout, pad refurbishment, and mobile launch refurbishment)are derived from the STS processing assessment, S T A R - 0 2 7 .

V. TEST P R O G R A MThe major elements of the test programs for the H L L V are depicted in Figure 23and detailed in Appendix I. The protoflight approach given in M I L - S T D - 1 5 4 0 B is tobe used and is best exemplified by the large structural test items, which will be

tested to levels exceeding flight loads but lower than yield values. Following rework,the test items can be used as flight hardware, thus avoiding the significant cost ofdedicated hardware for testing only.The test program is unique in the number and size of the tests to be conducted.An evolutionary or derivative approach would require more testing than the directdevelopment approach. Every stage will have to be tested to the loads expected foreach flight application. .It is important that all vehicle elements be designed to,facilitate testing. Designpersonnel should be involved not only in the test planning process but should beengaged in all subsequent phases of the test program.

VI. D E V E L O P M E N T S C H E D U L E SThe scope of the H L L V project is similar in size to the development of theSaturn V and STS launch vehicle systems. New manufacturing and launch facilitieswill be required to avoid interference with the operational STS program.The development of this launch vehicle system may be approached either directlyor in an evolutionary fashion. The direct approach, which can include derivatives,leads to development of a mature flight system without intermediate steps. Theevolutionary or derivative approach takes longer than the direct approach; however,

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the former may be more cost effective than the direct approach. A f amily of vehicleswith diverse capabilities, all built w i t h the same tooling, avionics, and propulsion,will be produced. Additionally, other vehicles can be derived including some thatare sized to fit the STS cargo bay diameter.The summary schedule is shown in Figure 24 and the supporting details con-tained in Appendix C. The development of derivative launch vehicles may be concur-rent with H L L V development or be delayed until the H L L V is operational. The directapproach has the earliest IOC but requires extensive funding early in the program.The evolutionary development is characterized by a building block approach wherethe STBE and booster sub-stages are developed initially for application with othervehicles, to be later integrated with the second stage, P/A M o d u l e , and payloadfairing to comprise the H L L V . This approach reduces initial funding requirementsat the penalty of a later IOC for the H L L V . The S T B E is the pacing i t e m for eitherdevelopment approach and this engine's design, performance, and operating charac-teristics must be established early.

V I I. E V O L U T I O N A R Y C O N C E P T SThe modular design allows the development of alternate launch vehicles from thebasic H L L V stage elements or the use of these elements in other space applications.It is not intended that elements or stages be directly exchanged, rather thatthe design and manufacturing data base, tooling, and assembly processes be effec-tively used for other applications (Fig. 25). The reference H L L V configuration hasexcellent growth capability by interchanging booster stages and increasing the powerlevel setting of the stage two engines. Replacing the single engine boosters withtw o engine boosters for a total of 4 two engine boosters, results in an increase ofthrust and available propellant. The crossfeed propellant capacity is increased, andto consume this requires increasing the stage two thrust during booster burn. Thislarger booster configuration will launch approximately 600 Klb of payload into a refer-ence 100 x 100 n.mi., 28.5 deg inclination orbit;A small payload two stage vehicle was derived from the two engine booster hard-ware. The first stage has two STBE's and the upper stage a single S T M E . Theperformance of this configuration is approxiamtely 123KIb to the 100 n.mi., 28.5 degorbit and is detailed in Appendix E. Shuttle configurations in which the SRBs havebeen replaced with modified H L L V liquid rocket boosters (LRB )were studied. Ini-tially, 2 two-engine LRBs with a full propellant load were investigated for the Shuttle.This results in a payload capability of approximately 167KIb which exceeds the cur-rent Shuttle load carrying capability; however, this capability could possibly be used

for added mission flexibility (e.g., higher orbits, plane changes, etc.). A secondoption was investigated in which three single engine HLLV-derived LRBs were usedwith the Shuttle. This results in a payload capability of 95K Ib to the above refer-enced orbit.An intermediate class heavy lift vehicle, the Shuttle Derived/Heavy L i f t Vehicle( S D / H L V ) , was studied which consists of a modified Shuttle ET with a reusable pro-pulsion/avionics module as the second stage, boosted by a pair of the two engineH L L V boosters. This configuration has a payload capability in excess of 3 0 0 K Ib tothe reference orbit.


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In addition to the above, the single engine booster diameter (176 in. ) wasselected for potential application within the Shuttle cargo bay. The cryogenic tankshave potential application as orbital tankers, or with the addition of a propulsionsystem, an OTV. These particular applications could utilize tankage with reducedskin gauge.The 33 ft diameter second stage H L L V tanks could be configured into a large

propellant storage facility, launched by the H L L V . Propellant loading requirementsof this nature have been defined by the interplanetary mission studies.The next section details some of the thought processes required during thedefinition phase of this or the next generation of vehicles.

V I I I . D E S I G N A N D M A N U F A C T U R I N G

The design and manufacturing philosophy for the H L L V emphasizes versatilityand flexibility of operation rather than unique application. The more expensivestructures are the tank domes, thrust structure, interfaces, and attach points. Thelength of the cylindrical barrel sections can be changed easily if the original designprocess has foreseen this requirement and the required strength may be attained byreprogramming the numerical milling machines that are used for panel cutting. Thisprocess can also be applied to the tank domes but trade studies should verify areasonable payback for the change in metal thickness versus reduction of payload orincreased propellant load required to deliver a given payload.

This adaptability can be achieved by addressing the multiple use aspects ofcomponents early in the design process. Computer aided design and computer aidedengineering ( C A D / C A E ) may be effectively applied. Before designing the manufactur-ing facility, CAD layouts should be performed to eliminate interference of differentoperations. Skin thicknesses and tolerances should be traded against manufacturingcosts and other processes.

