Future Requirements and Applications for Orbital Transfer Vehicles

The Space Congress® Proceedings 1983 (20th) Space: The Next Twenty Years

Apr 1st, 8:00 AM

D. E. Charhut General Dynamics Convair Division, San Diego, California

W. J. Ketchum General Dynamics Convair Division, San Diego, California

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Scholarly Commons Citation Charhut, D. E. and Ketchum, W. J., "Future Requirements and Applications for Orbital Transfer Vehicles" (1983). The Space Congress® Proceedings. 1. https://commons.erau.edu/space-congress-proceedings/proceedings-1983-20th/session-iic/1

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]. FUTURE REQUIREMENTS AND APPLICATIONS FOR ORBITAL TRANSFER VEHICLES

By: D.E. Charhut* W.J. Ketchum**

General Dynamics Convair Division San Diego, California

ABSTRACT — Storage/reconstitution — Threat avoidance/defense will be enhanced The capability of the Space Shuttle I Sp and low payload High OTV performance (through high by use of the high-energy Centaur to provide propellants and maximizes etc.) and vehicle weight) minimizes transfer to higher orbits (geosynchronous, Hydrogen-oxygen is currently the highest transfer payload. for planetary escape missions. Future orbital performing chemical combustion rocket propulsion vehicles (OTV) requirements for NASA, military, and system (50% higher ISp than solids or storables). Cen commercial exploitation of space will require taur (Figure 1) is being incorporated into the Space improvements and technological developments such Transportation System (STS) to take advantage of as increased performance, increased reliability, and this in the near future at affordable cost, and methods increased mission versatility. Eventual OTV space to further improve performance and versatility are basing should offer further cost reductions through being considered. As part of the STS, Centaur mis vehicle reuse, freedom from Shuttle constraints, and sion assignments include Galileo, ISPM, and two for Force pro possible STS propellant recovery. DoD (Figure 2). Under a joint NASA/Air gram, NASA and the Air Force are sharing develop of the Centaur G (short version). NASA is This paper summarizes Centaur characteristics, per ment costs funding Centaur G ' (long version). formance, and program status and presents future into the considerations for orbital transfer vehicles OTV are expected to improve and opera Hydrogen-oxygen Space Station era, including their capabilities, endure for many years until noncombustion propul tional requirements, and the technology develop sion (chemical-electric) becomes available to effec ments required to make them a reality. tively remove Isp limits. INTRODUCTION CENTAUR

The goal of space transportation is to provide in Centaur is the world's first liquid-hydrogen-powered creased launch opportunity at lower cost. This paper space vehicle. Today, it is the United States' premier addresses orbital transfer vehicles (from low earth upper stage for launching large geosynchronous com orbit to higher orbits) and how improvements to this munications satellites, solar exploration spacecraft, segment of the space transportation system can and observatories to study the farthest reaches of contribute to this goal for a number of applications: space. • Commercial programs placement The flight-proven Centaur system has been launched — Satellite per — Satellite servicing 67 times on Atlas and Titan boosters. It has • NASA programs formed flawlessly during the last decade of operation, missions due in part to the improved guidance, navigation, and — Planetary It — Satellite placement electronics systems incorporated in the early 1970s. orbital operations is currently undergoing performance improvements —- Manned the • DoD programs for INTELSAT that will enhance its capabilities for — Satellite placement 1980s. stage Advanced Space Programs Integrating a modified Centaur high-energy • Director, a significant increase in •* Project Manager, Orbital Transfer Vehicles with the Space Shuttle offers

