SEARCH · Engineering Papers

Results for “Expendable launch vehicles”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

SourceResearch connectionFrom yearThrough year

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 451 records · Page 25

All sources NASA NTRS DOE OSTI

NASA NTRS2000

Advances in Space Transportation Technology Toward the NASA Goals

NASA has set goals to increase safety by one hundred times while reducing cost tenfold over the next decade. This dramatic increase in safety will come by departing from a past emphasis on cost and performance to a new paradigm of safety and reliability, which, in turn, will drive down cost. To accomplish this, vehicle systems must be inherently reliable, functionally redundant wherever practical and designed to minimize or eliminate catastrophic failure modes. Over the next twenty-five years, NASA has set goals to increase safety ten thousand times while reducing cost one hundred fold. Safety will increase towards today's airline safety and the low price per flight has the potential to enable a 15-fold increase in the size of the current projected space launch market. This level of improvement is comparable to the developments in the 1970s in the computer microprocessor when microchips went from hundreds of thousands to tens of thousands of dollars - hailing the era of the personal computer. In order to achieve such ambitious cost goals, today's partial and fully expendable rockets must be replaced by fully reusable systems. A RLV can eliminate assembly and checkout of the large number of complex interfaces on today's Space Shuttle. Full reusability will eliminate the necessity to throw away expensive hardware and the need for on-going production. Systems in ten years will have to accommodate 50 to 100 missions per year and could be commercially certified for hundreds of flights. In twenty-five years, the number of flights per year could jump to over 1,000, which will require certification for thousands of flights. The large increase in flights per year will demand that current operations and maintenance procedures be revolutionized. Unlike the current shuttle, which requires several thousand personnel over five months to process, the next generation system must be turned around in one week with less than one hundred personnel. In contrast to the rigorous disassembly and inspections required for the Space Shuttle's subsystems, the next generation vehicle's on-board health monitoring systems will could tell the ground crews which systems need replacement before landing. In twenty-five years, vehicles will be re-flown within one with crews numbering less than one hundred. Fully automated ground processing systems must require only a handful of personnel to launch the vehicle. Due to the increased intelligence of on-board systems, only cursory walk-around inspections would be required between flights An assessment of the progress in breakthrough technologies toward these goals by the NASA Advanced Space Transportation Program is presented. These breakthrough technologies include combined rocket and air breathing propulsion, high strength lightweight structures, high temperature materials, vehicle health management, and flight operations.

Lyles, Garry M.↗

NASA NTRS1999

From the Rocket Equation to Maxwell's Equations: Electrodynamic Tether Propulsion Nears Space Test

The US space program is facing a growing challenge to its decades-long, global leadership position, as current launch costs consume valuable resources and limit achievements in science, exploration, and commercial development. More than 40% of projected launches over the next 10 years have payloads with intended destinations beyond low-Earth orbit. Therefore, more cost-effective upper stages and on-board propulsion systems are critical elements in reducing total space transportation costs. A new type of space propulsion, using electrodynamic tethers, may be capable of performing multiple sequential missions without resupply and have a potential usable lifetime of several years. They may provide an in-space infrastructure that has a very low life cycle cost and greatly enhanced mission flexibility, thus supporting the goal of reducing the cost of access to space. Electrodynamic tether thrusters work by virtue of the force the Earth's magnetic field exerts on a wire carrying an electrical current. The effect is the basis for electric motors and generators. The Propulsive Small Expendable Deployer System (ProSEDS) experiment, planned for launch in the summer of 2000, will demonstrate the use electrodynamic tether thrust by lowering the altitude of a Delta-H rocket's upper stage on which it will be flying. Applications of the technology include a passive deorbit system for spacecraft at their end-of-life, reusable Orbit Transfer Vehicles, propellantless reboost of the International Space Station, and propulsion and power generation for future missions to Jupiter.