Manufacturing facilities should be designed to fabricate multiple stage sizesvarying in diameter and length. Machining will be performed by high speed computeraided machines ( C A M ) and changes in skin thicknesses can be controlled f rom thedesign computer.Procedural changes to vertical buildup and assembly of the tank structures isrequired to accommodate different cylindrical lengths without m a j o r floor modificationsor extensive foundation restructuring. This must be addressed during plant designand before construction. Vertical buildup requires fewer internal sectional hoop jigsto maintain roundness during welding. Trades will determine whether the weldingturret will be movable or the assembly table move vertically within either a high bayor a pit. Correct design will allow welding of different diameters by reprogrammingthe drive computer.Plasma arc welding technology should be investigated as a method of reducingproduction time, weld inspection and ultimately reducing costs. Extrusion forming ofshort interstage panel sections may reduce costs for high production rates. Thisprocedure would have structural weight penalties due to the constant skin thicknessbut may be more cost effective than tapered panels. Upper stages may benefit fromthe use of composite materials, however, high production rate techniques will berequired to be cost effective.

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Trade studies must determine the cost differences between unique and generaldesigns and if payload penalties are justified .by easier manufacturing operations.Reducing machining tolerances or allowing thicker panel sections could allow contract-ing to lower overhead rated machine shops.The guideline of all these trade studies is to define a system that is affordable.The mass production state-of-the-art techniques must be implemented.

I X . C O N C L U S I O N S

The pacing development i t e m for the H L L V is the new L O X / J P 4 engine. Tomeet the projected capabilities of post-1995, the development of the S T B E should bestarted very soon. New recovery methods and hardware for the boosters and P'/AModule are required. A terrene landing instead of water touchdown could allowcomplete reuse of the boosters while turnaround times and refurbishment costs arereduced. A sub-scale reentry flight test of the P/A Module will be required to verifystability, attitude control, and TPS during reentry, and the steerable parachutesystem for the final landing must be demonstrated.

Updated design, engineering, and manufacturing processes must be applied tothis next generation of launch vehicles. The building block use of stage componentsfor evolution or derivation of other vehicles must be recognized at the beginning ofdesign and carried through the systems' lifetime. Tooling and factory layout mustreflect operational flexibility with a goal of no down time for conversion f rom onesize component to another. High speed, machinery and m i n i m u m inspection weldingprocesses will make large contributions to cost reduction.High technology materials may be implemented as product improvements within

the life of this system. Goals must be defined before implementation consideringpossible payload capability increases versus cost increased incurred by implementinga technology.N e w methods of efficient configuration management are required. One conceptis to make vendors responsible for designated end items, rather than assuminggeneral liability. This will reassign the warranty to the vendor and create incentivesto reduce nonproductive costs. The data presented above is to be considered as a point of departure forfollowing launch vehicle studies in the areas of design, materials, propulsion, con-trol, all avionics functions, manufacturing, and operations. Product improvementstudies in all areas are welcomed, especially those leading to a reduction in pro-

gram costs.

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R E F E R E N C E S

1 . Marshall, W . R . , a n d Shelton, B , W . : Advanced Launch Vehicles. 1984S A EAerospace Congress and Exposition, N A S A / G e o r g e C. Marshall Space FlightCenter.2 . Garner, Doyle: Control Theory Handbook. N A S A / G e o r g e C . M a r s h a l l SpaceFlight Center, T M X - 5 3 0 3 6 , April 2 2 , 1964.3 . Geissler, E . D . : M e m o r a n d u m o n N o s e Shape o n Saturn Vehicles. GeorgeC. Marshall Sapce Flight Center, April 9, 1963.

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PROPELLANTS NOZZLE AREA RATIO THRUST (SEA LEVEL)- KLBF DELIVERED SEA LEVEL ISP SEC CHAMBER PRESSURE PSIA MIXTURE RATIO (O/F) LENGTH IN NOZZLE EXIT DIAMETER IN ENGINE INSTALLED WT LBM

LOX/JP4251500 TO 200028920002.8199 TO 226116 TO 13116340 TO24160

Figure 2 . Space Transportation Booster Engine ( S T B E ) .

PROPELLANTS NOZZLE AREA RATIO( S T O W E D / E X T E N D E D ) VACUUM THRUST KLBF VACUUM ISP SEC CHAMBER P R E SSU R E PSIA MIXTURE RATIO(O/F) LENGTH IN NOZZLE EXIT DIAMETER IN ENGINE INSTALLED WT LBM SEA LEVEL THRUST-KLBF( S TO W ED) SEA LEVEL ISP SEC FLOWRATE LB/SEC

LOX/LH255/150468/481449/46130066.0139/21976.2/126.37142397380.41P43.4

S T O W E D

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S T B E '-1.616M LBF- LOX/JP4 S T M E 4 8 1-481K LBF- LO X/LH2

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540 X 540 NMIP A Y L O A O O R B I T 98 X 2 2 S N M IPHASIN G O RBIT-10 X 175 NMIR E E N T R Y O R B I T

E V E N T TIME ~ H R S1 - LIFTO FF 0.0002 - I N J E C T 100 X 540 NMI 0.1463 - P A Y L O A D S E P A R A T I O N 0.600*4 - K IC K S T A G E C I R C U L A R IZ E S P A Y L O A D @ 1st APO GEE 0 .9515 - S E P A R A T E P/AM O D U L E F R O M S T A G E 2 2.4*6 - S T A G E 2 DEO RBIT N EAR 2NDAPOGEE 2.7*7 - STAGE 2 SPLASHDOWN 3.2*8 - P/A M O D U L E PHASIN G BURN @ 2ndPERIGEE AFTE R IN SERTIO N 3.3649 -P /A M O D U L E DEO RBIT BURN 11.090

10 - P/AMODULE LANDING 11.93' A P P R O X I M A T E

Figure 6 . Design r e f e r e n c e .mission profile.