IIC-26 the high earth orbit and earth-escape performance craft. With on-orbit capabilities of the STS. Centaur rendezvous and assembly of the is the only affordable Centaur and spacecraft payload near-term solution for attaining this capability. weights of more than Figure 3 shows 20,000 pounds can be placed in geosynchronous Centaur performance for planetary equatorial orbit. and geosynchronous earth orbit missions, indicating its capability to perform the Galileo "direct" mission This approach allows the spacecraft as well as placement of heavy payloads at GEO for to use the full DoD and other users. 60-foot length of the cargo bay; however, it uses only about one-third of the lift capability of the Orbiter Studies were carrying the spacecraft. Adding a propulsion stage to begun in 1979 to integrate Centaur into the spacecraft the Space Shuttle using would use some of this excess capabil the current D-l configuration ity, and Centaur with 30,000 pounds of propellant could thus place even greater and cylindrical 10- spacecraft weights into geosynchronous foot-diameter tanks. Delay of the Galileo mission, orbit. however, increased energy requirements to the extent FUTURE REQUIREMENTS that additional propellant is required to perform the mission. A decision was made to increase the LH2 The basic tank diameter to the maximum allowed within the drivers for increased capability orbit cargo bay (14.2 feet), and to lengthen the existing transfer vehicles (OTV) include performance, opera LO2 tank by 2.5 feet (Figure 4). These modifications tions, and cost-effectiveness. Spacecraft growth from increase the usable propellant weight to 45,000 the current 5,000-pound range to 10,000 pounds pounds and and use the Shuttle cargo bay more efficient more is already beginning to happen. High velocities ly. The resulting configuration — Centaur G' — has are needed for planetary and maneuvering spacecraft. a vehicle length of 29.1 feet, with 30 feet of the Orbi- In the 1990s, missions are contemplated with round- ter cargo bay available to the spacecraft. trip capability to service satellites at GEO and others for manned By maintaining sortie. For this, OTV must fulfill the the Centaur D-l LO2 tank diameter, following requirements: the basic vehicle proplusion system remains un • Very changed, although the larger LH2 tank does require high performance: 15,000 to 30,000-pound an increased-diameter forward stub adapter and payloads, servicing and round trip. equipment module. • Low acceleration: maximize size of large space structures. For longer spacecraft, a shorter Centaur version uses • Reusability: reduce operating costs and return only 29,500 pounds of propellant at a slightly higher payloads. engine mixture ratio of 6:1 (the standard ratio is 5:1). • Aerobraking: double round-trip payload (versus This configuration, called Centaur G, is only 19.5 feet all propulsive). long and allows spacecraft up to 40 feet long in the Orbiter. • Manrating: Orbit transfer and sortie for crew modules. • Spacebasing: free from Shuttle Shuttle/Centaur G will have the geosynchronous constraints. capability to deliver large communications satellites Operations weighing 10,600 pounds with in space require quick reaction, restart for lengths up to 40 feet. orbit This is more than double the IUS capability, relocation, or low acceleration to transfer very with large, only five percent less spacecraft length. For heavier delicate spacecraft. Future space transportation spacecraft, Centaur J, with propellant tanks resized systems (Figure 5), including growth versions of ex to hold 35,000 pounds, will be capable of launching a isting or new vehicles, should be more cost effective. 14,000-pound spacecraft up to 37.6 feet long into Increased capability tends to lower the cost per pound geostationary orbits. of payload delivered to high orbit.

These capabilities will dramatically enhance the ENGINES United States' ability to launch large communication satellites. Without Shuttle/Centaur, the size of com Design studies and tests by the major engine contrac munication spacecraft would be limited until further performance tors indicated the feasibility of increasing hydrogen- improvements can be realized with oxygen Ariane or the Isp to 480 seconds (compared with 450 IUS, or until the United States can seconds afford to develop a new cryogenic currently). Figure 6 indicates several can OTV. didate engines. The limiting factor for Shuttle/Centaur geosyn Relatively small increases are not without chronous payload capability great conse is the maximum lift quence for such future missions as capability of the Orbiter: 65,000 GEO round trips. pounds. One way to For example, a 20-second Isp increase increase this capability is to use one Orbiter could double to launch the round trip payload of a shuttle-launched, Centaur and a second Orbiter to launch the space all- propulsive, reusable (ground-based) OTV.

IIC-27 the Air Force and NASA have LOW THRUST Studies performed for concluded that the use of a toroidal liquid oxygen engine mounted in the that are much larger and more complex tank with the main propulsion Spacecraft space provides the minimum length are being proposed (large geostationary communica central void separated tanks. The use of this ap tions and/or surveillance systems). These large OTV, assuming the potential for increasing the geosyn systems are stowed in the Orbiter cargo bay during proach offers capability (for a stage and ASE of and will require subsequent deployment and chronous payload launch length as Centaur G) to 17,000 checkout (Figures 7, 8, 9). the same installed pounds (Figure 10). Checkout before transfer to final destination orbit success by allowing malfunction Required toroidal tank technology development in maximizes mission of the spacecraft leaves the Orbiter. cludes control of propellant slosh, reduction corrections before control, load systems are designed for low loads residuals, propellant mixing for thermal Because these and manufacturing-assembly when deployed, it is necessary to limit acceleration destribution/support, methods. during transfer. This is accomplished with a low- thrust engine.