Johnson, Les↗

NASA NTRSDate not supplied

A Crew and Logistics Lander for the Common Habitat Architecture

The Common Habitat Architecture is a conceptual study that explores the use of a large habitat derived from the Space Launch System (SLS) liquid oxygen tank as a core habitation element that can be used for crew missions or training in 0g, 1/6g, 3/8g, and 1g. This is not part of current NASA Artemis mission planning, but instead represents an architecture that could potentially follow after the Artemis missions and initial human Mars landings. This architecture leverages Starship-derived vehicles for crew landing on and ascent from the Moon and Mars in support of long-duration surface missions with pressurized crew transfer baselined as a nominal capability. This is not the Human Landing System variant of Starship but represents a number of modifications to enable long-duration surface missions with an eight-person crew size. A rapid brainstorming study was conducted in February of 2021, using only public data, to identify options to deliver the 90-ton Common Habitat to the surface and emplace it at the intended habitation site. This study compared three lunar lander concepts and three Mars lander concepts before ultimately selecting the SpaceX Starship as the most viable lander. The Common Habitat Architecture assumes that this Starship variant can further be modified for crew and cargo delivery, assuming that the use of a common system will lead to cost benefits. It is assumed that the Starship will expend too much propellant in landing crew and logistics to be able to launch, given the assumed absence of surface propellant production. Consequently, a separable ascent stage is used for crew ascent. The pressurized elements of this modified Starship are discussed: Starship Ascent Module, Airlock, Transfer Tunnel, Pressurized Crew Transfer Module, and Logistics Modules. For each element, a description, dimensions, rough mass estimates, core capabilities, and design features are presented. Key mechanisms and internal structures of the starship are also discussed. This will include flame diverters for the ascend module, the orbital docking hatch, Pressurized Crew Transfer Module Garage door, Starship Ascent Module fairing, flame diverter blow-out panels, Pressurized Crew Transfer Module lift system, logistics module lift system, contingency crew ascent via the logistics module lift system, internal catwalks, and structural interfaces. Concept of Operations will be discussed for both Moon and Mars. This will include in-space crew rendezvous and transfer, crewed landing, shirtsleeve crew transfer, crew departure, contingency surface operations, and final element disposition. Key differences for Mars will be discussed, such as the Martian atmosphere, dust storms, the absence of crew handover, and Deep Space Exploration Vehicle rendezvous. Contingency microgravity maintenance access will also be discussed. This work will demonstrate viable pressurized crew transfer with a Starship-based lander architecture. Forward work includes Garage lighting and camera systems, Starship Ascent Module propulsion system and docking port trades, contingency habitation trades, and mass/power equipment estimation. Finally, forward work includes developing a heavy cargo return system derived from the Crew and Logistics Starship. The goal of this system is the return of at least 10,000 kg payloads from the surfaces of the Moon and Mars.

Lunar Lander↗

NASA NTRS1991

Shuttle-C auxiliary propulsion system certification approach and low-cost evolution concepts

Shuttle-C provides NASA with a heavy-lift cargo capability that is a low-cost derivative of the current Space Shuttle. This system can deliver 80,000 to 140,000 pounds of payload to low earth orbit within 4 years of authority to proceed. Shuttle-C will share th existing Shuttle launch facilities at Kennedy Space Center (KSC). The Shuttle-C vehicle incorporates an auxiliary propulsion system (APS) which provides separation from the external tank (ET), orbit circularization, on-orbit attitude hold, and deorbit capability. The APS is derived from the Shuttle orbiter aft reaction control system (ARCS) and orbital maneuvering system (OMS) with the objective of achieving minimum life cycle cost for this expendable vehicle. The resulting APS system will be described, its plan for a minimum-cost flight certification utilizing a cold-flow test program explained, and low-cost evolution by component design modification and part substitution presented.