A L T I T U D E -60.000 F TV E L O C I T Y -MACH 1

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Figure 15. H L L V propellant f e e d line layout.

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central avionics package has overall control of these auxiliary functions. In addition,there is an Auxiliary Subsystems Processor in each -booster, w h i c h assumes controlof the booster after separation f rom the m a i n vehicle.Electrical Power Subsystem

Independent power sources are located in the P/A M o d u l e central avionicspackage and in each booster. The P/A M o d u l e power source consists of f u e l cellsand/or batteries. Batteries are used for the booster power source and to providepower for the deboost system on the core stage. The power distribution and controlsubsystem is s h o w n in Figure B-13. The possibility of electromechanical thrustvector actuators in lieu of hydraulic actuators will be evaluated for i m p a c t on the# electrical power system design.

Concepts for highly distributed power sources wil l be evaluated in the tradestudies. These concepts w i l l e m p l o y a f a m i l y of high energy density, long-lifebatteries w h i c h can be distributed and optimized to local load requirements. Thef a m i l y w i l l include both primary and rechargeable batteries and special purpose .batteries for applications such as high power, very short duration loads. The dis-t tributed source concept m i n i m i z e s weight and complexity of the power transmissionsystem and enables partitioning to m a t c h vehicle modularity and maintenance require-ments. Concepts wil l be chosen to m i n i m i z e cost and t i m e of maintenance and refur-bishment between missions.

Distributed power system concepts will be evaluated to: (1) determine i m p a c tto avionics data interfaces, (2) maintain an acceptable E M I / I M C environment throughuse of a hybrid grounding scheme, and (3) determine the performance and charac-teristics of the energy storage devices. Information gained will be used to identifythe technology status and needs as they relate to the electrical power subsystem.For larger reusable vehicles such as the H L L V , that have greater amounts ofbuilt-in test, autonomous systems, and redundancy, the energy and power needs wil lincrease significantly. The use of a distributed power system to m e e t these needsm a y result in a m o r e s i m p l i f i e d electrical power system with higher reliability and theability to f u l l y utilize emerging technologies such as lithium and sodium sulphurbatteries.

Range S a f e t y SystemThe Range Safety Subsystem is shown in Figure B-14. Each m a j o r vehiclee l e m e n t contains a completely independent system. Cross strapping is provided asan additional assurance of c o m p l e t e vehicle destruction if necessary. W i t h the

exception of the cross strapping feature, the range safety system shown is essentiallythe s a m e as f l o w n currently on the STS.H L L V Avionics Studies

The foregoing description of desired attributes and characteristics for an H L L Vavionics system has been formulated without f u l l benefit of a comprehensive set ofstudies, and should be treated as a reference for comparison. There are m a n y areasopen for trades, analysis, and further definition, that must be resolved very earlybefore c o m m i t t i n g large expenditures to an advanced avionics development. A numbero f these areas for further study are identified in Table B-3.5 3


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RANGESAFETY

RECOVERYCOMM & TRACK

Figure B - l . H L L V avionics interfaces.T A B L E B - l. A V I O N I C S F U N C T I O N A L R E Q U I R E M E N T S

GUIDANCE N A V I G A T I O N , C O N T R O L- A N D FLIGHT S EQ UENC I NGS A F E T Y C O N T R O L

S E C U R I T YVEHICLE/PAYLOADS E R V I C E S

F A U L T D E T E C T I O N / I D E N T I F I C A T I O N . A N DR E C O V E R YF L I G H T I N S T R U M E N T A T I O N

O N B O A R D C H E C K O U T A N D L A U N C H

- P A Y L O A D D E L I V E R Y- R E T U R N / R E C O V E R Y- R A N G E S A F E T Y- C H E C K O U T / L A U N C H O P E R A T I O N S S A F E T Y- V E H I C L E /P A Y L O A D S A F E T Y- I N T E R N A L / E X T E R N A L C O M M U N I C A T I O N S- POST LANDING/SPLASHDOWN- S Y S T E M I NTEG RA TI O N- DATA PROCESSING- E L E C T R I C A L POWER- V E H I C L E S U B S YS T E M F A U L T S

R E A L T I M E F L I G H T S T A T U SE N V I R O N M E N T / D A M A G E A S S E S S M E N TT R E N D A N A L Y S I SP R O B L E M I N V E S T I G A T I O NO P E R A T I O N A L R E A D I N E S S V E R I F I C A T I O NC O M P O N E N T / S U B S Y S T E M S V E R I F I C A T I O NS Y S T E M I N T E G R A T I O N V E R I F I C A T I O NFLIGHT SIMULATION & C O U N T D O W ND E M O N S T R A T I O NC O U N T D O W N AN D LAUNCH

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T A B L E B - 2 . A V I O N I C S T E C H N O L O G Y N E E D ST E C H N O L O G Y N E E D

I K N O W L E D G E B A S E D E X P E R TS Y S T E M SA P P L I C A T I O N

H I G H L Y A U T O M A T E D G R O U N DSUPPO RT S Y S T E M S F O R :- M I S S I O N PLANNINGAND T A R G E T I N G- S O F T W A R E DE V E L O P M E N TA N D V E R I F I C A T I O N

R A T I O N A L E R E D U C E S L A B O R AN D TIME BE-T W E E N MISS IONS9 R E D U C E S R E A L T I M E S U P PO R TP E R S O N N E L E N A B L E S R A P I D P R O B L E MD I A G N O S I S A N D D E C I S I O N M A KI N G

F A U L T T O L E R A N T D I S T R I B U T E DP R O C E S S I N G

A U T O N O M O U S C H E C K O U T

I M P R O V E D I N S T R U M E N T A T I O NO A D A P T I V E G U I D A N C E

A D A P T I V E CONTROL P R E C I S I O N R E E N T R Y CN&C

9 HIGH P O W E R E L E C T R O -M E C H A N I C A L A C T U A T O R S I M P R O V E D S E N S O R SO LONG-LIFE. H I G H - P O W E R