A new low-thrust engine offers high ISp, low weight, AEROBRAKING and small size, but requires technology developments payload pumps, long burn times, multiple starts, etc.). For round-trip missions to* GEO (manned, (small and The RL10 engine has demonstrated a low-thrust servicing or return, etc.), aerobraking (Figures 11 capability at a reduced ISp when running at extreme 12) offers twice the payload than all propulsive by us and off-design conditions. Solutions for feed system in ing atmospheric drag to dissipate kinetic energy stabilities have been devised, such as oxidizer heat ex reduce return propulsive AV by 50 percent. Pro- changer, cavitating venturi, etc. pellants for the last burn are eliminated and therefore do not have to be transported to GEO and b^k. Using multiple perigee burn trajectories for low- light thrust LEO to GEO transfer, the ideal velocity re Required technology development includes Improved quirements are about the same as for high-thrust weight, high temperature brake designs. tests are (two-burn) trajectories. analytical techniques and further wind tunnel needed to predict temperatures and loads in the tran Mission time for LEO to GEO is one to two days for sition and slip flow regimes. low thrust, multiple (9 to 17) perigee burns versus one-quarter day for high thrust (two burn). With effi SPACE-BASED OTV cient thermal control systems, minimal boiloff oc features of space could curs. A 1 kW fuel cell (sufficient for OTV needs) con The unique environmental fraction space-based OTV sumes less than 0.5 Ib/hr. Attitude control is also on permit a very high mass weight and cost advantages the order of 0.5 Ib/hr. There are negligible losses with significant payload systems: associated with engine start/stop since tank head idle over heavier ground based (burning) is used. Other losses such as leakage, etc., Advantages are negligible. • Free from Shuttle constraints (size, loads) • Reusable (lower cost) While multiple passes through the Van Alien belts in (mix and match capability) cur increased radiation dose, avionics systems are in • Modularity creasingly being hardened to withstand extended time Key Issues in space and to operate during peak solar activity • Long-term space exposure for multiple periods, etc. Therefore little penalty • Orbital integration, servicing through the Van Alien belts over one to two passes • Efficiency (low weight, high Isp) days is expected for vehicle or payload systems. • Low-cost operations (propellant delivery to LEO) • Deployment and retrieval TORUS TANKS • Future payloads and mission characteristics •;*• Technology needs vehicle, the As long as the shuttle is used as a launch (thin-gage) tanks (explosive forming, of single shut • Lightweight payload bay length will limit the utility advanced sub-minimum gage chemical milling) tle launches (payload and OTV together in cargo bay) • Lightweight (composite) structure unless the OTV can be made very short.

IIC-28 • Lightweight/high temperature aerobrake mater Required technology development includes rapid ials (fixed aerobrakes, adaptive contour control) rendezvous, propellant extraction, etc. The alter • Long life/space maintainable engine (modular native concept, a dedicated ET tanker, is a simpler, components, quick release fittings, fixed high e but probably more costly approach. Larger quantities nozzle) of propellants can be delivered for increased OTV • Low vapor pressure (2-5 psia) cryogenic pro- traffic. pellant management — thermal control (MLI in sulation, mixing, venting), propellant acquisition, SUMMARY gaging • Meteoroid and space debris protection (multilayer Advances in OTV are needed to enhance our nation's tanks, self-sealing tanks, space station hangar ability to operate effectively in space, particularly to facility, onboard space debris location, classifica reduce the overall cost of space operations for tion & avoidance system) transfer of increasing numbers and size of payloads to • Redundant, fault-tolerant, hardened avionics geostationary orbit. • Auto rendezvous/docking Because the costs and performance of the vehicles Figure 13 compares performance of ground-based necessary to perform space missions are directly and space-based OTV, showing that the potential related to the available technology, advanced devel high mass fraction of a space-based OTV design opment programs must be supported and sustained. could result in twice the payload of ground-based OTV (GEO payload placement, OTV-only return, all Liquid hydrogen-oxygen rocket propulsion systems, propulsive). An aerobrake is needed to maximize because of their high performance and versatility, are payload return/round trip missions, but the high essential. Increases in specific impulse, reductions in mass fraction of a space-based OTV mitigates the ad total system mass, and increased degrees of reusabili vantage of aerobraking for OTV-only return mis ty are also needed. While Centaur/Shuttle (IOC 1986) sions. evolution can satisfy near-term (1986-1995) objec tives, a space-based OTV is envisioned for the far- Reduced weight is achieved with low tank pressures term requirements. Advanced technology & low cost and low acceleration loads to permit very lightweight operations are essential. tanks and structure (Figure 14).