Harsh, M. George↗

NASA NTRS2010

Spaceport Performance Measures

Spaceports have traditionally been characterized by performance measures associated with their site characteristics. Measures such as "Latitude" (proximity to the equator), "Azimuth" (range of available launch azimuths) and "Weather" (days of favorable weather) are commonly used to characterize a particular spaceport. However, other spaceport performance measures may now be of greater value. These measures can provide insight into areas of operational differences between competing spaceports and identify areas for improving the performance of spaceports. This paper suggests Figures of Merit (FOMs) for spaceport "Capacity" (number of potential launch opportunities per year and / or potential mass' to low earth orbit (LEO) per year); "Throughput" (actual mass to orbit per year compared to capacity); "Productivity" (labor effort hours per unit mass to orbit); "Energy Efficiency" (joules expended at spaceport per unit mass to orbit); "Carbon Footprint" tons CO2 per unit mass to orbit). Additional FOMS are investigated with regards to those areas of special interest to commercial launch operators, such as "Assignment Schedule" (days required for a binding assignment of a launch site from the spaceport); "Approval Schedule" (days to complete a range safety assessment leading to an approval or disapproval of a launch vehicle); "Affordability" (cost for a spaceport to assess a new launch vehicle); "Launch Affordability" (fixed range costs per launch); "Reconfigure Time" (hours to reconfigure the range from one vehicle's launch ready configuration to another vehicle's configuration); "Turn,Around Time" (minimum range hours required between launches of an identical type launch vehicle). Available or notional data is analyzed for the KSC/CCAFS area and other spaceports. Observations regarding progress over the past few decades are made. Areas where improvement are needed or indicated are suggested.

Finger, G. Wayne↗

NASA NTRS2008

Development Status of the J-2X

In June 2006, the NASA Marshall Space Flight Center (MSFC) and Pratt & Whitney Rocketdyne began development of an engine for use on the Ares I crew launch vehicle and the Ares V cargo launch vehicle. The development program will be completed in December 2012 at the end of a Design Certification Review and after certification testing of two flight configuration engines. A team of over 600 people within NASA and Pratt & Whitney Rocketdyne are currently working to prepare for the fall 2008 Critical Design Review (CDR), along with supporting an extensive risk mitigation test program. The J-2X will power the Ares I upper stage and the Ares V earth departure stage (EDS). The initial use will be in the Ares I, used to launch the Orion crew exploration vehicle. In this application, it will power the upper stage after being sent aloft on a Space Shuttle-derived. 5-segment solid rocket booster first stage. In this mission. the engine will ignite at altitude and provide the necessary acceleration force to allow the Orion to achieve orbital velocity. The Ares I upper stage, along with the J-2X. will then be expended. On the Ares V. first stage propulsion is provided by five RS-68B engines and two 5-segment boosters similar to the Ares I configuration. In the Ares V mission. the J-2X is first started to power the EDS and its payload. the Altair lunar lander. into earth orbit, then shut-down and get prepared for its next start. The EDS/Altair will remain in a parking orbit, awaiting rendezvous and docking with Orion. Once the two spacecraft are mated, the J-2X will be restarted to achieve earth departure velocity. After powering the Orion and Altair, the EDS will be expended. By using the J-2X Engine in both applications, a significant infrastructure cost savings is realized. Only one engine development is required, and the sustaining engineering and flight support infrastructures can be combined. There is also flexibility for changing, the production and flight manifest because a single production line can support both missions with minimal differences between each engine configuration kit.

Kynard, Mike↗

NASA NTRS1985

Overview od STS ground operations/orbiter turnaround STS-1 through STS-7

A review of STS-1 processing is presented as a reference for Shuttle processing time and the magnitude of the associated modifications, discrepancies, technical requirements, and ground systems activities. STS-1 processing provided the basis from which a plan to perform operational turnaround was established. Turnaround processing for Launch and Landing are treated separately to depict more clearly the progress made to reduce turnaround time and manhours expended. Turnaround time was reduced from 187 days for STS-2 to 60 days for STS-7 and landing turnaround was reduced from 14 days to 5 days. Modifications were reduced from 114 to 51. Requirements changes for launch readiness verification were reduced from 536 to 107. Special tests or inspections were reduced from 292 to 52. Anomalies resolved concurrent with processing were reduced from 13,000 to 4,000. Integration in the Vehicle Assembly Building (VAB) is about 10 percent of the flow, and Launch Pad operations comprise 40 percent of the turnaround. The work force has matured and has demonstrated the capability to perform planned and contingency operations. The turnaround program is fast approaching the goal of becoming operational.

