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Studies determined that shuttle optimized design, allowing the large space system and transfer vehicle in one shuttle flight, greatly reduces transportation costs and minimizes orbital operations. Careful attention to design resulted in efficient payload packaging. A minimum volume, high energy (liquid oxygen and liquid hydrogen) transfer vehicle is described that allows maximum volume for the payload in the orbiter cargo bay.
Studies are underway to design reusable orbital transfer vehicles that would be used to transfer payloads from low-earth orbit to higher orbits and return. One promising concept is to use an atmospheric pass on the return leg to reduce the amount of fuel for the mission. This paper discusses a six-degree-of-freedom simulation analysis for two configurations, a low-lift-to-drag ratio configuration and a medium-lift-to-drag ratio configuration using both a predictive guidance technique and an adaptive guidance technique. Both guidance schemes were evaluated using the 1962 standard atmosphere and three atmospheres that had been derived from three entries of the Space Shuttle. The predictive technique requires less reaction control system activity for both configurations, but because of the limited number of updates and because each update used the 1962 standard atmosphere, the adaptive technique produces more accurate exit conditions.
A reuseable orbit transfer vehicle concept was defined and subsequent recommendations for the design criteria of an advanced LO2/LH2 engine were presented. The major characteristics of the vehicle preliminary design include a low lift to drag aerocapture capability, main propulsion system failure criteria of fail operational/fail safe, and either two main engines with an attitude control system for backup or three main engines to meet the failure criteria. A maintenance and servicing approach was also established for the advanced vehicle and engine concepts. Design tradeoff study conclusions were based on the consideration of reliability, performance, life cycle costs, and mission flexibility.
An aerobrake for a lunar transfer vehicle was designed through the use of the Taguchi design method, structural finite element analyses, and structural sizing routines. A minimum weight aerobrake structural configuration was the objective of the study. Six design parameters were chosen to represent the aerobrake structural configuration. The design parameters included honeycomb core thickness, diameter-to-depth ratio, shape, material, number of concentric ring frames and number of radial frames. Each parameter had 3 levels. The optimum aerobrake configuration resulting from the study was approximately half the weight of the average of all the other experimental configurations. The parameters having the most significant impact on the aerobrake structural weight were identified.
The results of systems analysis conducted on low life drag ratio (L/D) aero-assisted orbit transfer vehicle (AOTV's) are presented. The objectives for this class of vehicle and formulate technology development plans and funding levels to bring the required technologies to readiness levels, as well as develop a credible decision data base encompassing the entire range of low L/D concepts for use in future NASA Aeroassist Orbit Transfer Vehicles studies. Each candidate low L/D concept, the aerobrake, the lifting brake, and the aeromaneuvering concept could be made to work with technologies achievable by the early 1990's. All concepts require flexible structure with flexible thermal protection system (TPS) to be successfully integrated into the shuttle orbiter for launch, all required improvements in guidance and control to fly the dispersed atmospheres at high altitude, and all concepts had potential to evolve from ground-based to space-based operations.
A new technique has been developed to enhance the use of upper atmosphere aerobraking for increased performance from orbital transfer vehicles. This technique utilizes a pressure supported drag brake and the orbital transfer vehicle main engine to modulate aerodynamic drag and also to alleviate the aerodynamic heating during a grazing pass through the atmosphere. Performance analyses of vehicles utilizing all-propulsive or aerobraking during round trip missions from low earth orbit (LEO) to geo-synchronous earth orbit (GEO) and back shows that aerobraking allows a given vehicle to deliver approximately twice as much payload to GEO and return. Aerobraking also provides more than twice the round trip payload.
Industry studies were undertaken to establish the technology base for an advanced engine for Orbital Transfer Vehicles for mid-1990s IOC. This paper presents the results of a study conducted by General Dynamics Convair Division, under contract to Aerojet TechSystems Company for NASA-LeRC, to define requirements, interface conditions, and operational design criteria for new LO2/LH2 propulsion systems applicable to future Orbit Transfer Vehicles, and to assess the impacts of space basing, man rating, and low-g transfer on propulsion system design requirements. The primary study emphasis was to determine what the OTV engine thrust level should be, how many engines are required on the OTV, and how the OTV engine should be designed. This was accomplished by evaluating planned OTV missions and concepts to determine the requirements for the OTV propulsion system, conducting tradeoffs and comparisons to optimize OTV capability, and evaluating reliability and maintenance to determine the recommended OTV engine design for future development.
