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Paper Session I-A - Liquid Rocket Boosters for Shuttle The Space Congress® Proceedings 1989 (26th) Space - The New Generation Apr 25th, 2:00 PM Paper Session I-A - Liquid Rocket Boosters for Shuttle James E. Hughes Manager, LRB Studies, Marshall Space Flight Center, NASA Follow this and additional works at: https://commons.erau.edu/space-congress-proceedings Scholarly Commons Citation Hughes, James E., "Paper Session I-A - Liquid Rocket Boosters for Shuttle" (1989). The Space Congress® Proceedings. 8. https://commons.erau.edu/space-congress-proceedings/proceedings-1989-26th/april-25-1989/8 This Event is brought to you for free and open access by the Conferences at Scholarly Commons. It has been accepted for inclusion in The Space Congress® Proceedings by an authorized administrator of Scholarly Commons. For more information, please contact [email protected]. LIQUID ROCKET BOOSTERS FOR SHUTTLE James E. Hughes, Manager LRB Studies Marshall Space Flight Center, NASA ABSTRACT The Liquid Rocket Booster study was initiated vehicles, and a pressure fed system, once by NASA to define an alternative to the Solid referred to as the "Big Dumb Booster". The Rocket Boosters used on the STS. These prime study contractors, Martin Marietta Cor­ studies have involved MSFC, JSC and KSC poration and General Dynamics Space Sys­ and their contractors. The prime study con­ tems, were assisted considerably by the ef­ tractors, Martin Marietta Corporation and forts of Lockheed Space Operations Co. General Dynamics Space Systems, have (LSOC) at the Kennedy Space Center and identified Liquid Booster configurations which Lockheed Engineering and Sciences Co. would replace the SRB's in the Shuttle stack. (LESC) at Johnson Space Center, as well as wind tunnel testing at MSFC, and other sup­ The Liquid Rocket Booster increases Shuttle port. performance to 70K LBS, provides improved reliability, hold down and verification prior to The Liquid Rocket Booster was required by vehicle release, engine out and improved NASA Headquarters to provide the STS with abort capability, and is phased into the STS a payload capability of 70,500 pounds to a 160 launch operations without adversely affecting NM , 28 1/2 degree orbit with SSME power flight rate. level @ 104 % ( Rated Power Level ). An alternate case of 62,500 pounds to a 160 NM INTRODUCTION orbit, 28 1/2 degree orbit with SSME power The Challenger accident caused NASA to level @ 104 % was also required reconsider all possible options for STS boost (The larger 70K case covered both needs with propulsion, including the possibility of replac­ only a minor difference in Booster length for ing the Solid Rocket Booster (SRB) with a the diameters of interest). All STS require­ Liquid Rocket Booster (LRB). Accordingly, ments and specifications were to be met, and studies were initiated in October 1987 to iden­ impacts to the launch vehicle and facilities tify two Liquid Rocket Boosterconceptsforthe minimized. The most important requirement Shuttle. These concepts were to include a was to provide engine out capability, to im­ conventional liquid system using pump fed prove the abort capability of the entire Shuttle engines similar to those used on the Saturn system, and enhance the probability of safe 1-9 Orbiter and crew return in the event of major but later rejected due to the severe environ­ malfunction anywhere in the system. An addi­ mental and safety problems with these propel­ tional ground rule was adopted which said the lants in the large quantities required by the LRB would be able to be integrated into the STS. The metallized gels were also investi­ KSC ground processing flow without interrupt­ gated, but were rejected due to their limited ing the planned STS launch rate. experience and the open technology issues. The propellant combination that seemed to BOOSTER SIZING best fit the criteria at this point in the study Initial concerns in sizing the Liquid Booster activity was LOX / RP-1, since a diameter of centered around the fact that for all liquid approximatly 15 feet could be accomodated, propellants considered, the LRB diameter the attach points to the External Tank would exceeded the current SRB diameter of 1 2 feet, be in non - pressurized structure, and the pro­ since all liquid propellants are less dense than pellant combination is one with which NASA the solid propellant used in the SRB. and industry has a firm technology base and The Booster diameter is of critical significance considerable engine experience. to the Orbiter wing loads based on the results of initial wind tunnel tests at MSFC. These Propellant combinations using LOX and Hy­ tests were conducted with models scaled to drogen and LOX / Methane were thought ini­ represent diameters in the 12- 20 foot range, tially to require diameters in excess of what and lengths in the 150- 200 foot range based the early wind tunnel test data showed as on inital sizing