Schwartz, R.↗

NASA NTRS1993

Reliability of solid rocket motor cases and nozzles

A recent article in Aerospace America claims that 'the average success ratio of the current U.S. stable of launch vehicles, including upper stages, is about 92 percent (without upper stages it is close to 95 percent). The 8 percent failure probability implies an expected loss of $12M per flight, not including the lost opportunity costs'. Since payload costs are likely to be much greater than launch costs and even more so for the new launch vehicles for the Advanced Launch Development Program (ALDP), the cost of rocket motor unreliability at the current 8 percent rate can run into billions of dollars if expected increases in demand are realized. At an 8 percent failure rate, it is extremely unlikely that failure will occur during the first few ground tests of a new system. At that time, most of the design, analysis and tooling costs of the program have not expended. Since most systems are expected to be used ten to a hundred or more times, the likelihood of one or more failures is very large, and it can be expected that the above losses will be realized in the future. This will occur unless the problems are addressed and remedied. Recent trends suggest the problem is not being addressed adequately.

Crose, James G.↗

NASA NTRS2022

Concept for 2033 Crewed Mars Orbital Mission with Venus Flyby

The 2033 Mars launch period provides a unique opportunity for a round trip mission with a total flight time of only 1.6 years. A concept is presented to perform a crewed Mars orbital mission in 2033 that would minimize development and mission risk by using conventional hypergolic in-space propulsion stages with a common design. The Mars mission vehicle would include a Mars transit habitat, an Orion spacecraft, and chemical propulsion stages. It would be launched by a combination of SLS and commercial rockets to be aggregated in high Earth orbit or at the Lunar Gateway. The crew would launch to the mission vehicle in Orion as the final element in the assembly. After transit to Mars, the crew would spend about 30 days in a high Mars orbit and then return to Earth via a Venus flyby and gravity assist. A sunshade would be deployed for thermal control of the mission vehicle inside 1 AU. The crew would return directly to Earth in the Orion crew module, with the transit habitat being expended. The 2033 mission would not be a “one-off”, but could be a pathfinder for crew transport for landing missions to follow, perhaps starting as early as 2037.

Woolley, Ryan↗

NASA NTRS1987

Configuration evaluation and criteria plan. Volume 2: Evaluation criteria plan (update). Space Transportation Main Engine (STME) configuration study

Candidate main engine configurations which enhance vehicle performance, operation and cost are identified. These candidate configurations are evaluated and the configurations which provide significant advantages over existing systems are selected for consideration for the next generation of launch vehicles. The unbiased selection of the Space Transportation Main Engine (STME) configuration requires that the candidate engines be evaluated against a predetermined set of criteria which must be properly weighted to emphasize critical requirements defined prior to the actual evaluation. During a prior study of the STME a Gas Generator Cycle engine was selected for conceptual design, with emphasis on reusability, reliability and low cost while achieving good performance. In this study emphasis is on expendable application of the STME while maintaining low cost and high reliability.

Bair, E. K.↗

NASA NTRS2006

RS-25 for the NASA Cargo Launch Vehicle: The Evolution of SSME for Space Exploration

A key element of the National Vision for Space Exploration is the development of a heavy-lift Cargo Launch Vehicle (CaLV). Missions to the Moon, Mars, and beyond are only possible with the logistical capacity of putting large payloads in low-earth orbit. However, beyond simple logistics, there exists the need for this capability to be as cost effective as possible to ensure mission sustainability. An element of the CaLV project is, therefore, the development of the RS-25, which represents the evolution of the proven Space Shuttle Main Engine (SSME) into a high-performance, cost-effective expendable rocket engine. The development of the RS-25 will be built upon the foundation of over one million seconds of accumulated hot-fire time on the SSME. Yet in order to transform the reusable SSME into the more cost-effective, expendable RS-25 changes will have to be made. Thus the project will inevitably strive to maintain a balance between demonstrated heritage products and processes and the utilization of newer technology developments. Towards that end, the Core Stage Engine Office has been established at the NASA Marshall Space Flight Center to initiate the design and development of the RS-25 engine. This paper is being written very early in the formulation phase of the RS-25 project. Therefore the focus of this paper will be to present the scope, challenges, and opportunities for the RS-25 project. Early schedules and development decisions and plans will be explained. For not only must the RS-25 project achieve cost effectiveness through the development of new, evolved components such as a channel-wall nozzle, a new HIP-bonded main combustion chamber, and several others, it must simultaneously develop the means whereby this engine can be manufactured on a scale never envisioned for the SSME. Thus, while the overall project will span the next eight to ten years, there is little doubt that even this schedule is aggressive with a great deal of work to accomplish.