Usage of low vapor pressure propellants and an ad vanced space engine with low inlet pressure and REFERENCES NPSH requirements combine to reduce tankage skin gages. Use of a low-thrust engine reduces acceleration 1. "The Centaur loading on the tanks and further improves their effi Family," W.F. Rector III, General Dynamics/Convair, ciency. Technology development required includes San Diego, Ca. 33rd Con gress of The thin-wall tanks, low vapor pressure propellants, low International Astronautical Federa thrust-space maintainable engine, etc. tion, Sept 27-Oct 2, 1982, Paris, France. 2. "Orbital Transfer Vehicle Concept Definition Reduced cost is achieved with a reusable space-based Study," NAS8-33533, General Dynamics Convair OTV and low propellant refueling cost (Figures 15 Division, ASP-80-012, February 1981. and 16). 3. "Orbit Transfer Vehicle Servicing," NAS 8-35039, General Dynamics Convair Division In Two concepts have been proposed: The "Honeybee," terim Review, 12 January 1983 and the dedicated ET tanker. The "Honeybee" con 4. "Study of Space Station Needs, Attributes and cept calls for the OTV to leave the Space Station and Architectural Options," NASW 3682, General dock to the aft end of an ascending ET shortly after Dynamics Convair Division Midterm Briefing, 17 MECO. The OTV would then load itself with residual November 1982 propellants from the ET through a special docking port. The OTV would fire its engine during propellant 5. "Large Space System Cryogenic Deployment loading to settle propellants and possibly deorbit the System Study," General Dynamics Convair Divi ET. After separation, the OTV either returns to the sion, ASP-81-021. Space Station and off loads propellant to dewar 6. "Low Thrust Vehicle Concept Study," storage tanks or is immediately integrated with a NAS8-33527, General Dynamics Convair Divi payload for ascent to a higher orbit. sion, ASP-80-010, September 1980

IIC-29 Figure 1. Shuttle/Centaur.

Calendar year 1986 1987 1988 1989 Shuttle/Centaur Galileo T International Solar Polar T Venus Radar Mapper V V DoD T V V V V Geoplatform demonstration V Asteroid or comet V Commercial communication V V V V V V Saturn Orbiter/probe V 5 Total Centaur flights 2 5 6 T Firm v Potential and Figure 2. Potential Shuttle/Centaur missions for NASA, DoD, commercial applications.

Planetary missions Geosynchronous missions

25 65,000-lb Shu ttle lift limit ^s. ^

20 Storable Shuttle/Centaur | ; 7^-G 15 •G Injected Spacecraft J mass mass (1Q3|b) (1Q3|b) x/ 10 •<

Solar Polar (1986),

IUS

100 120 140 IDS Centaur 20 40 60 80 Storable Centaur Energy-C3 (KM^ per sec^) low-thrust Figure 3. Centaur dramatically enhances Shuttle capability.

IIC-30 Increase LH2 Lengthen tank& LH2 tank forward & forward structure structure diameter Lengthen 30ft 45-deg aft LO2 tank cone angle 24-deg 6:1 mixture aft cone ratio angle 19.5ft 5:1 mixture ratio

• Uses 15-ft diameter available in orbiter bay • Increases length available for spacecraft • Maintains LC>2 tank diameter & propulsion system unchanged Figure 4. Shuttle/Centaur is minimum modification to current Centaur. Ground-Based Space Based Centaur Initial 1986 Uprated 1993

Payload (Ib) 10-14K 10-25K (12K round trip) Payload (ft) 40-37 35-41 30 Engines Two RL1OA3-3A, phased-in low Single RL10 Cat MB Advanced propulsion system thrust & deployable skirt with idle mode man-rated

Avionics Current single-string phase- Triple-string fault- Advanced round trip in triple string tolerant, hardened

Tanks Wide body CRES 30-45K Increased capacity Separated tanks, ~ 10OK propellant capacity (<55K) propellant, orbital tanking

Structure Titanium adapters — phase-in Reduced boiloff, aerobrake May be dual stage, aerobrake composite adapters

Operations Expendable, SMM for Reusable, may use kick stage May be space based, on-orbit spacecraft survivability Orbital assembly assy, manned Figure 5. Potential OTV evolution.