A review of past high energy upper stage evolution leading to the first earth-lunar transfer vehicle, the Apollo S-IVB, is given and the current STS (Space Transportation System) upper-stage characteristics are discussed as a basis for future OTV (Orbital Transfer Vehicle) planning. Attention is given to requirements, operational factors, design features, and program alternatives as a background for a general review of OTV planning considerations for STS performance enhancement. The alternative discussed and other issues are under study by NASA and will be carefully assessed during the next few years as options are considered versus mission needs.
A cyclical Mars exploration scenario is described which combines the use of libration points in the earth-moon and sun-Mars systems with a pair of reusable nuclear electric propulsion (NEP) transfer vehicles. In this approach, the NEP transfer vehicles are stationed at the cislunar and cismartian libration points, and involves short cislunar point escape and capture spirals. It eliminates any spiral requirements at Mars by utilizing the cismartian point as a transportation node. Chemical/aerobrake vehicles are used to transport humans between earth and the cislunar point as well as between Mars and the cismartian point.
NASA-LeRC is sponsoring industry studies to establish the technology base for an advanced engine for orbital transfer vehicles for mid-1990s IOC. Engine contractors are being assisted by vehicle contractors to define the requirements, interface conditions, and operational design criteria for new LO2-LH2 propulsion systems applicable to future orbit transfer vehicles and to assess the impacts on space basing, man rating, and low-G transfer missions on propulsion system design requirements. The results of a study is presented. The primary study emphasis was to determine what the OTV engine thrust level should be, how many engines are required on the OTV, and how the OTV engine should be designed. This was accomplished by evaluating planned OTV missions and concepts to determine the requirements for the OTV propulsion system, conducting tradeoffs and comparisons to optimize OTV capability, and evaluating reliability and maintenance to determine the recommended OTV engine design for future development.
Simulations of aerobraking trajectories of aeroassisted orbital transfer vehicles (AOTV's) returning from geosynchronous orbit were analyzed to examine the effects of high-altitude viscous interactions and off-nominal atmospheres on AOTV return weight, heating, and loads performance. Viscous interaction effects encountered at high altitudes had little detrimental effect on the return weight capabilities for AOTV's representing a range of lift/drag ratios. Most of the AOTV return weight increase over an all-propulsive OTV occurred for a low lift/drag ratio. Smaller increases in return weight were observed for higher lift/drag ratios, at the expense of significantly higher heating and aerodynamic loads. Off-nominal atmospheres based on Shuttle-derived data and multipliers on a U.S. Standard Atmosphere were considered. AOTV's intended for entry under standard atmospheric conditions either deorbited during the pass through the off-nominal atmospheres or missed the target phasing orbit by wide margins. The AOTV's could successfully negotiate these atmospheres when new bank-angle histories were implemented with little loss and sometimes with a gain in return weight.
A slant-nose-cylinder aeroassisted orbital transfer vehicle configuration is described and analyzed in this study. The vehicle is sized for a 12,000 lb roundtrip payload between low earth orbit and geosynchronous orbit and is assumed to be space based. The vehicle can be fabricated using near-term technologies and is fully reusable. Optional advanced technologies offer potential for improved performance. The vehicle can be assembled on the ground and carried to orbit in the Shuttle cargo bay. An enclosed payload bay is provided in the vehicle to protect payloads during the pass through the atmosphere. The payload bay capacity can be increased from a 10 ft to a 14 ft diameter payload by replacing a modular section of the payload bay in space. The results of calculations used to size the vehicle and to predict its performance and weight are presented.
An aerobrake concept for a Lunar transfer vehicle was weight optimized through the use of the Taguchi design method, structural finite element analyses and structural sizing routines. Six design parameters were chosen to represent the aerobrake structural configuration. The design parameters included honeycomb core thickness, diameter to depth ratio, shape, material, number of concentric ring frames, and number of radial frames. Each parameter was assigned three levels. The minimum weight aerobrake configuration resulting from the study was approx. half the weight of the average of all twenty seven experimental configurations. The parameters having the most significant impact on the aerobrake structural weight were identified.