L O W M A I N T E N A N C E ,P R O P E L LA N T G R A D E F U E L C E L L

F A M I L Y OF HIGH E N E R G YDENSITY, LONG-LIFEB A T T E R I E S

- V E H I C L E C H E C K O U T ANDL A U N C H- MISSION O P E R A T I O N S

V E R Y R E L I A B L E M O D U L A RP R O C E S S I N G S Y S T E M F O R :- V E H I C L E A V IO N I C S- C H E C K O U T AND L A U N C HS U P P O R THIGH D E G R E E O F ON-BOAROC H E C K O U T AND SELF-TEST

D E T E R M I N A T I O N O F C O N D I T I O NA N D P E R F O R M A N C E O F E N G I N E S ,E T C . O P T I M A L R E T A R G E T I N GD U R I N G F L I G H T O R J U S TP R I O R T O L A U N C H

> DETERMINATION OF OPTIMALC O N T R O L PA R A M E T E R SDURING FLIGHT

P R E C I S I O N T A R G E T I N GF O R R E T U R N O F R E C O V E R A B L EM O D U L E S

MAIN E N G I N E T H R U S T V E C T O RC O N T R O L A T T I T U D E , A T T I T U D E R A T E S ,A C C E L E R A T I O N , E T C .

C E N T R A L P O W E R S O U R C E F O RS E V E R A L H O U R S M I S S I O ND U R A T I O N

I D I S T R I B U T E D S O U R C E S O P T I M I Z E D& D E D J C A T E D TO L O C A L L O A DR E Q U I R E M E N T S- P R I M A R Y B A T T E R I E S- R E C H A R G E A B L E B A T T E R I E S- B A T T E R I E S F O R H I G H P O W E R ,V E R Y S H O R T D U R A T I O N L O A D S

E N A B L E S P A R T I T I O N I N G OFP R O C E S S I N G T O F I T V E H I C L EM O D U L A R I T Y

I M P R O V E S S A F E T Y A N DM I S S IO N R E L I A B I L I T Y E N A B L E S R A P I D C H E C K O U TAND L A U N C H R E D U C E S G R O U N D E Q U I P M E N TAND C A B L I N G C O N N E C T I O N S I M P R O V E S A S S E S S M E N T O F W E A R ,D A M A G E , A N D F L I G H T W O R T H I -N E S S O F R E U S A B L E H A R D W A R ER E D U C E S / E L I M I N A T E SP R E - F L I G H T S IM U L A T I O N A L L O W S M IS S IO N R E T A R G E T -I N G A T A N Y T I M EA C C O M M O D A T E S OFF-NOMINALE V E N T S S U C H A S E N G I N E

O U T , E N G I N E P E R F O R M A N C EA T M O S P H E R I C P E R T U R B A T I O N S .C . G. E R R O R S , E T C . R E D U C E S G R O U N D O P E R A T I O N S

E N H A N C E S R E U S A B I L I T Y E L I M I N A T E S NEEFORAPU ANDH Y D R A U L I C S I M P R O V E D L O N G T E R M S T A B I L I T YA N D R E L I A B . T O A C C O M M O D A T EL A U N C H ON D E M A N D W I T H MINIMAlR E C A L I B R A T I O N A N D C H E C K O U T U T I L I ZE S P R O P E L L A N T B O IL O F F

MI NI MI ZES M A I N T E N A N C E ANDREFURBI S H MENT B E T W E E N MI SSI O NS MIN IM IZ E S W E I G H T ANDC O M -PLEXITY OF P O W E R TRA NS MI S S I O NS Y S T E M E N A B L E S PARTITIONING TOFITV E H I C L E O F P O W E R T R A N S M I S S I O N

I M I N I M I Z E S M A I N T E N A N C E AN DR E F U R BI S H M E N T B E T W E E NMISS IONS.

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T A B L E B-3. A V I O N I C S S T U D I E S

TECHNOLOGY UTILIZATION AND ADVANCED DEVELOPMENT PLANNING FAULT TOLERANCE DEGREE ANDMETHODOLOGY

HARDWARESOFTWARE INNOVATIVE CONCEPTS FOR MINIMIZING TURN-AROUND COSTS AND

TIMEHARDWARESOFTWAREOPERATIONS

DEGREEOFREUSABILITY DEGREEOF ONBOARD CHECKOUT AUTOMATION9 UTILIZATION OF KNOWLEDGE BASED EXPERT SYSTEMS

FOR CHECKOUT AND LAUNCH SUPPORTFOR MISSION PLANNING AND TARGETINGFOR SOFTWARE DEVELOPMENT

PARTS/COMPONENTS RELIABI LITY APPROACH DEFINITION OFSOFTWARE DEVELOPMENT AND SUPPORT ENVIRONMENT DEGREEOF COMMONALITY AND STANDARDIZATION

FLIGHT HARDWARE AND SOFTWAREDEVELOPMENT AND OPERATIONAL SUPPORT SYSTEMS

SYSTEMS/SUBSYSTEMS DESIGN TRADES AND ANALYSESDATA PROCESSING SUBSYSTEM ARCHITECTURESOFTWARE SIZING ANALYSISELECTRICAL POWER SOURCES & DISTRIBUTION CONCEPTSELECTRICAL VS. HYDRAULIC THRUST VECTOR ACTUATORSG&N SENSOR COMPLEMENT- GUIDANCE METHODOLOGYGN&C PERFORMANCE/ACCURACY ANALYSISCONTROL LAWS, AND SENSORS

56


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c P O O R AuHiC U R R E N T S T 5 A V I O N I C ST E C H N O L O G Y IS 10 TO 15 V E A H S O LDIMITS. GPC'S . & M O M ' S O U T O F P R O D U C T I O NNON-STANDARD D A T A B U S P R O T O C O L SL I M I T E D F A U L T D E T E C T I O N A N D' R E D U N D A N C Y M A N A G E M E N TH A L - S S O F T W A R ES L O W & C O S T L Y M I S S I O N P L A N N I N GA N D R E C O N F I G U R A T I O N

C U R R E N T & E V O L V I N G T E C H N O L O G Y D I S T R I B U T E D P R O C E S S I N G P R O C E S S I N G S P E E D 8,C A P A C I T Y E N E R G Y S T O R A G E C A P A C I T Y G N & C S EN S O R A C C U R A C Y & R E L I A B I L I T Y F A U L T T O L E R A N T S Y S T E M S ADA S O F T W A R E K N O W L E D G E B A S E D S Y S T E M S( A R T I F I C I A L I N T E L L I G E N C E !