I \ RL-10A3-3A RL-10 Cat MB Advanced Engine Low Thrust Manufacturer Pratt & Whitney Pratt & Whitney P&W, R/D, ALRC WD Status 160 successfully Phase B studies Conceptual studies flown component hardware Isp (sec) 448.4 460 475 to 480' 490* 15 1 0 to 20 CJU8 Thrust 15-16.5 (1,000 Ib) ( + 10% idle) ( + 10% idle) fiLffliii bum Development: ($M) none required 80 250 to 350 ISO (years) available 5 10

Figure 6. Candidate OTV engines*

IIC-31 • Platforms have significant economic advantage over individual satellites — Sharing of subsystems & structure — Built-in reliability & redundancy • LEO deployment/checkout of payload prior to transfer to GEO increases reliability — STS- or LEO-base for deployment/checkout — High energy, low thrust orbit transfer vehicle

• Limited automated revisit is beneficial (servicing) — Replenish consumables — Exchange predictable wearout components — Allow payload update £ growth Figure 7. Large space platforms (geostationary satellites). Figure 8. NASA LSS/Centaur.

Uses atmospheric drag to reduce GEO return AV by 7,000 fps, resulting in double the round trip payload of all propulsive OTV. OTV returning from GEO,

Brake stowed Graphite/polyimide rib/beams Flow stream Closure doors Offset eg

ply Nextel 312 ceramic fiber material Note: Main engine does not fire during aerobraking }. DoD LSS/OTV. Figure 11. Lifting brake OTV for GEO return missions (servicing, manned).

Centaur G'

Torus tank concept

Geo payload, tb Payload length, ft Thrust, Ib 30,000 30,000 500 Isp, sec 446.4 (MR=5.0) 437.2 (MR=6.3) 476.7 (MR=6) IOC 1986 1987 *65KSTS

Figure 10. Compact tankage (TORUS) maximizes payload length and weight.

IIC-32 Figure 12. Lifting aerobrake OTV.

Payload to GEO Reference Payload to GEO & return OTV only returns Expendable OTV Reference' (all propulsive) 20 20 * Expendable OTV Aerobraking {all propulsive) 18 -eturns) 18 16 16 14 14 GEO 12 GEO 12 Aerobrafcirig pay load, payload, (roundl Wp paytad} 1,000 Ib 10 1,000 Ib 10 8 8 6 6 4 4 ]:rwjm>d! lirip paytoadl 2

0.80 0.84 0.88 0.92 0.96 0.80 0.84 0,88 O.92 0.96 Mass fraction - Main Impulse Propellant Mass fraction Main: impulse Propefaitf Loaded OTV (w/o payload) Loaded QT¥ (w/o

OTV LO2/LH2 OTV AV assumed: Delivers 60K Ib (P/L + OTV + — 470 sec ISP OTV engine 14,000 up-all pmapufeiM' oir propellant) to LEO if ground based. aerotowraked, 2,000 Ib aerobrake 14,OOO bacfc-al! prapulswe- O Ground based (see refs) * Delivers 60K Ib (P/L + propellant 7.000 back-aerobraked D Space based: (reduction in dry + transfer system) if space based: weight — tanks & structure) "Ref: GD&ASP-8&O1IO OTV sized for this condition Figure 13. OTV performance.

Wp = 57,080 Ib Figure 14. Representative space-based OTVconcqtf*

IIC-33 Space station

Payload integration & servicing facility Recoverable engine pod Aerodynamic f airing x Avionics-

Deployable MLI Rendezvous & shield dock OTV docking port

Honeybee concept E.T. tanker concept per flight • 9-36 klb residuals recovered per STS flight • 214-220 klb delivered cost — % 235 $/lb • Propellant delivery cost — essentially free • Propellant delivery concept to support (no tanks in Orbiter) • Supplements Honeybee GEO • Supports 90-129 klb year to GEO 360 + klb year to Figure 15. Low-cost propellant delivery concepts.

/Ground-based reusable OTV All propulsive

SBOTV baseline Isp — 482 Usable prop, mass fraction 15 2 tank- .86 (.91)* 4 tank- .91 (.94)* *AII propulsive

GEO payload deliver 1 Q SBOTV cost All propulsive ($k/lb)

0 5 10 15 20 25 30 GEO payload weight (klb) Figure 16. OTV performance comparison.

IIC-34