Kynard, Michael H.↗

NASA NTRS2008

Reusability Studies for Ares I and Ares V Propulsion

With a mission to continue to support the goals of the International Space Station (ISS) and explore beyond Earth orbit, the United States National Aeronautics and Space Administration (NASA) is in the process of launching an entirely new space exploration initiative, the Constellation Program. Even as the Space Shuttle moves toward its final voyage, Constellation is building from nearly half a century of NASA spaceflight experience, and technological advances, including the legacy of Shuttle and earlier programs such as Apollo and the Saturn V rocket. Out of Constellation will come two new launch vehicles: the Ares I crew launch vehicle and the Ares V cargo launch vehicle. With the initial goal to seamlessly continue where the Space Shuttle leaves off, Ares will firstly service the Space Station. Ultimately, however, the intent is to push further: to establish an outpost on the Moon, and then to explore other destinations. With significant experience and a strong foundation in aerospace, NASA is now progressing toward the final design of the First Stage propulsion system for the Ares I. The new launch vehicle design will considerably increase safety and reliability, reduce the cost of accessing space, and provide a viable growth path for human space exploration. To achieve these goals, NASA is taking advantage of Space Shuttle hardware, safety, reliability, and experience. With efforts to minimize technical risk and life-cycle costs, the First Stage office is again pulling from NASA's strong legacy in aerospace exploration and development, most specifically the Space Shuttle Program. Trade studies have been conducted to evaluate lifecycle costs, expendability, and risk reduction. While many first stage features have already been determined, these trade studies are helping to resolve the operational requisites and configuration of the first stage element. This paper first presents an overview of the Ares missions and the genesis of the Ares vehicle design. It then looks at one of the most important trade studies to date, the "Ares I First Stage Expendability Trade Study." The purpose of this study was to determine the utility of flying the first stage as an expendable booster rather than making it reusable. To lower the study complexity, four operational scenarios (or cases) were defined. This assessment then included an evaluation of the development, reliability, performance, and transition impacts associated with an expendable solution. The paper looks at these scenarios from the perspectives of cost, reliability, and performance. The presentation provides an overview of the paper.

Williams, Thomas J.↗

NASA NTRS1989

Advanced manned launch systems

Advanced Manned Launch Systems (AMLS) studies underway at the NASA Langley Research Center are described. The AMLS studies have focused on investigating mission needs and identifying technologies, operational requirements, and life-cycle costs. Two-stage vertical take off rocket vehicles with significant levels of reusability are shown to be cost effective for moderate to high flight rates. All-hydrogen fuel vehicles are preferred over methane boosted systems for both operational and cost considerations. Drawings of possible AMLS design solutions are presented. Fully reusable, partially resusable, expendable stage and horizontal take off airbreather/rocket solutions are compared.

Talay, Theodore A.↗

NASA NTRS2008

Reusability Studies for Ares I and Ares V Propulsion

With a mission to continue to support the goals of the International Space Station (ISS) and explore beyond Earth orbit, the United States National Aeronautics and Space Administration (NASA) is in the process of launching an entirely new space exploration initiative, the Constellation Program. Even as the Space Shuttle moves toward its final voyage, Constellation is building from nearly half a century of NASA spaceflight experience, and technological advances, including the legacy of Shuttle and earlier programs such as Apollo and the Saturn V rocket. Out of Constellation will come two new launch vehicles: the Ares I crew launch vehicle and the Ares V cargo launch vehicle. With the initial goal to seamlessly continue where the Space Shuttle leaves off, Ares will firstly service the Space Station. Ultimately, however, the intent is to push further: to establish an outpost on the Moon, and then to explore other destinations. With significant experience and a strong foundation in aerospace, NASA is now progressing toward the final design of the First Stage propulsion system for the Ares I. The new launch vehicle design will considerably increase safety and reliability, reduce the cost of accessing space, and provide a viable growth path for human space exploration. To achieve these goals, NASA is taking advantage of Space Shuttle hardware, safety, reliability, and experience. With efforts to minimize technical risk and life-cycle costs, the First Stage office is again pulling from NASA s strong legacy in aerospace exploration and development, most specifically the Space Shuttle Program. Trade studies have been conducted to evaluate life-cycle costs, expendability, and risk reduction. While many first stage features have already been determined, these trade studies are helping to resolve the operational requisites and configuration of the first stage element. This paper first presents an overview of the Ares missions and the genesis of the Ares vehicle design. It then looks at one of the most important trade studies to date, the "Ares I First Stage Expendability Trade Study." The purpose of this study was to determine the utility of flying the first stage as an expendable booster rather than making it reusable. To lower the study complexity, four operational scenarios (or cases) were defined. This assessment then included an evaluation of the development, reliability, performance, and transition impacts associated with an expendable solution. This paper looks at these scenarios from the perspectives of cost, reliability, and performance.