U P G R A D E D S T S A V I O N I C S R E P L A C E M E N T O F OUT-OF-PRODUCTIONC O M P O N E N T S I M P R O V E D A U T O M A T I O N O F M IS S I O N

P L A N N I N G A N D R E C O N F I G U R A T I O N

2 N D G E N E R A T I O N S T S A V IO N I C S' U T I L I Z A T I O N O F E V O L V I N G T E C H N O L O G Y> S O M E H E R I T A G E F R O M P A S T

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- F A U L T T O L E R A N C E- O N - B O A R D C H E C K O U T- A U T O M A T E D C/O AND LAUNCH- A U T O M A T E D M I S S IO N P L A N N I N GA N D R E C O N F I G U R A T I O N

Figure B-2. Avionics/software technology.

L O C A LN E T W O R K

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F T P - F A U L T T O L E R A N T PR O C E S S O RDID - D E V I C E I N T E R F A C E UNITN N E T W O R K N O DE

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A P P E N D IX H . S T R U C T U R A L A N A L Y S I S

The selected launch vehicle consists of a four-element booster and a core secondstage. The booster is composed of two 171 in. tank diameter elements w i t h one LOX/hydrocarbon engine and two 2 4 6 - i n . tank diameter elements w i t h two L O X / H y d r oc a r b o nengines. The second stage is a 396-in. diameter L O X / L H 2 stage powered by arecoverable Propulsion/Avionics (P /A) M o d u l e w i t h five L O X / L H . ^ engines. A l l enginesare pad ignited. Propellants for the second stage burn are crossfed f r o m the boostere l e m e n t s so the core stage is f u l l w h e n the first stage is jettisoned. Structuralintegration components are a forward adapter cone and booster attach ring w i t h aftstruts connecting the booster elements to the second stage.

A . LoadsThe H L L V load characteristics were determined using trends and load analysis

techniques developed on the Saturn and Space Shuttle programs. An early IOC wasdesired and new technology was not used in the design and materials selection.These may be studied using this design as a point of departure for improvements.A factor of safety of 1.25 was used on structural elements. Prelaunch on-pad, lift-off,m a x i m u m dynamic pressure (max q), max q times angle-of-attack (max q*alpha),m a x i m u m acceleration, staging, and P/A M o d u l e land i m p a c t recovery events wereanalyzed or considered in sizing the vehicle components. Structural design loadswere generated at the trajectory point of max q*alpha assuming an aerodynamic angle-of-attack of 6 deg and all six booster engines gimballed 6 deg.Aerodynamic Loads

The m a x i m u m shear loads were derived for the two worst conditions; first atl i f t - o f f prelease, and second at max q*alpha. The worst condition was determined tobe the latter with a three s i g m a wind profile.M a x i m u m Bending M o m e n t

The m a x i m u m bending m o m e n t distribution was calculated at the max q*alphacondition and based on aerodynamic load distribution with an angle-of-attack of 6 degand all six booster engines gimballed 6 deg. The combination of these two separateloads yield the m a x i m u m l i m i t structural bending m o m e n t at max q*alpha as shown inFigure H-l. The longitudinal acceleration history used for loads calculations is shownin Figure H-2.The booster load paths are s h o w n in Figures H-3 and H-4. The forward attach-m e n t , designed into the large, adapter ring .plus the aft attach stru ts, cause the bend-ing loads to be reacted by the booster elements rather than the second stage. Thispenalty in booster weight is m o r e than o f f s e t by the reduction in second stage tankweight.

>

3. BoostersThe single engine boosters are located in the pitch plane and will react a largepercentage of the vehicle pitch and s o me of the yaw bending m o m e n t s . The two

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engine boosters are m o u n t e d in the yaw plane, and react a large percentage of they a w a n d a s ma l l a m o u n t o f t h e pitch bending m o m e n t s . 'The booster forward adapters are unique structures and can be described asoblique cones, similar in shape to a "duckbill" (Fig. H-5). These are designed asa l u m i n u m skin and stringer shells that carry m o s t of the loads in the stringers w h i c h

are straight columns tied together by the skin and f r a m e s . The 5 0 - f t diameter struc-tural ring, that combines w i t h the second stage adapter and the payloaci aerodynamicfairing base to integrate the components into a launch vehicle, is e m b e d d e d in the"duckbill" section of each booster. Figure H-6 displays the f o r w a r d integration hard-ware detailing the individual components and separation planes. .The booster stagesections of the ring f o r m a torque box/ring f r a m e that separates f r o m the uppersection of the ring at booster burnout. The structural w e i g h t estimates for theboosters are detailed in Tables H-l and H-2.Propellant Tanks .