performed by both contractors limits. Additional wind tunnel tests were per­ for a variety of propellant combinations. Early formed, in which the SRB protuberance ef­ results showed that booster diameters of 14- fects caused by the aft ET attach ring and the 16 feet were probably maximum to avoid IEA (Integrated Electronics Assembly) boxes unacceptable Orbiter wing loads. Also to be were evaluated. These results showed that considered were physical issues such as for­ protuberances had a much larger effect than ward and aft External Tank attach point loca­ was earlier believed; and, if the LRB could be tions, engine nozzle exit plane location , keep­ made without protuberances adjacent to the ing engine plumes within the flame trench Orbiter wing, that diameters of up to 18 feet limitations at pad 39 and VAB door clearence. were possible without violating wing load con­ straints. Due to the operational complications PROPULSION CONSIDERATIONS of another propellant at the launch pad ( RP- The only liquid propellant options which ap­ 1 ), the KSC personnel supported the LOX / peared to offer the total impulse required, and Hydrogen selection, since these propellants stay close to the dimensions of the SRB were are already used for the SSME'S. storable propellants and metallized gels. The As shown in Figure 1 "Range of LRB Con­ storable propellant combination cepts" , the propulsion options considered ( N2 04 / MMH ) was baselined by Martin both new and existing engines for the pump Marietta initially in orderto minimize diameter, fed concepts as well as a variety of propellant 1-10 RANGE OF LRB CONCEPTS Primary reasons for the selection of RP-1 are the density and the well known characteristics LRB PROPULSION OPTIONS in rocket engine use. The selection of LH2 1— 1 PRESSURE-FED 1 was primarily influenced by alternate applica­ «JEWENGINEl tions, and the ease of integration at the launch —1 » ———| L02/LH2 :] pad. Either of these options could be inte­ —I S L02/ —\ hb grated into the STS stack and provide major L02/C3H8 H L advantages over the SRB currently used. L02/C3H8/LH2 - -." METAL1ZED I RECOVER OR EXPEND ? An analysis of life cycle costs to develop and operate the recoverable and the expendable Liquid Rocket Booster showed a cost advan­ combinations for both pump and pressure fed tage for the recoverable case; however, the designs. Four engines per booster were answer is highly dependent on the assump­ baselined to provide "engine out" capability. tions used and the cost sensitivity to such Existing engines were rejected for various parameters as water impact damage (refur­ reasons - the SSME due to cost and produci- bishment % of new ), flight rate, engine cost, bility limitations, the AJ-23 due to the propel- salt water compatibility, mission model size, lant decision to avoid the highly toxic N2O4 / etc. [ Note: only ballistic boosters similarto the MMH, and the F-1 due to it's 1.5 M Ib thrust SRB are covered in this study - flyback options level where only two engines would be re­ are not considered.] The pressure fed system quired and would not allow for engine out would have very rugged tanks (rated at 600- capability. Propane was also considered as a 1000 psi) and would more nearly represent fuel, but the severe coking of injectors experi­ the SRB recovery, i.e. the entire stage recov­ enced in recently conducted technology tests ered by parachute. The pump fed system resulted in that fuel being dropped. Methane would have lower pressure tanks which could presented an attractive option for a fuel, and not survive the water impact loads; therefore, was investigated fully using the split expander recovery schemes for pump fed systems in­ cycle engine being considered in the concur­ volved only the main engine portion of the rent ALS propulsion studies by MSFC. Meth­ Booster, ie a Booster Recovery Module. ane was not selected since it offered no signifi­ Due to the relatively low engine unit cost fore­ cant benefits over hydrogen in sizing, and it's cast by the ALS engine contractors use as a rocket engine fuel is relatively new (approximatly $4M ea), the design impacts and unproven. of incorporating recovery, and the additional The selected fuel,by both contractors, for the front end funding required by the recoverable pressure fed booster was RP-1, and for the concept, the expendable mode was selected. pump fed, MMC selected RP-1 and General The penalty of error in the initial assumptions Dynamics, LH2. for the recoverable case was also a factor in 1-11 REUSABLE VS. EXPENDABLE LIFE CYCLE COST FLIGHT RATE SENSITIVITY REFURBISHMENT % SENSITIVITY 30000 _ 26000 EXPENDABLE 24000 20000 _ 22000 - REUSABLE LIFE CYCLE 20000 COST 18000 $INM 16000 14000 20 30 40 50 5 10 15 20 AVG ENGINE REFURB % STS FLIGHT RATE 12 18 24 30 36 AVG STAGE REFURB % FIGURE 2 establishing the expendable as the conserva­ optimum weight at a tank pressure of approxi- tive choice.
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