Williams, Thomas J.↗

NASA NTRS2003

X-43C Flight Demonstrator Project Overview

The X-43C Flight Demonstrator Project is a joint NASA-USAF hypersonic propulsion technology flight demonstration project that will expand the hypersonic flight envelope for air-breathing engines. The Project will demonstrate sustained accelerating flight through three flights of expendable X-43C Demonstrator Vehicles (DVs). The approximately 16-foot long X-43C DV will be boosted to the starting test conditions, separate from the booster, and accelerate from Mach 5 to Mach 7 under its own power and autonomous control. The DVs will be powered by a liquid hydrocarbon-fueled, fuel-cooled, dual-mode, airframe integrated scramjet engine system developed under the USAF HyTech Program. The Project is managed by NASA Langley Research Center as part of NASA's Next Generation Launch Technology Program. Flight tests will be conducted by NASA Dryden Flight Research Center off the coast of California over water in the Pacific Test Range. The NASA/USAF/industry project is a natural extension of the Hyper-X Program (X-43A), which will demonstrate short duration (approximately 10 seconds) gaseous hydrogen-fueled scramjet powered flight at Mach 7 and Mach 10 using a heavy-weight, largely heat sink construction, experimental engine. The X-43C Project will demonstrate sustained accelerating flight from Mach 5 to Mach 7 (approximately 4 minutes) using a flight-weight, fuel-cooled, scramjet engine powered by much denser liquid hydrocarbon fuel. The X-43C DV design flows from integrating USAF HyTech developed engine technologies with a NASA Air-Breathing Launch Vehicle accelerator-class configuration and Hyper-X heritage vehicle systems designs. This paper describes the X-43C Project and provides the background for NASA's current hypersonic flight demonstration efforts.

Moses, Paul L.↗

NASA NTRS2003

X-43C Plans and Status

X-43C Project is a hypersonic flight demonstration being executed as a collaboration between the National Aeronautics and Space Administration (NASA) and the United States Air Force (USAF). X-43C will expand the hypersonic flight envelope for air breathing engines beyond the history making efforts of the Hyper-X Program (X-43A). X-43C will demonstrate sustained accelerating flight during three flight tests of expendable X-43C Demonstrator Vehicles (DVs). The approximately 16-foot long X-43C DV will be boosted to the starting test conditions, separate from the booster, and accelerate from Mach 5 to Mach 7 under its own power and autonomous control. The DVs are to be powered by a liquid hydrocarbon-fueled, fuel-cooled, dual-mode, airframe integrated scramjet engine system developed under the USAF HyTech Program. The Project is managed by NASA Langley Research Center as part of NASA s Next Generation Launch Technology Program. Flight tests will be conducted by NASA Dryden Flight Research Center over water off the coast of California in the Pacific Test Range. The NASA/USAF/industry project is a natural extension of the Hyper-X Program (X-43A), which will demonstrate short duration (~ 10 seconds) gaseous hydrogen-fueled scramjet powered flight at Mach 7 and Mach 10 using a heavyweight, largely heat sink construction, experimental engine. The X-43C Project will demonstrate sustained accelerating flight from Mach 5 to Mach 7 (~ 4 minutes) using a flight-weight, fuel-cooled, scramjet engine powered by much denser liquid hydrocarbon fuel. The X-43C DV design flows from integrating USAF HyTech developed engine technologies with a NASA Air Breathing Launch Vehicle accelerator-class configuration and Hyper-X heritage vehicle systems designs. This paper describes the X-43C Project and provides background for NASA s current hypersonic flight demonstration efforts.