The propellant tank w a l l s m u s t be designed to resist the sum of the folicwir.; . .loads due to supporting mass, drag loads, bending compressive loads, engine thru.=; r.loads, and internal tank pressure loads. Using the worst case at max q*alpha. therequired w a l l thickness and stiffness are established. All tanks are designed ofintegrally stiffened construction, 2219 a l u m i n u m alloy with stabilizing rings that serveas both prime structure and slosh baffles. The skin thickness is tapered accordingto pressure and compressive load requirements at all positions along the stage. Thelongerons and ribs are integral parts of numerically machined plates, f o r m e d to thecorrect radius and welded into the cylindrical sections. The welded ring f r a m e s arespaced for o p t i m u m stability.The propellant tank domes are elliptical "square root of two" geometry. Exceptfor the c o m m o n bulkhead between the LOX and LH2 tanks, all domes are m a d e by

mechanical and chemical milling of 2219a l u m i n u m plate, f o r m e d into gore sections,and welded into a complete unit. The c o m m o n bulkhead is of honeycomb constructionwith a l u m i n u m f a c e sheets and a thermal barrier core. A phenolic core was selectedfor this design but trade studies may result in a m o r e o p t i m u m material or concept.A "Y" ring is welded to the outer edge of each d o m e to allow connection to eitherthe propellant tank w a l l or the intertank sections.

The intertank sections are non-pressurized load carrying cylindrical sectionsbetween tanks. They are designed as corrugated a l u m i n u m alloy w i t h rings spacedfor stability.Thrust Structure

The thrust structure consists of a l u m i n u m box beams that support and reactthe engine thrust (Figs. H-7 and H-8). The thrust structure is supported by acylindrical skin and stringer skirt with stabilizing ring frames. The stiffness wasdesigned to m i n i m i z e b e a m deflection during normal operation and engine out condi-tions. The a l u m i n u m cylindrical section and ring f r a m e s between the thrust structureand the f u e l - t a n k was sized to react Phe loads during all operating conditions,including transportation, on-pad stackup, prelaunch, and all flight environments.Attachments for the aft struts to react the kick loads to the second stage and postsfor pad erection and holddown are located on the thrust structure ring (Figs. H-7,H - 8 , and H-9).

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The booster aft skirt provides aerodynamic and t h e r m a l protection to the boosterengine(s), thrust vector control s y s t e m , and recovery system. Ring f r a m e s are addedto provide shear support for the a t t a c h m e n t of the aerodynamic fins.A e r o d y n a m i c Fins

The aerodynamic fins are required to provide directional stability duringatmospheric flight and reduce the required engine gimbal angle due 'to external dis-turbances. The all a l u m i n u m structure was sized for stiffness to react the worstw i n d shear experienced by any fin dui'ing ascent flight. Each fin is attached to thebooster by 16 explosive bolts required for jettisoning f r o m the booster after burnout.The structural w e i g h t e s t i m a t e s for the boosters are detailed in Tables H-l and H-2.C. Second Stage

The second stage consists of a 3 3 - f t diameter L O X / L H . - , vehicle, nestled in thecenter of the f o u r booster stages. It is composed of a conical f o r w a r d adapter /'payloada t t a c h ring, cryogenic propellant tanks, aft skirt, and a recoverable propulsion/avionics m o d u l e .Conical Adapter

The conical adapter is a structural transition between the 3 3 - f t core stage tothe 50-ft diameter booster attach ring/payload fairing adapter. The cone is ana l u m i n u m skin and stringer assembly w i t h stabilizing ring frames. It is designed totake all first stage thrust tension loads m i n u s the second stage thrust compressiveloads, stage inertia loads, and bending m o m e n t s . These loads are s u m m e d to estab-lish the m a x i m u m conditions 'for both compressive and tension loads. The worstcompression load condition wil l be during l i f t o f f and the m a x i m u m tension load w i l loccur at m a x i m u m booster acceleration. The adapter components are designed forthe worst condition and therefore, can react any other flight loads.The f o r w a r d end of the conical adapter is a 5 0 - f t ring w h i c h is an integralpart of the payload fairing base as shown in Figure H-6. The l o w e r half of this ringis f o r m e d f rom the booster forward adapters and the upper half from the payloada t t a c h m e n t and aerodynamic shroud a t t a c h m e n t . The upper half is a torque box,25-in. high and 2 4 - i n . w i d e . The payioad lairing is attached to the ring and isjettisoned by pyrotechnic devices during second stage flight.

Propellant TanksThe propellant tanks of the second stage are 33-ft in diameter with a c o m m o n

bulkhead between the L O X / L H 2 (Fig. H-10). All d o m e s are elliptical "square-root-of-two" geometry. The tanks are designed'of 2219a l u m i n u m alloy, integrally stiffenedconstruction w i t h stabilizing ring f r a m e s that also serve as slosh baffles. The basicm e m b r a n e for both tanks tapers in thickness according to strength requirements.The longeron skin and ribs are numerically machined from thick plate, then formedto be w e l d e d into a cylinder.The aft d o m e of.the LH~ tank is sized for the m a x i m u m ullage pressure of

24 psi plus the pressure head due to m a x i m u m longitudinal acceleration during boosterburn. The c o m m o n bulkhead is designed of honeycomb construction w i t h a l u m i n u m

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face sheets and phenolic core for the t h e r m a l barrier. The LOX tank is designedfor 27 psi ullage and the pressure head due to m a x i m u m longitudinal accelerationburn. "Y" rings are used to connect the tank sections with the forward and aftskirt sections.The forward skirt is a 7-ft long cylinder of a l u m i n u m skin and stringer designw i t h three stabilizing ring f r a m e s . It provides the transition between the LOXtankand the conical adapter.

A f t SkirtThe second stage aft skirt is a cylindrical section f o r m i n g the transitionbetween the LH2 tank and the P/A module. It distributes all the P/A m o d u l e longi-

tudinal loads and the booster kick loads to the second stage (Fig. H -9) . The boosierstrut attach ring is a welded a l u m i n u m assembly w i t h four strut attach fittings.designed to withstand the m a x i m u m loads resulting f r o m reaction forces of enginegimbaling and vehicle dynamics. ' These loads'are introduced tangentially t h v c u g ' hthe strut attachments.The thrust loads f r o m the P/A M o d u l e are transferred to the aft skirt at f o u r .hard points f r o m the P/A M o d u l e thrust structure. These attach points m u s t distributethrust loads evenly to the second stage tanks. Table H-3 is a detailed structuralweight s u m m a r y of the second stage.