An aerobrake structural concept for a lunar transfer vehicle was weight optimized through the use of the Taguchi design method, finite element analyses, and element sizing routines. Six design parameters were chosen to represent the aerobrake structural configuration. The design parameters included honeycomb core thickness, diameter-depth ratio, shape, material, number of concentric ring frames, and number of radial frames. Each parameter was assigned three levels. The aerobrake structural configuration with the minimum weight was 44 percent less than the average weight of all the remaining satisfactory experimental configurations. In addition, the results of this study have served to bolster the advocacy of the Taguchi method for aerospace vehicle design. Both reduced analysis time and an optimized design demonstrated the applicability of the Taguchi method to aerospace vehicle design.
Freebird is a space-based orbital transfer vehicle designed to repair and deorbit orbital assets. Freebird is based at International Space Station Alpha (ISSA) at an inclination of 51.6 deg and is capable of three types of missions: crewed and teleoperated LEO missions, and extended robotic missions. In a crewed local configuration, the vehicle can visit inclinations between 30.8 deg and 72.4 deg at altitudes close to 390 km. Adding extra fuel tanks extends this range of inclination up to 84.9 deg and down to 18.3 deg. Furthermore, removing the crew module, using the vehicle in a teleoperated manner, and operating with extra fuel tanks allows missions to polar and geosynchronous orbits. To allow for mission flexibility, the vehicle was designed in a semimodular configuration. The major system components include a crew module, a 'smart box' (which contains command, communications, guidance, and navigation equipment), a propulsion pack, extra fuel tanks, and a vehicle storage facility (VSF) for storage purposes. To minimize risk as well as development time and cost, the vehicle was designed using only proven technology or technology which is expected to be flight-qualified in time for the intended launch date of 2002. And, because Freebird carries crew and operates near the space station, it must meet or exceed the NASA reliability standard of 0.994, as well as other standard requirements for such vehicles. The Freebird program was conceived and designed as a way to provide important and currently unavailable satellite repair and replacement services of a value equal to or exceeding operational costs.
Lunar transportation architectures exist for several different mission scenarios. Direct flights from Earth are possible, as the Apollo program clearly demonstrated. Alternatively, a space transfer vehicle could be constructed in space by using the Space Station as a base of operations, or multiple vehicles could be launched from Earth and dock in LEO without using a space station for support. Similarly, returning personnel could proceed directly to Earth or rendezvous at the Space Station for a ride back home on the Space Shuttle. Multiple design concepts exist which are compatible with these scenarios and which can support requirements of cargo, personnel, and mission objectives. Regardless of the ultimate mission selected, some technologies will certainly play a key role in the design and operation of advanced lunar transfer vehicles. Current technologies are capable of delivering astronauts to the lunar surface, but improvements are needed to affordably transfer the material and equipment that will be needed for establishing a lunar base. Materials and structures advances, in particular, will enable the development of more capable cryogenic fluid management and propulsion systems, improved structures, and more efficient vehicle assembly, servicing and processing.
This report presents the results of systems analyses and conceptual design of space transfer vehicles (STV). The missions examined included piloted and unpiloted lunar outpost support and spacecraft servicing, and unpiloted payload delivery to various earth and solar orbits. The study goal was to examine the mission requirements and provide a decision data base for future programmatic development plans. The final lunar transfer vehicles provided a wide range of capabilities and interface requirements while maintaining a constant payload mission model. Launch vehicle and space station sensitivity was examined, with the final vehicles as point design covering the range of possible options. Development programs were defined and technology readiness levels for different options were determined. Volume 1 presents the executive summary, volume 2 provides the study results, and volume 3 the cost and WBS data.
A concept for a lightweight space-based orbital transfer vehicle (OTV) featuring thin, spherical, pressure-designed aluminum liquid-oxygen and liquid-hydrogen tanks and a truss structure of composite materials is used as a baseline design for a large-cargo OTV. Vehicle sizing, fleet analysis, and parametric cost analysis are used to evaluate the effects of engine technology, vehicle staging, and high-thrust versus low-thrust transfer. Results indicate that Earth-to-orbit launch costs and OTV engine performance are strong drivers in orbital transportation cost and that there is no significant benefit in staging vehicles. At the low values of Earth-to-orbit cost representative of advanced launch vehicles, low-thrust and high-thrust OTV's are competitive.