Moses, Paul L.↗

NASA NTRS1998

Measuring Thermal Conductivity and Moisture Absorption of Cryo-Insulation Materials

NASA is seeking to develop thermal insulation material systems suitable for withstanding both extremely high temperatures encountered during atmospheric re-entry heating and aero- braking maneuvers, as well as extremely low temperatures existing in liquid fuel storage tanks. Currently, materials used for the high temperature insulation or Thermal Protection System (TPS) are different from the low temperature, or cryogenic insulation. Dual purpose materials are necessary to the development of reusable launch vehicles (RLV). The present Space Shuttle (or Space Transportation System, STS) employs TPS materials on the orbiter and cryo-insulation materials on the large fuel tank slung under the orbiter. The expensive fuel tank is jettisoned just before orbit is achieved and it burns up while re-entering over the Indian Ocean. A truly completely reusable launch vehicle must store aR cryogenic fuel internally. The fuel tanks will be located close to the outer surface. In fact the outer skin of the craft will probably also serve as the fuel tank enclosure, as in jet airliners. During a normal launch the combined TPS/cryo-insulation system will serve only as a low temperature insulator, since aerodynamic heating is relatively minimal during ascent to orbit. During re-entry, the combined TPS/cryo-insulation system will serve only as a high temperature insulator, since all the cryogenic fuel will have been expended in orbit. However, in the event of an.aborted launch or a forced/emergency early re-entry, the tanks will still contain fuel, and the TPS/cryo-insulation will have to serve as both low and high temperature insulation. Also, on long duration missions, such as to Mars, very effective cryo-insulation materials are needed to reduce bod off of liquid propellants, thereby reducing necessary tankage volume, weight, and cost. The conventional approach to obtaining both low and high temperature insulation, such as is employed for the X-33 and X-34 spacecraft, is to use separate TPS and cryo-insulation materials, which are connected by means of adhesives or stand-offs (spacers). Three concepts are being considered: (1) the TPS is bonded directly to the cryo-insulation which, in turn, is bonded to the exterior of the tank, (2) stand-offs are used to make a gap between the TPS and the cryo-insulation, which is bonded externally to the tank, (3) TPS is applied directly or with stand-offs to the exterior so the tank, and cryo-insulation is applied directly to the interior of the tank. Many potential problems are inherent in these approaches. For example, mismatch between coefficients of thermal expansion of the TPS and cryo-insulation, as well as aerodynamic loads, could lead to failure of the bond. Internal cryo-insulation must be prevent from entering the sump of the fuel turbo-pump. The mechanical integrity of the stand-off structure (if used) must withstand multiple missions. During ground hold (i.e., prior to launch) moisture condensation must be minimized in the gap between the cryo-insulation and the TPS. The longer term solution requires the development of a single material to act as cryo- insulation during ground hold and as TPS during re-entry. Such a material minimizes complexity and weight while improving reliability and reducing cost.

Lambert, Michael A.↗

NASA NTRS2019

Electric Sail Propulsion for Deep Space Missions

Electrostatic Sail (E-sail) propulsion extracts momentum from the solar wind, which has minimal speeds of 400 km/sec, through electrostatic repulsion of the positively charged solar wind ions. This momentum exchange is accomplished by an array of multi-kilometer-length charged tethers biased to a high positive voltage, producing an electric field around the tethers that deflect the positively charged solar wind particles. This electric field grows in diameter as the spacecraft moves away from the sun, increasing the E-sail effective area. The growth of the E-sail effective area allows the created propulsive force to decrease at a rate closer to 1/r out to distances from the sun of up to 20 Astronomical Units (AU). This is unlike solar sail propulsion, where the thrust decreases at a rate of 1/r2 and is only effective out to distances of ~5 AU. The propulsive force is applied without expending any propellant. Although the thrust generated by an E-Sail is low, it can be applied continuously over a period of years (depending on the mission type), and can push a 500 kg spacecraft to tremendous velocities—as high as 12 AU per year—making the voyage to 600 AU in as few as 50 years after launch. The NASA Marshall Space Flight Center recently completed a NASA Innovative Advanced Concept (NIAC) study in which all of the major system- and subsystem-level aspects of an E-Sail-propelled vehicle were assessed and a maturation plan developed. If implemented, E-Sail propulsion could be realized and deployed as a viable option for space missions within a decade.

Interstellar Probe↗