D. Propulsion/Avionics M o d u l eThe recoverable P/A M o d u l e is the heart of the H L L V . It houses the propulsionsystem composed of m a i n L O X / L H 2 engines w i t h two position nozzles, orbital maneuver-

ing and attitude control engines (Appendix F), and m a j o r avionics systems (Appendix .B). The geometric design of the P/A M o d u l e is a 33-ft diameter spherical d o m e w i t ha 33-ft radius and a conical skirt which tapers to 41.7-ft diameter at its base (Fig.H-l l ) . The forward section of the conical section attaches to the thrust structure.

The forward d o m e is designed as an a l u m i n u m honeycomb structure capable ofwithstanding the high aerodynamic pressure of reentry and absorbing the. groundi m p a c t loads at landing. Several honeycomb f a c e plates *.vere investigated '.vithdifferent core thicknesses. It was determined, due to critical pressure, that thebasic honeycomb core should be 5.2 in. thick w i t h the outside face plate of 0.06 in.and the inner face plate 0.04 in. To reduce the nose cap weight, the .honeycomb corewas sized of two densities. To absorb the crushing loads of land impact, the center310-ft diameter was designed using a density of 12 I b / f t . The remaining outer ringoused a 6 I b / f t density because of lighter loads. The spherical cap, including TPS,is considered expendable, i.e. , removable f r o m the basic structure and replaced afterevery flight.The conical section was designed of isogrid construction to withstand theexternal reentry dynamic pressure. The thrust structure attaches directly to theforward conical section. Four thrust post fittings distribute the shear and com-pressive load to the second stage. These load attach points must be. designed insuch a way that they can be stowed and protected during reentry. Ring frames forstiffness have been designed into the lower portion of the conical section which could

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be used for attaching thermal curtains. Other m a i n f r a m e rings are located at thea f t of the c r u c i f o r m thrust structure to react the m o m e n t s introduced by the enginegimbaling. OMS and RCS engines are mounted on removable panels located on eachquadrant at the aft end of the conical skirt.The thrust structure consists of all a l u m i n u m box b e a m s that f o r m a cruciform

b e a m assembly that reacts the engine thrust and g i m b a l loads. The box beams are52-in. high and 16-in. w i d e and designed for stiffness to m i n i m i z e deflection. Thea l u m i n u m box b e a m design was selected as a m e t h o d of meeting an early IOC date,realizing that other materials and structure 'designs are available. Table H-4 s u m m a r i z e sthe detailed structural w e i g h t , s u m m a r y of the P/A M o d u l e .E. Aerodynamic Payload Fairing

The aerodynamic payload fairing (PLF) protects the payload during prelaunchand atmospheric flight. The PLF is jettisoned during second stage flight a f t e r thed y n a m i c pressure is low enough to preclude damage to the payload f rom atmosphericheating. The fairing wil l have access ports for support services during ground pro-cessing. The aerodynamic, axial, bending, and shear loads were determined forseveral transient and quasi-steady-state conditions. The m a x i m u m loading occursfrom the combination of these loads at max q*alpha. The m a x i m u m bending m o m e n tof 42 mill ion ft-lb occurs at the fairing base.Payload Fairing Adapter Ring

This 4 0 - f t diameter ring is part .of the second stage adapter ring. It consistsof a bolted angle or machined flange that will be the base of the PLF. The fairingwil l be bolted or clamped to the base which wil l be the separation plane.The basic cylindrical shell is 50 ft in diameter and 168-ft long as shown in

Figure H-12. The shell is divided into seven longitudinal segments of 24 ft each,allowing for variable length payload fairings. The three forward segments are notheavily loaded. The l imit bending m o m e n t at the forward segment is 4 mill ion ft-lb.A t the aft end of the third segment, the l i m i t bending m o m e n t is 10 mill ion ft-lb w i t ha shear of 1 2 0 , 0 0 0 Ib. The three forward segments are designed as a l u m i n u m honey-c o m b . The aft segments consist of a l u m i n u m skin and stringers w i t h stabilizing ringf r a m e s . The f o r w a r d s e g m e n t honeycomb has f a c e sheets 0 . 0 6 5 in. thick and a 2-in.3thick a l u m i n u m core w i t h a density of 2.5 I b / f t .. The aft third segment has faceosheets 0 . 0 9 0 in. and a core density of 3.5 I b / f t . The stabilizing rings for thesesegments are "Z" sections, 12-in. w i d e w i t h 4-in. flanges, located every 4 ft. Thefour aft segments of the fairing are designed of a l u m i n u m skin and stringer construc-tion w i t h stabilizing ring frames. The aft segment is the m o s t heavily loaded w i t h al i mi t bending m o m e n t load of 42 mill ion ft-lb at the-base. The l i m i t shear load atthis location is 4 2 0 , 0 0 0 Ib. The aft segment skin is 0.090-in. thick w i t h stringer andrib spacing of 2 and 24 in. apart, respectively. Both the stringer and rib areas2are 0.21 in. each. The stabilizing "hat section" ring f r a m e s wil l be located every4 ft. -. tThe PLF f o r w a r d end is a double angle nose cone developed at Marshall SpaceFlight Center. The frustrum cone, 31.3 ft long w i t h a 12.5 deg angle, is designedw i t h a 2-in. thick honeycomb having 0.040-in. thick face sheets. The core density

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o.is 2 I b / f t . The shell is designed to withstand the external pressure introduced byshock waves. Four stabilizing rings are included in the frustrum design. The for-ward cone is 3 7 . 4 - f t long w i t h a 25 deg angle. The design is a l u m i n u m skin andstringer with stabilizing ring f r a m e s . The skin thickness is 0 . 0 4 2 in. and the 9stringers are at 2-in. spacing. The stringer required cross section area is 0.11 in."2The rib spacing is 40 in. and has a required cross section area of 1.0 in. . Eightstabilizing rings are required. The stabilizing rings are "Z" sections 12-in. wide,having 34-in. flanges, spaced every 4 ft. Table H-5 lists the structural weightestimates for the PLF.

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,. /STABILITY RINGS(1\ / 24 I N CH E S A P A R T

F I R S T S T A G E S T R U TI N T E R C O N N E C T

THRU S T BEA MRINGST H R U S TS T R U C T U R E

Figure H-8. Two engine booster structure and tank adapter skirt,

FW D R I N G T OLH2 T A N K A D A P T E R

S T A B I L I T Y R I N G S2 4 I N C H E S P A R T

B O O S T E R S T R U TA T T A C H M E N T

Figure H-9. Structural description of second stage aft adapter skirt.117


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600

I1701680

986

284_L

396DIMENSIONS IN INCHES

Figure H-10. Second stage configuration.396 R

396

12R

| 1 2 .5 1 5 D E G229

500

DIMENSIONS IN I N C H E S

4 IN X 4 IN X 1/2

. C R O S S SEAMS E C T I O N O FT H R U S TS T R U C T U R E

S E C T I O N A-A

52

4 - 1 -Figure H-ll. Propulsion/avionics module.

118


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THERMALACOUSTICBLANKETSSECTION A-A

OPTIONAL INTERIORTHERMA L ACOUSTICINSULATION BLANKETS

Figure H-12. Aerodynamic payload fairing.T A B L E H-l . O N E E N G I N E B O O S T E R S T R U C T U R A L C O M P O N E N T W E I G H T S

ITEMF O R W A R D S K I R TC Y L I N D R I C A L S E C T I O N S K I R T F W DL H 2 F WD DO M ELH2 CYLINDRICAL DOMELH.2 COMMON DOMEL O X C Y L I N D R I C A L DO M EL O X A F T D O M EI N T E R T A N KRP FWD DO M ERP C Y L I N D R I C A L S E CT IO NRP A F T D O M EA FT A T T A C H . S T R U T S A F T S K I R T C Y L I N D R I C A L S E C T IO NT H R U S T S T R U C T U R EA F T S K I R TF INS

W E I G H T / L B S .52002100540

7350720

1639093550605404680640300066403300268010390

119


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T A B L E H - 2 . T W O E N G I N E B O O S T E R S T R U C T U R A L C O M P O N E N T W E I G H T SITEM . W E I G H T ~ LBS.F O R W A R D S K I R T (C O N I C A L S U R F A C E ) 4465C Y L I N D R I C A L S E C T I O N ( F W D . S K I R T ) 3790LH2 DOM E 1042L H 2 C Y L I N D R I C A L S EC T I ON 7065LH2/LOX C O M M O N D O M E 1 55 0LO X C Y L I N D R I C A L S E C T O R 16352L O X A F T DO ME 2680I N T E R T A N K 7 1 1 0RP FWD DOME 1042RP C Y L I N D R I C A L S E C T I O N 5 2 80RP A FT D O M E . 1620A F T S T R U T S 3 00 0C Y L I N D R I C A L S E C T IO N 8 06 0T H R U S T S T R U C T U R E 1 3136AFT S KIRT 3 880FINS 10390

T A B L E H - 3 . S E C O N D S T A G E S T R U C T U R A L C O M P O N E N T W E I G H T S

ITEM . W E I G H T ~ L B S .P/L F A I R I N G A D A P T E R R IN G 1 55 20S T A G E P A Y L O A D A D A P T E R 1 032 5FW D S K I R T 5 0 3 0L OX FW D DOME 4505L O X C Y L I N D R I C A L S E C T IO N 8 08 5LOX/LH2 CO M MO N DO ME 7970L H 2 C Y L I N D R I C A L S E C T I O N 3 41 00LH2 AFT DOME 3700A F T S KIRT 17950A F T A T T A C H S T R U T S F IT T IN G S 8 0 0

120


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T A B L E H-4: P / A M O D U L E S T R U C T U R A L C O M P O N E N T S W E I G H T S

ITEMTHRUST BEAMS 52 X 16 IN BOX ALUMINUMDOME-SPHERICAL CAP 6" THICK ALUMINUMHONEYCOMBCONE - SKIRT ISOGRID ALUMINUMUPPER CONE RINGL O W E R CONE RINGTHRUST BEAM RING L O W E RTHRUST BEAM UPPERTHRUST PO S T SBRACKETRYFASTENERS & MISCELLANEOUS

WEIGHT-LB896Q4760

5776532530712685252300100

T A B L E H-5. P A Y L O A D F A I R IN G S T RU C T U R A L C O M P O N E N T S W E I G H T S

NOOF ITEMSEGMENTS DESCRIPTION

TOTALWEIGHTPOUND

ONE/4 FWD CONE SEGMENTSFWD CONE FRAMES

89701220

ONE/4 FRUSTUM CONE SEGMENTSFRUSTUM CONE FRAME

99004800

FOUR/4 FWDCYLINDRICAL SEGMENTSFWD CYLINDRICAL SEGMENT FRAMES

2507012800

FOUR/4 AFT CYLINDRICAL SEGMENTSAFT CYLINDRICAL SEGMENTS FRAMES

4036018800

121


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A P P E N D I X I . T E S T STest Program

The m a j o r details of the test program are shown graphically in Figures 1-1through 1-6. V e r b a l discussion on upper level structure of program is given in thef o l l o w i n g paragraphs.

Structural Testing

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Heavy Lift Vehicles for 1995 and Beyond

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